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Produced by the NASA Center for Aerospace Information (CASI) 



BOEING DOCUMENT^ 
NO. D5-17142 


VOLUME II OF 1! 
TECHNICAL VOLUME 


(NASA-CR-124C57) SPACE TUG AEROBRAKIHG 
STUDY. VOLUME 2: TECHNICAL Final 

Report (Boeing Co. , Huntsville* Ala.) 

519 p HC $28.00 CSCL 22A 




N73-18827 1 - 


Onclas 

G3/30 17104 


< 1 ? iS-’ 


r / s / 

'-J / \ ! 


[A i 


SPACE TUG WITH AEROBRAICING KIT 



\ CIRCULARIZE, 

£arYh1 rendezvous with 
S i SHUTTLE, RETURN 
TO EARTH 

iv ./ 


GEOSYNCHRONOUS 

ORBIT ) 


DEORBIT 


DECREASING 

APOGEE 


AEROBRAKING RETURN TRAJECTORY PROFILE 






D5-17142 


FINAL REPORT 


SPACE TUG AERQBRAKING STUDY 


VOLUME II OF II 
TECHNICAL VOLUME 

PREPARED BY : p 
C.' J. eORSO 





.Ylkwvtt&d 

J^TEPHU.lUNRUb 


VJ 


PREPARED UNDER CONTRACT NAS8-27S01 
FOR 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 
GEORGE C. MARSHALL SPACE FLIGHT CENTER 
MARSHALL SPACE FLIGHT CENTER, ALABAMA 35S12 


APRIL 12,1972 


DB-17142 


THE BOEING COMPANY 
SATURN/APOLLO/SKYLAB DIVISION 
HUNTSVILLE, ALABAMA 


PRECEDING PAGE BLANK NOT FIT, MED 


05-17142 






05-17142 


ABSTRACT 


The feasibility and practicality of employing an aerobraklng trajectory 
for return of the reusable Space Tug from geosynchronous and other high 
energy missions was investigated. The aerobraking return trajectory 
modes from high orbits employ transfer ellipses which have low perigee 
altitudes wherein the earth's sensible atmosphere provides drag to 
reduce the Tug descent delta velocity requirements and thus decrease 
the required return trip propulsive energy. An aerobraked Space Tug, 
sized to the Space Shuttle payload capability and dimensional con- 
straints, can accomplish 95 percent of the geosynchronous missions with 
a single Shuttle/Tug launch per mission. Orbital assembly and/or 
orbital propellant transfer operations are not required. The same size 
Space Tug using conventional trajectory modes cannot deliver any pay- 
load to the geosynchronous orbit. 

Aerodynami cs , aerothermodynami cs , trajectory, guidance and control, 
configuration concepts, materials, weights and performance parameters 
were identified. Sensitivities to trajectory uncertainties, atmos- 
pheric anomalies and re-entry environments were determined. New 
technology requirements and future studies required to further enhance 
the aerobraking potential were identified. 


KEY WORDS 


Space Tug 
Aerobraking 
Return trajectories 
Synchronous missions 
Orbit-to-Orbit Shuttle (00S) 
Earth- to-Orbit Shuttle (EOS) 


Space Tug Kits 
Flare Concepts 
Heat Shield Concepts 
Propulsion modules 
Astrionics modules 
Aerodynamic drag 


iii 


FOREWORD 


This Technical Volume is one of two volumes presenting the results of a 
study to investigate the feasibility of an aerobraking trajectory mode 
for return of the reusable Space Tug from geosynchronous equatorial 
orbit. The Executive Summary Volume presents a brief outline of the 
objectives, summarizes the results and gives conclusions and recommendations 
of this study. This Technical Volume presents the detailed technical re- 
sults. 

The Boeing Company performed this study at the Boeing-Huntsville facility 
for the NASA Marshall Space Flight Center, Huntsville, Alabama. The 
NASA/MSFC Technical Monitor was Thomas W. Barrett, Advanced Systems 
Analysis Office, Vehicle Systems Group. Subcontractor to The Boeing 
Company for the navigation requirements impact on the astrionic module 
was the International Business Machine Corporation, Huntsville facility. 


D5-17142 


CONTENTS 


PARAGRAPH 


REVISIONS 

ABSTRACT 

FOREWORD 

CONTENTS 

LIST OF FIGURES 

ABBREVIATIONS AND ACRONYMS 

SECTION I - INTRODUCTION 

1.0 GENERAL 

1.1 STUDY OBJECTIVES AND APPROACH 

1.2 BACKGROUND 

SECTION II - SUMMARY 

2.0 CONCLUSIONS AND RECOMMENDATIONS 

2.1 SUMMARY OF TECHNICAL STUDY RESULTS 

2.1.1 Aerodynamics Analysis 

2. 1.1.1 Aerodynamics Conclusions 

2.1.2 Conf igurati ons 

2. 1.2.1 Configuration Conclusions 

2.1.3 Trajectory Analysis 

2. 1.3.1 Trajectory Conclusions 

2.1.4 Control Analysis 

2. 1.4.1 Control Conclusions 

2.1.5 Thermal Analysis 

2. 1.5.1 Thermal Conclusions 

2.1.6 Astrionlcs Analysis 

2. 1 . 6 . 1 As tri oni cs Concl us ions 

2.1.7 Materials 

2. 1.7.1 Materials Conclusions 

2.1.8 Weights 

2.1.B.1 Weights Conclusions 

2,2 SUMMARY OF SENSITIVITY ANALYSIS RESULTS 

2.2.1 Sensitivities of Payloads and Temperatures for Tug 
Configurations for Standard Atmosphere 

2.2.2 Sensitivities of Payloads and Temperatures for Tug 
Configurations for Atmospheric Perturbations and 
Navigation Errors 

2.2*3 Conventi onal/Aerobraktng Tug Performance Comparison 

2.2.4 Sensitivity Conclusions 




PAGE 

ii, 

* » * 
m, 

iv 

v 

x 

XX 

1-1 

1 - 1 . 

1-2 

1-9 


2-1 

2-4 

2-4 

2-8 

2-8 

2-14 

2-14 

2-16 

2-18 

2-18 

2-21 

2-24 

2-24 

2-29 

2-30 

2-30 

2-31 

2-34 

2-38 

2-38 

2-41 


2-43 

2-46 


V 



D5-17142 


CONTENTS (Continued) 

PARAGRAPH PAGE 

SECTION III - GROUNDRULES, GUIDELINES AND 
ASSUMPTIONS 

3.0 GENERAL 3-1 

3.1 OVERALL SPACE TUG 3-1 

3.2 PROPULSION MODULE 3-2 

3.3 ASTRIONICS MODULE 3-3 

SECTION IV - PERFORMANCE AND TRADE STUDIES 

4.0 GENERAL 4-1 

4.1 AERODYNAMIC ANALYSIS 4-1 

4.1.1 Basic Tug Aerodynamic Analysis 4-4 

4. 1.1.1 Drag Data 4-4 

4. 1.1. 2 Static Stability Data 4-7 

4. 1.1. 3 Nose Pressure Distribution 4-11 

4.1.2 Flared Tug Configuration Analysis 4-14 

4.1. 2.1 Flow Field and Flare Sizing Analysis 4-14 

4. 1.2. 2 Flare Selection Rationale 4-16 

4. 1.2. 3 Flared Tug Drag Characteristics 4-25 

4. 1.2. 4 Static Stability Analysis 4-26 

4.1. 2. 5 Pressure Coefficient Distribution Data 4-31 

4.2 AEROBRAKING CONFIGURATION CONCEPTS 4-38 

4.2.1 Conventional Trajection Space Tug Configuration 4-39 

4.2.2 Aerobraking Kit Elements 4-41 

4.2.2. 1 Aft Heat Shield 4-43 

4.2. 2. 2 Sidewall Thermal Protection System 4-55 

4.2.2. 3 Aerodynamic Flare 4-57 

4.. 2, 2.4 Payload Adapter 4-63 

4.2.3 Configuration Design Options and Recommendations 4-65 

4.3 TRAJECTORY ANALYSIS 4-67 

4.3.1 Analytical Model 4-67 

4.3. 1.1 Assumptions and Groundrules 4-69 

4.3.2 Preliminary Trajectory Analysis 4-70 

4.3.3 Final Trajectory Analysis 4-70 

4.3.4 Dispersed Atmosphere Effects 4-87 

4. 3.4.1 Trajectory Correction Techniques 4-91 

4. 3.4.2 Dispersed Atmosphere Effects on Inertial Velocity 4-117 

4.3.5 Lunar, Solar and Earth Harmonics Perturbations 4-123 


D5-17142 


PARAGRAPH 

4.4 

4.4.1 

4.4.2 

4.4.3 

4.4.4 

4.4.5 

4.5 

4.5.1 

4.5. 1.1 

4.5.2 

4.5.3 

4.5.4 

4.5.4. 1 

4.5. 4.2 

4. 5. 4.3 

4.5. 4.4 

4.6 

4.6.1 

4.6.2 

4.6.3 

4.6. 3.1 

4.6. 3. 2 

4.6. 3. 3 

4. 6. 3. 4 

4.6.4 

4.6.4.1 

4.6.4. 2 

4.6. 4. 3 

4.6. 4. 4 

4.6. 4.5 

4. 6. 4.6 

4. 6. 4.7 

4.6.5 

4.6. 5.1 

4.6. 5. 2 

4.6.6 

4.6.6. 1 

4.6. 6. 2 

4.6. 6. 3 

4.6. 6. 4 

4. 6. 6. 5 

4.6. 6. 6 




CONTENTS (Continued) 


CONTROL ANALYSIS 
Control Methods 

Atmospheric Effects on Control 
Astrlonlcs Effect on Control Requirements 
Control Options 

Dpcr tl "he 

THERMAL ANALYSIS 

Analytical Methods 

Assumptions 

Heating Rates 

Equilibrium Temperatures 

Thermal Protection System Requirements 

Heat Shield Dome 

Flare Structure 

Sidewall Area 

Thermal Protection System Weights 
ASTRIONICS ANALYSIS 
Introduction 

Astrionic System Study Groundrules and Guidelines 

Astrionic System Analysis Summary and Observations 

Astrionic System Aerobraking Weight Deltas 

Navigation Analysis 

Radiation Impacts Analysis 

Observations and Conclusions 

Navigation Accuracy Analysis 

Introduction 

Navigation Sensors . 

Analytical Tools 

Synchronous Orbit Navigation Uncertainty Analysis 
Aerobraking Descent Orbit Navigation Accuracy 
Analysis 

Trajectory Correction Burn Uncertainty Analysis 
Aerobraking Navigation Uncertainty Analysis 
Sensitivities 

Astrionic System Configuration 

Astrionic System Configuration Description 

Weight and Power Summary 

Redundancy Analysis 

Redundancy Study Assumptions 

Analytic Programs 

Reliability Enhancement/Coverage 

Nominal Mission without Aerobraking 

Aerobraking Redundancy Impacts 

Redundancy Analysis Observations 


vi i 


PAGE 

4-125 

4-125 

4-127 

4-130 

4-130 

4-131 

4-139 

4-139 

4-140 

4-140 

4-140 

4-156 

4-156 

4-156 

4-156 

4-163 

4-163 

4-163 

4-166 

4-167 

4-167 

4-167 

4-169 

4-169 

4-171 

4-171 

4-172 

4-174 

4-174 

4-175 

4-192 

4-195 

4-200 

4-202 

4-202 

4-202 

4-207 

4-207 

4-207 

4-209 

4-209' 

4-209 


D5-17142 


CONTENTS (Continued) 

PARAGRAPH PAGE 

4.6.7 Power-Weight Impacts 4-211 

4.6.7. 1 Space Tug Power Requirements 4-211 

4.6. 7. 2 Aerobraking Weight Deltas 4-211 

4.6.8 Astrionic System Radiation Impacts 4-211 

4.6.8. 1 External Radiation Envi ronments 4-213 

4. 6.8.2 Internal Radiation Environments 4-213 

4. 6. 8. 3 Electronic Device Effects 4-215 

4. 6. 8. 4 Observations 4-218 

4.6.9 New Technology Implications 4-219 

4.6.9. 1 Redundancy Implementation 4-219 

4. 6. 9. 2 Navigation Sensor Integration 4-219 

4.6.9. 3 Navigation and Guidance Analyses 4-220 

4.6.10 Follow-On Study Effort 4-220 

4.6.10.1 Astrionic System Configuration Analysis 4-220 

4.6.10.2 Redundancy Analysis 4-220 

4.6.10.3 Navigation Timeline Analysis 4-221 

4.6.10.4 Advanced Sensor Systems 4-221 

4.6.10.5 Updating Capability versus Control Requirements 4-221 

4.6.10.6 Radiation Analysis 4-221 

4.6.10.7 Astrionic System New Technology Component Analysis 4-221 

4.7 Aerobraking Kit Materials Selection 4-222 

4.7.1 Materials Groundrules ar.d Criteria 4-222 

4.7.2 Thermal Environments of Aerobraking Kit Elements 4-223 

4.7.3 Materials Selection 4-226 

4. 7.3.1 Aft Heat Shield Materials Selection 4-244 

4. 7.3.2 Sidewall Insulation 4-244 

4. 7.3. 3 Flare Materials Selection 4-245 

4. 7. 3. 4 Payload Adapter Materials Selection 4-245 

4.7.4 Options and Recommendations 4-246 

4.8 WEIGHTS AND MASS PROPERTIES 4-246 

4.8.1 Mass Properties Summary 4-246 

4.8.2 Structural Weights 4-246 

4. 8. 2.1 Aft Heat Shield 4-254 

4. 8. 2. 2 Flared Skirt 4-254 

4.8.2. 3 Payload Adapter/Flared Skirt Support 4-254 

4.8.3 Thermal Protection System Weight 4-254 

4.8.4 Astrionics System Weight 4-256 

4.8.5 Control System Weight 4-256 

4.8.6 Total Tug Weight 4-256 

5.0 SENSITIVITY ANALYSES 5-1 

5.1 PAYL0AD/AER08 RAKING PERFORMANCE OPTIMIZATION , 5-1 

5.1.1 General Parametric Results 5-3 

5. 1.1.1 Inert Weight and Delta Velocity Effects 5-3 

5. 1.1. 2 RCS Isp Effects 5-7 

5. 1.1. 3 Circularization Altitude and EOS Rendezvous Effects 5-12 

5. 1.1. 4 General Parametric Conclusions 5-20 


vlli 


D5-17142 

CONTENTS (Continued) 

PARAGRAPH PAGE 

5.1.2 Specific Payload Capability Assessment 5-20 

5. 1.2.1 Delta Velocity Budgets 5-24 

5. 1.2. 2 Basic (No Flare) Configurations 5-27 

5. 1.2. 3 30° Flare Configuration 5-30 

5. 1.2. 4 45° Flare Configuration 5-34 

5. 1.2. 5 60° Flare Configuration 5-34 

5. 1.2. 6 Configuration Payload Comparison 5-34 

5.2 ATMOSPHERIC PERTURBATION AND NAVIGATION ERROR 

SENSITIVITIES 5-47 

5.2.1 Effects on Nominal Perigee Altitudes 5-47 

5.2.2 Payload Sensitivities 5-53 

5.2.3 Perturbation Summary and Conclusions 5-56 

5.3 AERQBRAKING/CONVENTIONAL TRAJECTORY TUG 

COMPARISONS 5-59 

5.3.1 Conventional Trajectory Tug Size Comparison 5-59 

5.3.2 Sensitivities to Performance Parameters 5-61 

5.3.2. 1 Mass Fraction 5-61 

5. 3. 2. 2 Specific Impulse 5-63 

5. 3. 2. 3 Delta Velocity 5-63 

5.4 SENSITIVITY ANALYSIS CONCLUSIONS AND RECOMMENDATIONS 5-63 

LIST OF REFERENCES 1 

APPENDICES 

A TWO PASS AEROB RAKING SPACE TUG ANALYSIS A-l 

B BRIEF DESCRIPTION OF KALMAN FILTERING B-l 

C ANALYSIS PROGRAMS C-l 

0 NAVIGATION COMPONENTS D-1 

E RECOMMENDED AEROBRAKXNG FOLLOW-ON ACTIVITIES E-1 

F TWO PASS LIGHT WEIGHT LARGE FLARE F-l 


D5-17142 


FIGURE 

1.1. 0.0-1 
1.1. 0.0-2 

1.1. 0. 0-3 

1.1. 0. 0-4 

1.1. 0. 0-5 

1 . 2 . 0 . 0-1 

1 .2, 0.0-2 
2 . 0 . 0 . 0—1 
2 . 1 . 1 . 0-1 
2 . 1 . 1 . 0-2 

2.1. 2.0-1 
2.1. 2.0-2 

2.1. 2.0- 3 

2.1. 2.0- 4 

2.1. 3.0- 1 

2. 1.3. 0- 2 

2.1. 4.0- 1 

2.1. 5.0- 1 

2.1. 5.0- 2 

2.1. 5.0- 3 


2.1. 6.0-1 

2.1. 6.0-2 

2.1. 6.0- 3 

2. 1.8. 0- 1 
2.1 .8.0-2 

2.1. 8.0- 3 

2.1. 8.0- 4 

2.1 .8.0- 5 

2. 2. 0.0-1 

2 . 2 . 1 . 0 - 1 
2. 2. 2.0-1 
2.2. 2. 0-2 


LIST OF FIGURES 


Aerobraking Feasibility Study Logic 
Conventional Tug (Starting Point) 

Selected Space Tug Aerobraking Configuration 
Concepts (1st Phase) 

Selected Space Tug Aerobraking Configuration 
Concepts (Add-On Activity) 

Space Tug Aerobraking Study Schedule. 

Comparison of Conventional and Aerobraking 
Trajectory Profiles 

Payload Potential for Space Tug Aerobraking 
Geosynchronous Payload Capability of Aerobraked Tug 
Drag Coefficients (SO 0 ! 45° and 60° Flares) 

Drag Coefficients (Basic, Short 60° Flare and Large 
Nose Flare) 

Baseline Radiative Aft Heat Shield 
Baseline Flare Concept (60° Flare) 

Short 60° Flare Concept 
Large Nose Flare Concept 

Space Tug Aerobraking Return Time from Synchronous 
Orbit 

Space Tug Aerobraking Return from Synchronous Orbit 
(With Dispersed Atmosphere) 

RCS Propellant Consumption 

Comparison of Maximum Equilibrium Temperatures 

Maximum Equilibrium Temperatures 

Maximum Equilibrium Temperatures for Basic and 60° 

Flare Configurations - Nominal and (+) Density 

Atmosphere 

Radial Perigee Position Uncertanties (Basic Tug 
Configuration) 

Navigation Update History 

Orbital Correction Burn, Delta Velocity RCS Fuel 
Consumption, and Radial Perigee Position Error 
Aft Heat Shield Weight Vs. Number of Passes 
Flare Weight Vs. Number of Passes 
Thermal Protection System Weight Vs. Number of 
Passes 

Astrionics System Weight Vs, Number of Passes 
Total Tug Weight Vs. Number of Passes 
H-33 Configuration External H2 Tank Orbiter 
(Alternate Mission Capability) 

Synchronous Round Trip Payload Vs. Number of 
Passes (200 N.M. Recovery) 

Round Trip Payload Vs. Number of Passes 
(Depart and Recover at 200 N.M.) 

Atmospheric Anomaly and Navigation Error Effect 
on Round Trip Payload (10 Pass Mission) 


PAGE 

1-a 

1-4 

1-6 

1-7 

1-8 


1-10 

1-12- 

2-jJ 


2-6 

2-9 

2-11 

2-12 

2-13 

2-15 

2-17 

2-19 

2-20 

2-22 


2-23 

2-26 

2-27 

2-28 

2-32 

2-33 

2-35 

2-36 

2-37 

2-39 

2-40 

2-42 

2-44 


x 


D5-17142 


LIST OF FIGURES (Continued) 


FIGURE 


PAGE 


2. 2. 3.0- 1 

4.0. 0.0-1 

4.1. 1.0- 1 

4. 1.1. 1- 1 

4. 1.1. 1- 2 

4.1. 1.1- 3 

4. 1.1. 2- 1 

4.1. 1.3- 1 

4. 1.1. 3- 2 

4.1. 2.0- 1 

4. 1.2. 1- 1 

4.1. 2. 1- 2 

4. 1.2. 1- 3 

4. 1.2. 1- 4 

4.1. 2.2- 1 

4.1. 2.2- 2 


4. 1.2. 2-3 


4. 1.2. 3- 1 

4.1. 2. 3- 2 

4. 1.2. 4- 1 

4. 1.2. 4- 2 

4.1. 2. 5- 1 

4. 1.2. 5- 2 

4. 1.2. 5- 3 

4.1. 2.5- 4 

4. 1.2. 5- 5 

4.2. 1.0- 1 

4. 2.2. 1- 1 

4.2. 2.1- 2 

4. 2.2. 1- 3 

4.2.2. 1- 4 

4.2.2. 1- 5 

4.2.2. 1- 6 

4 . 2 . 2 . 1 - 7 

4.2.2. 1- 8 

4 . 2 . 2 . 2 - 1 


Round Trip Payload Capabilities of Conventional 

and Aerobraked Tugs 

Performance and Trade Study Logic 

Basic (No Flare) Configuration 

Drag Bridging Scheme Illustration 

Final Basic (No Flare) Tug Drag Characteristics 

Preliminary Basic (No Flare) Tug Drag 

Characterise cs 

Basic (No Flare) Tug Static Stability 
Characteristics 

Basic (No Flare) Pressure Coefficient 
Distribution 

Maximum Local Pressure Profile (Basic-No Flare) 

Flare Sizing Study - Simplified Technical Approach 
Ellipsoid Nose Bow Shock Profile 
Embedded Flow Field Characteristics 
Parametric Flare Drag Characteristics 
Parametric Flare Static Stability Characteristics 
Flared Tug Configuration(s) Geometry 
Preliminary Variation of Time to Decay with 
Respect to Ballistic Coefficient (W/C^A) 

C D and W/C^A Range for Aerobraked Space Tug 
Configurations 

Drag Coefficients for Selected Flared Configurations 
Drag Coefficients For Optional Flared Configurations 
Coefficient Of Normal Force Slope For Flared 
Configurations 

Center Of Pressure For Flared Configurations 
Local Pressure Coefficients Over Heat Shield And 
Cylinder (Flared Configurations) 

Local Pressure Coefficients Over Flares 
Maximum Local Pressure On Heat Shield Nose 
Maximum Local Pressure At Outer Edge Of Flare 
Maximum Local Pressure Profile (30 Pass Mission- 
Flared Configurations) 

Conventional Space Tug Configuration 
Aft Heat Shield 

Internal Cap Heat Shield Concept 
Clam Shell Aft Heat Shield Concept 
Single Pivot Point - Total Elliptical Dome 
Retraction Concept 

Double Pivot Point - Clam Shell Concept 
Single Pivot Point - One Piece Cap Concept 
Single Pivot Point - Clam Shell Concept 
Payload End Heat Shield Concept 
Sidewall Insulation - 60 Degree Flare Space Tug 
Configuration - 30 Pass (5 Day Mission) 


2-45 

4-2 

4-5 

4-8 

4-9 


4-10 

4-12 

4-13 

4-17 

4-18 

4-19 

4-20 

4-21 

4-21 

4-22 

4-24 


4-24 

4-27 

4-28 

4-29 

4-30 

4-32 

4-33 

4-35 

4-36 


4-45/4-46 

4-48 

4-49 


4-50 

4-52 

4-53 

4-54 

4-56 

4-58 


xi 


D5-17142 


LIST OF FIGURES (Continued) 


FIGURE 


PAGE 


4.2. 2. 3- 1 

4.2. 2. 3- 2 

4.2. 2.4- 1 

4.3. 1.0- 1 

4.3. 2.0- 1 

4.3. 3. 0- 1 

4. 3. 3. 0- 2 

4.3. 3. 0- 3 

4. 3.3. 0- 4 

4.3. 3.0- 5 

4.3. 3.0- 6 

4.3. 3.0- 7 

4. 3. 3.0- 8 

4.3. 3.0- 9 

4.3.3.0- 10 

4.3.3.0- 11 

4.3.3.0- 12 

4.3.3.0- 13 

4.3. 4.0- 1 

4.3. 4. 0- 2 

4. 3. 4. 0- 3 

4. 3. 4. 1- 1 

4.3.4. 1- 2 

4.3.4. 1- 3 

4.3.4.1- 4 

4.3. 4. 1- 5 

4.3.4. 1 - 6 

4.3.4. 1- 7 

4. 3. 4. 1- 8 

4. 3.4. 1- 9 

4.3.4.1- 10 

4.3.4.1- 11 

4.3.4.1- 12 

4.3.4.1- 13 


Baseline Flare Concept 
Alternate Flare Concept 

Payload Adapter Concept For Flared Aerobraking 

Tug Concepts 

Mission Profile 

Preliminary Trajectory Data 

Initial Perigee Altitude 

Tug Aerobraking Return Time From Synchronous 

Orbit 

Apogee Altitude Decay 
Apogee Altitude Decay 
Perigee Altitude Decay 
Perigee Altitude Decay 
Maximum Dynamic Pressure 
Maximum Dynamic Pressure 
Maximum Dynamic Pressure 

Maximum Inertial Velocity Per Pass - Basic (No Flare) 
Configuration 

Maximum Inertial Velocity Per Pass - 60° Flare Tug 
Configuration 

Comparison of Free Space and Re-Entry Times 
(Basic - No Flare Configuration) 

Comparison of Free Space and Re-Entry Times 
(60° Flare) 

Summer Atmospheric Density Dispersions 
Winter Atmospheric Density Dispersions 
Time to Lower Apogee to 270 N.M. 

Accumulated Trajectory Correction Delta Velocity 

for Impulse at Atmosphere Exit 

Vacuum Perigee Altitude at Atmosphere Exit 

Apogee Altitude at End of Pass 

Accumulated Trajectory Correction Delta Velocity 

for Impulse at Apogee 

Vacuum Perigee Altitude at Atmosphere Exit 
Apogee Altitude at End of Pass 
Final Pass and Apogee Error for Impulse at Apogee 
Final Pass and Apogee Error for Impulse at Apogee 
Increase ir* Delta Velocity to Go to 270 N.M. for 
Impulse at Exit 

Increase in Delta Velocity to Go to 270 N.M. for 
Impulse at Apogee 

Maximum Dynamic Pressure for Basic Tug Impulse at 
Exit (10 Pass) 

Maximum Dynamic Pressure for Basic Tug Impulse at 
Exit (30 Pass) 

Maximum Dynamic Pressure for Basic Tug Impulse at 
Exit (60 Pass) 


4-59/60 

4-62 

4-64 

4-68 

4-71 

4-72 

4-74 

4-75 

4-76 

4-77 

4-78 

4-79 

4-80 

4-81 

4-82 

4-83 

4-86 

4-88 

4-89 

4-90 

4-92 

4-94 

4-95 

4-96 

4-97 

4-98 

4-99 

4-101 

4-102 

4-103 


4-104 


4-105 

4-106 

4-107 


xii 


D5-17142 


LIST OF FIGURES (Continued) 

figure page 

4.3.4.1- 14 Maximum Dynamic Pressure for Large Flare Tug 

Impulse at Exit (10 Pass) 4-108 

4.3,4*1-15 Maximum Dynamic Pressure for Large Flare Tug 

Impulse at Exit (30 Pass) 4-109 

4.3.4.1- 16 Maximum Dynamic Pressure for Large Flare Tug 

Impulse at Exit (60 Pass) 4-110 

4.3.4.1- 17 Maximum Dynamic Pressure for Basic Tug Impulse 

at Apogee (10 Pass) 4-111 

4.3.4.1- 18 Maximum Dynamic Pressure for Basic Tug Impulse 

at Apogee (30 Pass) 4-112 

4.3.4.1- 19 Maximum Dynamic Pressure for Basic Tug Impulse 

at Apogee (60 Pass) 4-113 

4.3.4.1- 20 Maximum Dynamic Pressure for Large Flare Impulse 

at Apogee (10 Pass) 4-114 

4.3.4.1- 21 Maximum Dynamic Pressure for Large Flare Impulse 

at Apogee (30 Pass) 4-115 

4.3.4.1- 22 Maximum Dynamic Pressure for Large Flare Impulse 

at Apogee (60 Pass) 4-116 

4. 3. 4. 2- 1 Elapsed Time Trend from Maximum Inertial Velocity 

to Perigee (Fitst Pass) 4-118 

4. 3. 4. 2- 2 Sensitivity of First Pass Inertial Velocity to 

Atmosphere - Basic Tug (No Flare) Configuration 4-119 

4. 3. 4.2- 3 Sensitivity of Fifst Pass Inertial Velocity to 

Atmosphere (60° Flare Configuration) 4-120 

4.3. 5.0- 1 Vacuum Perigee History of the First Pass of a 

Ten Pass Mission 4-121 

4.3. 5.0- 2 Time to Lower Apogee to 270 n.m. 4-123 

4.3. 5. 0- 3 Altitude Time History of the First Pass of a Ten 

Pass Trajectory 4-124 

4. 4. 1.0- 1 RCS System Limit - Cycle Under Worst Case In- 

fluence of Aero-Moment 4-126 

4.4. 1.0- 2 Phase Plane Trajectory Showing Tug Oscillations 

From Time lug Enters Atmosphere Until It Passes 

Max Q 4-128 

4. 4. 2. 0- 1 Pitch Axis Propellant Consumption Due to the 

Aero-Moment for the Basic Tug 4-129 

4. 4. 5. 0- 1 Pitch Axis Propellant Consumption Due to Limit 

Cycling for the Basic and 60° Flare Tugs 4-132 

4.4. 5. 0- 2 Roll Axis Propellant Consumption Due to Limit 

Cycling for the Basic Tug 4-133 

4. 4. 5. 0- 3 Roll Axis Propellant Consumption Due to Limit 

Cycle for the Space Tug With 60° Flare 4-134 

4. 4. 5. 0- 4 RCS Propellant Consumed For the Basic Tug 

During Change of Orbit Period 4-135 

4.4. 5.0- 5 RCS Propellant Consumed for Tugs With Flares 

During Change or Orbit Period 4-136 




D5-17142 


LIST OF FIGURES (Continued) 

FIGURE PAGE 

4. 4.5. 0- 6 Effects of Atmospheric Density on RCS 

Propellant Consumption for the Basic Tug 4-137 

4. 4. 5.0- 7 Effects of Atmospheric Density on RCS 

Propellant Consumption for the 60° Flare 

Tug 4-138 

4. 5. 2. 0- 1 Maximum Heating Rates - Basic Configuration 4-141 

4. 5.2. 0- 2 Maximum Heating Rates - 30" Flare Con- 

figuration 4-142 

4. 5. 2. 0- 3 Maximum Heating Rates - 45° Flare Con- 

figuration 4-143 

4. 5. 2. 0- 4 Maximum Heating Rates - 60° Flare Con- 

figuration 4-144 

4. 5. 2. 0- 5 Comparison of Maximum Heating Rate 4-145 

4. 5.2. 0- 6 Heating Rate Distribution - 10 Pass Basic 

Configuration 4-146 

4. 5.2. 0- 7 Heating Rate Distribution - 10 Pass 30° 

Flare Configuration 4-147 

4. 5. 2. 0- 8 Heating Rate Distribution - 10 Pass 45° 

Flare Configuration 4-148 

4. 5. 2. 0- 9 Heating Rate Distribution - 10 Pass 60° 

Flare Configuration 4-149 

4.5.2.0- 10 Maximum Heating Rates for Basic and 60° 

Flare Configurations - Normal and (+) 

Density Atmosphere 4-150 

4. 5. 3. 0- 1 Maximum Equilibrium Temperatures - Basic 

Configuration 4-151 

4. 5. 3. 0- 2 Maximum Equilibrium Temperatures - 30° 

Flare Configuration 4-152 

4. 5. 3. 0- 3 Maximum Equilibrium Temperatures - 45° 

Flare Configuration 4-153 

4.5. 3.0- 4 Maximum Equilibrium Temperatures - 60° 

Flare Configuration 4-154 

4.5. 3.0- 5 Comparison of Maximum Equilibrium Tempera- 

tures 4-155 

4.5. 3.0- 6 Maximum Equilibrium Temperatures for Basic 

and 60° Flare Configurations - Nominal and 

(+) Density Atmosphere 4-157 

4. 5. 4. 3- 1 Space Tug Sidewall Cross-Section 4-158 

4. 5. 4. 3- 2 Space Tug Sidewall Insulation Thickness - 

Basic Configuration 4-159 

4. 5. 4. 3- 3 Space Tug Sidewall Insulation Thickness - 

30° Flare Configuration 4-160 

4. 5. 4. 3- 4 Space Tug Sidewall Insulation Thickness - 

45° Flare Configuration 4-161 

4.5. 4. 3- 5 Space Tug Sidewall Insulation Thickness - 

60° Flare Configuration 4-162 

4. 5. 4. 4- 1 Space Tug Sidewall Thermal Protection System 

Weights 4-164 


D5-17142 


FIGURE 

4.6. 1.0- 1 

4.6. 3.1- 1 

4.6. 3.2- 1 

4.6.4.2- 1 

4. 6.4. 4- 1 

4. 6.4. 5- 1 

4. 6. 4. 5- 2 

4. 6. 4. 5- 3 

4. 6.4. 5- 4 

4. 6.4. 5- 5 

4. 6. 4. 5- 6 

4.6. 4.5- 7 

4. 6. 4. 5- 8 

4.6. 4.5- 9 


4. 6. 4.6- 1 

4.6. 4.6- 2 


4. 6. 4. 6- 3 

4.6.4. 7- 1 


4.6. 4.7-2 


4. 6. 4. 7- 3 

4.6.4. 7- 4 

4.6.5. 1- 1 

4.6. 5. 2- 1 

4.6.5. 2- 2 

4.6.6. 1- l 
4. 6. 6. 5-1 

4.6. 7.2- 1 
4.6.8. 1-1 


LIST OF FIGURES (Continued) 


Astrionlc Study Approach 
Astrionic System Aerobraking Weight 
Delta Summary 

Perigee Navigation Position Uncertainty 
Summary 

Typical Navigation Sensor Characteristics 
Synchronous Orbit Analysis 
Navigation Update History 
Navigation Position Uncertainties (10 day - 
60 Pass) 

RCS Perigee Position Uncertainty (Basic 
Tug Configurations - 60 Pass) 

RCS Apogee Position Uncertainties (Basic 
Tug Configuration) 

Radial Perigee Position Uncertainties 
(Basic Tug Configuration) 

RSS Apogee Position Uncertainties (Flared 
Tug Configuration) 

Radial Perigee Position Uncertainties 
(Flared Tug Configuration) 

One Pass Navigation Position Uncertainties 
One Pass Mission Apogee Altitude as a 
Function of Delta Deboost Velocity (Basic 
Tug Configuration) 

Deboost Burn Uncertainty Analysis 
Orbital Correction Bum,4V RCS Fuel 
Consumption, and Radial Perigee Position 
Uncertainty 

Radial Position Uncertainty as a Function 
of Corrective Burn Timing 
Effects on Navigation Due to Variation of 
Navigation Update Rate During Landmark 
Tracking Prior to Perigee 
Effects of Navigation Due to Variation of 
Navigation Update Rate During Landmark 
Tracking after Perigee 
RSS Navigation Position for Various Navi- 
gation Update Sample Rates 
Second Apogee RSS Position Uncertainty Vs. 
Time of Reinitialization of Kalman Filter 
Tug Astrionic System Configuration 
Nominal Mission Weight and Power 
Weight and Power Summary 
Failure Rates of Critical Components 
Aerobraking Redundancy Weight Impacts 
Aerobraking Power Weight Deltas 
Summary of' External and Internal Van Allen 
Radiation Environments for Three Proposed 
Mission Profiles 


PAGE 

4-165 

4-168 

4-170 

4-173 

'•4-176 

4-179 

4-181 

4-182 

4-183 

4-184 

4-186 

4-187 

4-188 


4-190 

4-193 


4-194 

4-196 


4-197 


4-198 

4-199 

4-201 

4-203 

4-204/205 

4-206 

4-208 

4-210 

4-212 


4-214 


XV 


D5-17142 


LIST OF FIGURES (Continued 1 


I FIGURE PAGE. 

4. 6. 8. 2- 1 Prescribed Shielding Configuration 4-216 

4. 6. 8. 2- 2 Electron and Bremsstrahlung Dose/Unit 

f Fluence for a Van Allen Electron Fluence 4-217 

i 4. 7. 1.0-1 Materials Selection Criteria 4-224 

! 4.7. 2.0-1 Comparison of Maximum Equilibria Temperatures 4-225 

} 4. 7. 2. 0-2 RL-10A-3-8 Engine Maximum Allowable Tempera- 

l ture Limits for Non-Operating Engine 4-227 

i 4.7. 3.0-1 Material Properties 4-228/240 

3 4. 7. 3.0-2 Radiative Material Properties 4-241/242 

l 4.7. 3.0-3 Strength to Density Properties of High 

Temperature Structural Materials 4-243 

4. 8. 1.0- 1 Space Tug Mass Properties 4-247 

4.8. 2. 0- 1 Aft Heat Shield Materials Summary 4-248 

4.8. 2.0- 2 Aft Heat Shield Weight Vs. Number of Passes 4-249 

4. 8. 2. 0- 3 Flare Weight Vs. Number of Passes 4-250 

4. 8. 2. 0- 4 Payload Adapter Weight Vs. Number of Passes 4-251 

4.8. 2.0- 5 Weight Summary - Payload Adapter/Flared 

Skirt Support 4-252 

4.8.2. 0- 6 Total Structural Weight Vs. Number of Passes 4-253 

4.8. 3.0- 1 Thermal Protection System Weight Vs. Number 

of Passes 4-255 

4.8.4. 0- 1 Astrionics System Weight Vs. Number of Passes 4-257 

4.8. 5.0- 1 Reaction Control System Weight Vs. Number of 

Passes 4-258 

4.8. 6.0- 1 Total Tug Weight Vs. Number of Passes 4-259 

5.1. 0.0-1 Conventional (Non-Aerobraking) Space Tug 

Configuration 5-2 

5. 1.1. 1- 1 Effect of Aerobraking Modification Weights on 

Round Trip Payload 5-4 

5. 1.1. 1- 2 Effect of Aerobraked Return AV on Round Trip 

Payload 5-6 

5. 1.1. 1- 3 Effect of Aerobraked Return 4V on Placement 

Payload 5-8 

5. 1.1. 1- 4 Effect of Aerobraked Return 4V on Retrieval 

Payload 5-9 

5.1. 1.2- 1 Atmospheric Entry Weight 5-10 

5. 1.1. 2- 2 Initial Weight and Propellant Weight 5-11 

5. 1.1. 2- 3 Effect of RCS Usage on Payload 5-13 

5. 1.1. 3- 1 Mission Delta Velocity Sensitivity to Final 

Orbit Altitude 5-14 

5. 1.1. 3- 2 Effects of Circularization Altitude on Payload 

(Light Weight Aerobraking Kit Penalty - 1000 
| Pounds) 5-16 

! 5.1. 1.3-3 Effect of Circularization Altitude on Payload 5-18 

; 5. 1.1, 3-4 Effect of Tug Recovery Altitude and Recovery 

Method on Delta Payload 5-19 

5.1. 2.0-1 Selected Space Tug Aerobraking Configuration 

Concepts 5-21 


.s xvi 

j ■ 


■i 


D5-17142 


LIST OF FIGURES (Continued) 

FIGURE PAGE 

5. 1.2. 0- 2 Fully Fueled Weight Statement 5-22 

5. 1.2. 0- 3 H-33 Configuration External H 2 Tank Orbiter 

(Alternate Mission Capability! 5-23 

5.1.2. 1- 1 Mission Total Delta Velocity Sensitivity to 

Plane Change Angle and Configuration 5-25 

5. 1.2. 1- 2 Payload Sensitivity to Delta Velocity 5-26 

5. 1.2. 1- 3 Total Mission Delta Velocity Equivalent Vs. 

Number of Passes 5-28 

5. 1.2. 2- 1 Aerobraking Weight Vs. Number of Passes (Basic - 

270-100 n.m. Recovery) 5-29 

5. 1.2. 2- 2 Aerobraking Weight Vs. Number of Passes 

(Basic - Depart and Recover at 200 n.m.) 5-31 

5. 1.2. 3- 1 Aerobraking 'Weight Vs. Number of Passes (30° 

Flare - 270-100 n.m. Recovery) 5-32 

5. 1.2. 3- 2 Aerobraking Weight Vs. Number of Passes (30° 

Flare - Depart and Recover at 200 n.m.) 5-33 

5. 1.2. 4- 1 Aerobraking Weight Vs. Number of Passes (45° 

Flare - 270-100 n.m. Recovery) 5-35 

5. 1.2. 4- 2 Aerobraking Weight Vs. Number of Passes (45° 

Flare - Depart and Recover at 200 n.m.) 5-36 

5. 1.2. 5- 1 Aerobraking Weight Vs. Number of Passes (60° 

Flare - 270-100 n.m. Recovery) 5-37 

5. 1.2. 5- 2 Aerobraking Weight Vs. Number of Passes (60° 

Flare - Depart and Recover at 200 n.m.) 5-38 

5. 1.2. 6- 1 Round Trip Payload Vs. Number of Passes (270 - 

100 n.m. Recovery) 5-39 

5. 1.2. 6- 2 Round Trip Payload Vs. Number of Passes (De- 

parture and Recovery at 200 n.m.) 5-40 

5. 1.2. 6- 3 Round Trip Payload Vs. Number of Passes (270 - 

100 n.m. Recovery) 5-42 

5. 1.2. 6- 4 Synchronous Round Trip Payload Vs, Number of 

Passes 5-43 

5. 1.2. 6- 5 Placement Payload Vs. Number of Passes (270 - 

100 n.m. Recovery) 5-44 

5. 1.2. 6- 6 Retrieval Payload Vs. Number of Passes (Depart 

and Recover at 200 n.m.) 5-46 

5. 1.2. 6- 7 Geosynchronous Payload Capability of Aerobraked 

Tug : 5-48 

5. 2. 1.0- 1 Radial Position Uncertainty at Perigee 

5. 2. 1.0- 2 Range of Atmospheric Perturbations 5-49 

5. 2. 1.0- 3 Maximum Equilibrium Temperatures Vs. Initial i 5-51 

Perigee Error (1st Pass of 30 Pass Mission) 5-52 

5. 2. 2. 0- 1 Fully Fueled Weight Statements for High Density 

Atmosphere 5-54 

5.2. 2. 0- 2 Round Trip Payload Vs. Number of Passes (Depart 

and Recover at 200 n.m.) .• 5-55 

5. 2. 2. 0- 3 Atmospheric Anamoly and Navigation Error Effect i 

on Round Trip Payload (10 Pass Mission) i 5-57 


xvil 


D5-17142 


LIST OF FIGURES (Continued) 

FIGURE PAGE 

5. 3. 1.0- 1 Payload Capability Comparison - Aerobraked 

vs. Conventional Tug 5-60 

5. 3. 2. 1- 1 Comparison of Round Trip Payload Sensitivities 

to Mass Fraction 5-62 

5. 3. 2. 2- 1 Comparison of Round Trip Payload Sensitivities 

to Specific Impulse 5-64 

5. 3. 2. 3- 1 Comparison of Round Trip Payload Sensitivities 

to Mission Delta Velocity 5-65 

A-l. 0.0-1 Selected Space Tug Aerobraking Configuration 

Concepts (Add-On Activity) A-3 

A-l. 0.0-2 Geosynchronous Round Trip Payload A-4 

A-2. 1.0-1 Preliminary Flare Configuration/Net Payload 

Sensitivity Estimate A-7 

A-2. 1.1-1 Basic (No Flare) Pressure Coefficient Distribution A-9 

A-2. 1.2-1 Short 60° Flared Tug Configuration(s) Geometry A-10 

A-2. 1.2-2 Drag Coefficients For Short 60° Flare #1 A-l 1 

A-2. 1.2-3 Drag Coefficients For Short 60° Flare #2 

(Selected Configuration) A-l 2 

A-2. 1.2-4 Coefficient of Normal Force for Short 60° Flare #2 A-13 

A-2. 1.2-5 Center of Pressure for Short 60° Flare #2 A-14 

A-2. 1.2-6 Local Pressure Coefficients Over Short 60° 

Flare § 2 A-l 5 

A-2. 1.2-7 Drag Coefficients for Short 60° Flare #3 A-16 

A-2. 1.3-1 Drag Coefficients for “Ring Tail" Flare A-18 

A-2. 1.3-2 Candidate Configurations for W/CgA = 2 PSF A-19 

A-2. 1.3-3 Drag Coefficients for Nose Flare (Selected 

Configuration) A-20 

A-2. 2. 1-1 Ablative Heat Shield Concept A-22 

A-2. 2. 3-1 Short 60° Flare Concept A-24 

A-2. 2. 3-2 Large Nose Flare Concept A-25 

A-2, 2. 3-3 Large Nose Flare Deployment/Retraction A-26 

A-2. 3. 1-1 Short 60° Flare Maximum Dynamic Pressure A-30 

A-2. 3, 1-2 Short 60° Flare Initial Perigee Altitudes A-29 

A-2. 3. 3-1 Equivalent M/CqA Values A-32 

A-2. 4. 0-1 RCS Propellant Consumption A-34 

A-2. 5, 1-1 Heating Rate Distribution - Basic Configuration 

(2 Pass) A— 35 

A-2. 5. 1-2 Heating Rate Distribution - 60° Short Flare 

(2 Pass) A- 36 

A-2. 5. 1-3 Heating Rate Distribution - 60° Short Flare 

(30 Pass) A-37 

A-2. 5, 1-4 Heating Rate Distribution - Large Flare (2 Pass) A-39 

A-2. 5.2-1 Maximum Equilibrium Temperatures A-39 

A-2. 5. 3-1 Heat Shield Ablative Material Thickness A-41 

A-2. 5. 3-2 Space Tug Sidewall Insulation Thickness A-42 

A-2. 5. 3-3 Thermal Protection System Weight Summary A-43 

A-2. 7. 0-1 2 Pass Maximum Heating Rates A-44 

A-2. 7. 0-2 ESA-3560 UA Properties A-46 

A-2. 7. 0-3 ESA*3560 IIA Thickness Vs. Heat Input A-47 


xviii 


D5-17142 


LIST OF FIGURES (Continued) 

FIGURE PAGE 

A-2. 7.0-4 L-605 Cobalt Alloy Material Properties A-48 

A-2.8.0-1 Mass Properties for Add-On Activity A-50 

A-2. 8. 0-2 Aerobraking Kit Weights A-51 

A-2. 8.0-3 Materials Used for Aerobraking Kit Components A-52 

A-2. 8.0-4 Total Tug Weight Vs, Number of Passes A-53 

A-3. 1.0-1 Atmospheric Density Variations for Space Tug 

Aerobraking Studies A-55 

A-3. 1.0-2 Ten Pass Basic Tug - Target Apogee Vs. Pass 

Number A-56 

A-3. 1.0-3 Final Apogee Error Vs, Number of Passes to 

270 n.m. • A-57 

A-3. 1.0-4 Increase in AV to Go to 270 n.m. for Off- 

Nominal Atmosphere A-59 

A-3. 2.0-1 Sensitivity of Total Mission Navigation 

Error Correction Delta Velocity to Correction 
Burn Time A-60 

A-3. 3.0-1 Fully Fueled Aerobraking Tug Weight Statements A-62 

A-3. 3.0-2 Round Trip Payload Capabilities A-63 

A-3. 3.0-3 Fully Fueled Weight Statement for 30 Pass 

Short 60° Flare Tug (#3) A-65 

A-3. 3.0-4 Flare Configuration/Net Payload Sensitivity 

Estimate A-66 

A-3.4.0-1 Round Trip Payload Capabilities of Conventional 

and Aerob raked Tugs A-68 

A-3. 4. 0-2 Payload Placement Capabilities of Conventional 

and Aerobraked Tugs A-70 

A-3,4.0-3 Payload Retrieval Capabilities of Conventional 

and Aerobraked Tugs A-71 

B-l Comparison of Least Squares and Maximum 

Likelihood Estimation B-l 

B-2 Comparison of Straight Position 

Resets with Kalman Position Resets B-3 

C-l Autonomous Navigation Modes C-4 

F-l. 0.0-1 Light Weight Large Flare Concept F-3 

F-l .0.0-2 Aerobraking Kit Weight Statement for Two Pass 

Airmat Flared Configuration F-2 

F-l. 0.0-3 Candidate Light Weight Large Flare Concepts F-4 

F-2. 0.0-1 Round Trip Payload Capabilities F-5 


xix 


05-17142 


A 

ACT 

AM 

ANS 

A R 

BITE 

BOM 


% 

C M 

C N c 

C P 

CP 

CPU 

D(Dr) 

EOS 

FMF 

FOV 

h 


I 

Isp 

IBM 

IMU 

10 

IPEP 

ITT 

K 

K N 

KFT 

KLB 

m(H) 

L/D 


ABBREVIATIONS AND ACRONYMS 

Cross-sectional area 
Acquisition, Control and Test (Unit) 

Astrionics Module 

Autonomous Navigation Simulation (computer program) 

2 

Reference area (constant @ 154 FT ) 

Buil t-In-Test-Equi pment 
Basic Operating Memory 

Zero angle of attack drag coefficient = Drag/q w A R 
Pitching moment coefficient « Pitching moment/qe* A R D R 
Initial normal force coefficient slope = ©C^/eo/«s=o 
Local pressure coefficient = (P^-P*®)/q» 

Total vehicle center of pressure = C^/C^ Od^o 
Central Processor Unit (computer) 

Tug diameter (14 FT), also deadband 

Earth- to-Orbit Shuttle 

Free Molecular Flow 

Field-of-view 

Geometric altitude 

Hydrogen 

Inertia 

Specific impulse 

International Business Machines Corporation 
Inertial Measurement Unit 
Input/Output (computer) 

Inertial Platform Error Program (computer program) 
International Telephone and Telegraph Company 
Bridging parameter 

Knudsen Number (assumed herein « M//"P£) 

Conversion factor for C R at max loads 
Kilofeet (thousands of feet) 

Kilopounds (thousands of pounds) 

Kilowatt (Hour) 

Lift-to-drag 


XX 


D5-17142 


Jt 

M 

MEV 

MHz 

MOSFET 

MSFN 

N 

N 

NM 


OOS 

P 

PM 

PRA 

q» Q 

R 

* 

R 

R e 

RADS (Si )m 

RGS 

RE© 

RMS 

RSS 

S 

SA 

Sync 

T 

TBC 

TPS 

USB 

USBE 


ABBREVIATIONS AND ACRONYMS (Continued) 

Radius of vehicle 

Mach number = V*® /ambient sound speed 

Million electron volts 

Megahertz 

Metal -Oxide-Silicon/Field Effect Transistor 
Manned Space Flight Network 
Normal (Position), also number of thrusters 
Normal (velocity) 

Nautical Mile 
Oxygen 

On-board maneuvering systems 

Orbit-to-Orbit Shuttle 

Static pressure 

Propulsion Module 

Patrick Reference Atmosphere 

Dynamic pressure =1/2 pVoo 2 » also heating rate 

Bow shock radius = f(x), also gas constant, or radial (position) 

Radial (velocity) 

Reynolds number = pVD//f 

Radiation energy deposited in silicon 

Reaction Control System 

Equatorial radius 

Root mean square 

Root sum square 

Molecular speed ratio = V/vT2rT 
Standard 1962 Atmosphere 
Synchronous 

Temperature (absolute), also tangential Cposition/velocity), 

also time or minimum firing time 

The Boeing Company 

Thermal Protection System 

Unified S-Band 

Unified S-Band Equipment 


xxi 


D5-17142 


V 

Mt 


W PLD 

w/c d a 

W t 

X 

Y 


a 

y 

d 

AV 

A 

A* 

V 

P® 

P 

a 

* 

6 


c$ 

EQ 

1 

L 

r 

s 

w 

OB 


ABBREVIATIONS AND ACRONYMS (Continued) 

Velocity 

Weight of inerts 
Weight of propellant 
Weight of payload 
Ballistic coefficient 
Wei ght 

Axial position coordinate measured from 2:1 ellipsoid nose apex 

Vertical position coordinate measured normal to the Tug 
centerl i ne axi s 

Thermal accommodation coefficient, also angle of attack 

Ratio of specific heats 

Partial derivative, also flare angle 

Delta velocity 

Mean Free molecular path 

Mass fraction 

Coefficient of viscosity 

Gravity constant 

Density 

Standard deviation, also Stefan-Boltzman Constant 

Attitude 

Attitude rate 

Earth rotation constant 

Surface emissivity 


SUBSCRIPTS 


Center of gravity 
Equilibrium 

Incident flow characteristics 
Local condition 
Reflected flow characteristics 
Conditions behind normal shock 
Wall surface characteristics 
Ambient or free stream characteristics 


xxii 


D5-17142 


SECTION I - INTRODUCTION 


1.0 GENERAL 

This report describes the results of a study to investigate the feasibility 
and practicality of applying an aerobraking trajectory mode to the Space 
Tug for return to low earth orbit from geosynchronous orbit. This return 
scheme will reduce the overall mission delta velocity requirements and 
will allow for a smaller propulsion module to accomplish these missions. 

The Space Tug, consisting of the smaller propulsion module coupled with 
the astrionics module, payload adapter and payload, will conform with the 
Space Shuttle dimensional and weight capabilities and will be able to 
perform almost all ('^95 percent) of the proposed geosynchronous missions 
with a single Shuttle/single Space Tug launch per mission. No orbital 
assembly and/or orbital propellant transfer operations would be required. 
The aerobraking Space Tug concept will provide an economical ground based 
Space Tug system. 

The Space Tug is one of the new hardware elements required to support the 
Integrated Space Program. It must interface with the Space Shuttle, Space 
Station, Nuclear Shuttle, Orbiting Propellant Stations, Satellites and 
other payloads. In operating with these elements, it must operate in both 
a manned and unmanned mode, over a wide range of missions. These missions 
include low earth orbital resupply, earth orbital operations, geosynchro- 
nous, lunar orbit, lunar surface, translunar and Interplanetary. 

The versatility requirements necessitate that the Space Tug have different 
modules and/or kits to provide flexibility without Imposing undue per- 
formance penalties on the basic Space Tug. This study defined the special 
aerobraking kit elements and assessed their impact on the Space Tug 
performance. 

As the study was limited in scope, it was not possible to analyze all of 
the mission implications, operational modes, environmental effects and 
design options. However, the major factors were identified and their 
influence defined. No major problems appear to exist in applying the 
aerobraking re-entry trajectory mode to the Space Tug. 

Additional Space Tug study activities are required to bring the Aerobraking 
Space Tug missions, mission modes and configuration options to the same 
level of knowledge as the other Space Tug conceptual study options. Prin- 
ciple follow-on activities should -include: Mission analyses, operational 
modes, economic analyses, payload/Tug interfaces, Shuttle/Tug interfaces, 
more detailed Space Tug aerobraking design studies and impact of the 
aerobraking modes on the advanced technology Space Tug configurations. 

In performing this study, areas where new technology is required were 
also determined. Principle follow-on technical activities required 
include: Aerodynamics and aerothermodynamics studies, materials and 
structural concepts, drag configuration concepts, and astrionics and 
control systems.. 


' 1-1 


t 


D5-17142 


1 . 0 - 


(Continued) 


This report Is divided Into five sections: This Section I - Introduction, 
Section II - Summary , Section III - Groundrules, Guidelines and Assumptions, 
Section IV - Performance and Trade Results, and Section V - Sensitivity 
Analysis. In addition, there are six appendices: Appendix A - Two Pass 
Aerobraking Space Tug Analysis, Appendix B - Brief Description of Kalman 
Filtering, Appendix C - Analysis Programs, Appendix D - Navigation Com- 
ponents, Appendix E - Recommended Aerobraking Follow-On Activities, and 
Appendix F - Two Pass Light Weight Large Flare. The Appendix A activities 
present the results of tne contract add-on study. 


l.T 


STUDY OBJECTIVES AND APPROACH 


The objective of this study was to determine the feasibility of applying 
an aerobraking trajectory mode to the Space Tug for return from geosyn- 
chronous orbit. This objective was achieved by: 

o Defining sensitivities of Tug weights to various re-entry 
environments 

o Defining sensitivity of re-entry environments to trajectory 
anomalies 

o Determining sensor and sensor accuracy requirements 
o Determining position and velocity update requirements 

o Developing astrionics reliability weight and performance 

requi rements 

o Defining the impact of the radiation environment on the 
astrionics systems 

o Determing representative inert weight penalties associated 
with aerobraking (i.e., thermal and aerodynamics) 
o Comparing required Space Tug gross weights for equal payloads 
for conventional and aerobraking trajectories 
o Determining scar weights for aerobraking kit modifications 
o Identifying the new technology implications of the aerobraking 
concept 

o Identifying the follow-on Tug aerobraking activities 

To accomplish these objectives, the study logic as shown in Figure 
1.1. 0.0-1 was followed. The first phase of the study was directed 
towards maximizing the payload potential of the Aerobraked Tug. The 
second phase add-on activity was directed towards identifying the payload 
capability with the minimum mission duration. 

The Space Tug configuration developed by The Boeing Company for NASA/MSFC 
under a prior study (Reference 1.1. 0.0-1) consisted of a propulsion module, 
an astrionics module and a payload adapter which* when combined with the 
payload, matched the former Space Shuttle cargo limit of 54,000 pounds. 

For this study, the current Space Shuttle cargo limit of 65,000 pounds 
was used to update the Space Tug. The propulsion module was increased 
to reflect this change (as shown in the propulsion module weights in 
Figure 1.1. 0.0-2). The astrionics module was updated to reflect Shuttle- 
era technology. This updating reduced the astrionics weight from 2526 
pounds to 1960 pounds. The other modules and kits were unchanged in the 


1 - 2 - 


STUDY INPUT DATA 


i 

to 


® gkqundrul.es & 

GUIDELINES 


• PRE-PHASE A 
SPACE TUG STUDY 
RESULTS 


• PRELIMINARY TUG 
CONFIGURATIONS 
& MASS PROPERTIES 


TASK I 


TRAJECTORY & 
AERODYNAMICS 
ANALYSIS 


•TRAJECTORIES 

♦ AERODYNAMIC 

CHARACTERISTICS 

♦ MASS PROPERTIES 

♦ CONTROL REQM'TS. 

« A/B DESIGN CRITERIA 





SENSITIVITY 
ANALYSIS 


e OPT. WEIGHT/PLD 
« A/B VS. CONVENTIONAL 
PERFORMANCE 
® TRADES OF MISSION 
TIME, PERIGEE ALT., 
WEIGHTS, RCS, ETC. 


CONFIGURATIONS 
& MATERIAL 
SELECTION 


• DOME CONCEPTS 

• FLARE CONCEPTS 

• ACTUATION SYSTEMS 

• MATERIAL PROPERTIES 

& DESIGN CRITERIA 

• MATERIAL SELECTION 


TASK VI 


CONCLUSIONS 
& 

RECOMMENDATIONS, 


® RESULTS 
® CONCLUSIONS 
& RECOMM. 

• REPORTS 



• TRAJECTORY ACCURACY REQM’TS. 

• ERROR ANALYSIS 

« ASTRIONICS CONFIGURATION DEFINITION 
» RELIABILITY ANALYSIS 
® SYSTEM SELECTION & WEIGHTS 

• SYSTEM PARAMETRIC SENSITIVITIES 


FIGURE 1.1. 0.0-1: AEROBRAKING FEASIBILITY STUDY LOGIC 


STUDY OUTPUT DATA 


& FEASIBILITY OF SPACE 
TUG AEROBRAKING 


® NEW TECHNOLOGY 
REQUIREMENTS 


© FURTHER AEROBRAKING 
STUDIES 


© FURTHER SPACE TUG 
AEROBRAKING STUDIES 


D5-17142 





WEIGHT ESTIMATE (LBS) 


ASTRIONICS 

MODULE 


TANK 


14'0” D!A 




TANK y \ 


^ 


U 


® PROPULSION MODULE 

• PROPELLANT 
® INERTS 

® ENGINE 639 

s PROP/MECH 801 

• THERMO/MICRO 573 

• STRUCTURE 2912 
o CONT. & RESID. 943 

® ASTRIONICS MODULE 

® ELECTRICAL 515 

® AVIONICS 842 

• STRUCTURE 300 

® THERMAL 303 

• PAYLOAD ADAPTER 

TOTAL SPACE TUG WEIGHT 
MASS FRACTION 
SPECIFIC IMPULSE (SEC) 


■RL-10A-3-8 ENGINE 
(NOZZLE RETRACTED) 


45,000 

5,868 


FIGURE 1.1. 0.0-2: CONVENTIONAL TUG (STARTING POINT) 


D5-17142 


D5-17142 


1.0 (Continued) 

baseline Space Tug configuration. 

From this uprated Space Tug configuration, six aerobraking configuration 
concepts were selected (Figure 1.1. 0.0-3 - 1st phase and Figure 1,1, 0.0-4 - 
2nd phase). These aerobraking concepts were based on nozzle end first 
re-entry and employed a 2:1 elliptical domed heat shield mounted over 
the engine as an aft end radiative or ablative thermal protection system. 

In addition to the aft heat shield, five of the six selected aerobraking 
configurations reduced the ballistic coefficient by increasing the aero- 
dynamic drag with a flare. Five different flare lengths and four different 
angles were selected to examine the vehicle static stability and drag 
characteristics. Aerodynamics, trajectory and control analyses were 
performed on these configurations. The aerodynamic analyses provided 

(1) airloads data for determining pressure loads on the structure, 

(2) drag data for determining trajectory characteristics, and (3) sta- 
bility data for determining reaction control system requirements. Thermal 
environmental data was generated which were used to define and select 
materials and provide design criteria for the aerobraking kit elements. 

The trajectory data provided input data to the astrionic analysis to 
permit accuracy and navigation error evaluation, redundancy requirements, 
sensor systems definition and selection, astrionics module configuration 
update, and weights. The impact of the structural modifications, thermal 
protection system requirements, astrionics system modifications and 
reaction control systems weights as a function of mission duration was 
defined. Trades of mission time, perigee altitude, weights, operational 
modes, etc., were performed. The results of the above activities were 
assessed and conclusions and recommendations were developed. 

A majority of the study activity was directed to determining the aero- 
dynamics, astrionics and thermal aspects of the aerobraking analyses. A 
small portion of the study effort was directed toward conceptual designs. 
The design activity was only performed in sufficient depth to define the 
feasibility and weights associated with each concept. Performance analysis 
show that the round trip payload capability is traded pound for pound with 
aerobraking kit inerts. Therefore, lightweight aerobraking designs are a 
necessary follow-on study activity. 

The study approach used in the first phase was to determine the feasibility 
of the aerobraking mode and to determine the mission duration which would 
maximize the payload capability of the Tug in performing a round trip geo- 
synchronous mission* The second phase (add-on activity) determined the 
impact of too pass short duration missions on the payload capability. 

The task and milestone schedule is shown in Figure 1.1. 0,0-5. The first 
phase of the study was of four months duration with two months for final 
documentation and presentations. The second phase (add-on activity) was 
of two months duration with one-half month for documentation. The study 
flow and major milestones are identified from the Task I, Performance 
and Aerodynamic Analyses, through the Task VI, Reports, Conclusions and 
Recommendations. 


1-5 


BASIC 

30 FLARE 

45 FLARE 

60 FLARE 

(NO FLARE) 

r* PAYLOAD ADAPT. 




| PAYLOAO I 


AST Rl OH ICS 


PAYLOAD 

FLARED 


ASTRJOMICS ^ 


PAYLOAD j 
ASTRIONICS f& e 



PAYLOAD 


ASTRIONICS 


1 


v V* 

V 





\~X' 


AFT HEAT SHIELD- 




FIGURE 1.1. 0.0-3: SELECTED SPACE TUG AEROBRAKING CONFIGURATION CONCEPTS (1ST PHASE) 





"* *<StZG 


BASIC SHORT 60° URGE NOSE 

wo flare) FURE FURE 



FIGURE 1.1. 0.0-4; SELECTED SPACE TUG AEROBRAKING CONFIGURATION CONCEPTS 

(ADD-ON ACTIVITY) 


D5-17142 



I 

00 




FINAL 


FINAL 

CONTRACT 

MID-TERM 

REPORT 

ADD-ON 

REPORT 

GO-AHEAD 

REVIEW 

DRAFT 

ACTIVITY 

8. REVIEW 


MUDY 

ACTIVITIES 


MONTHS JUN JUL 


TASK 1 - PERFORMANCE & 
AERODYNAMIC 
ANALYSIS 


TASK 2- SELECTION OF 

CONFIGURATIONS 
& MATERIALS 


TASK 3 - THERMAL 
ANALYSES 


TASK 4 - ASTRIONICS 
ANALYSES 


TASK 5 - SENSITIVITY 
ANALYSES 


TASK 6- REPORTS. CON- 
CLUSIONS, 
RECOMMENDATIONS 



SEP 

r^n 

■ 



DEC 


MHBV OT mmmm Mwah ~ iwi • 4M* ■ ■ 


JAN 

- 

FEB 

1 



iHH 



L I 




'T 


HL. _ 

1 


J i 

1 

L i 


i* 



r 


y 

i 

MAR 


TIME ALLOWED FOR ITERATIONS 

FIGURE 1.1. 0.0-5: SPACE TUG AEROBRAKING STUDY SCHEDULE 


D5-17142 


D5-17142 


1.2 BACKGROUND 

The previous Pre-Phase A studies of the Space Tug have shown that to 
accomplish the geosynchronous missions and other high energy earth or- 
bital type missions, it is necfessary to use either a very large single 
stage Space Tug or complex multi-stage Space Tugs (two stages or stage 
plus drop tanks). The large single stage Tug would not meet the Shuttle 
constraint of 15 foot diameter by 60 foot length and 65,000 pounds payload 
capability to 100 n.m. /28. 5° orbit. Similarly, the multistage Tug con- 
figurations cannot meet the Shuttle constraints. The multi-stage Tug 
operational mode would either deliver the stages in two or more launches 
or deliver the unfueled (or partially fueled) Tug to the departure orbit 
followed by a delivery of fuel and payloads in a second Shuttle flight. 

Both approaches would necessitate multi-Shuttle missions and assembly 
and/or refueling operations in orbit for the accomplishment of a single 
Space Tug mission. Neither approach is compatible with the desired ground 
based Shuttle/Tug mode of operation wherein the Shuttle would deliver the 
Tug and its payload to the departure orbit in a single launch. The 
Shuttle then could return to earth or wait on orbit until the Space Tug 
and the replaced payload are recovered from the geosynchronous orbit. 

An attractive alternative mode which would reduce the propulsive delta 
velocity requirements and thereby reduce the size of the propulsion 
module stage is the use of an aerobraking trajectory for the return mode. 
Figure 1.2. 0.0-1 illustrates the conventional trajectory profile versus 
the aerobraked trajectory profile. The conventional profile uses two 
ascent delta velocity burns to (1) leave the departure orbit, and (2) to 
plane change and circularize at synchronous orbit. These ascent veloci- 
ties total to approximately 14,100 feet per second. A 400 feet per second 
additional allowance provides for rendezvous and docking at synchronous 
orbit. The descent trajectory also uses two burns, one burn to depart 
from synchronous orbit and the second burn to circularize at low earth 
orbit. Similarly, a 400 feet per second allowance is required to rendezvous 
and dock with the Shuttle. Thus the total delta velocity requirements for 
the conventional trajectory mode is 29,000 feet per second for the geo- 
synchronous round trip mission. 

The aerobraking trajectory profile has the same ascent velocity (14,100 
feet per second) and the same rendezvous and docking velocity C 40 Q feet 
per second) requirement. However, the descent mode has a considerably 
lower delta velocity requirement. The initial deorbjt burn from synchronous 
orbit requires a 5993 feet per second delta velocity. The aerobraking 
phase is initiated with the Tug entering the atmosphere, proceeding to a 
low perigee altitude, where some of the vehicle energy is reduced by the 
atmospheric drag. Subsequent passes have lower apogee altitudes and 
slightly lower perigee altitudes. When the apogee altitude of the last 
pass is equal to the desired circularization orbit altitude, a second 
bum circularizes the Tug orbit. For this study* this altitude was 
selected as 270 n.nn (Lower altitudes, such as 200 n.m., are desirable 
ways to increase the Tug payload capabilities and may be used.) The Tug 
then phases with the Shuttle and with two additional delta velocity burns 
descends to a Shuttle/Tug rendezvous orbit and circularizes. An additional 


Ol-l 


CONVENTIONAL PROFILE 


GEOSYNCHRONOUS 
ORBIT \ 
19,300 N.M. \ 



4 V BUDGET 

0 ASCENT 14,500 FPS 
© DESCENT 14.500 FPS 


AEROBRAKING RETURN 


/^TRANSFER,. 

/oyro 


->c 


GEOSYNCHRONOUS 
ORBIT \ 
19,100 N.M. ' 


\ 


EARTH 


L~ /PERIGEE 2 
I © U* 50 N.M. 


X 


DOCK WITH \ 


A \ 

\ DECREASEviA \\ 

1 APOGEE 

10 EACH ORBIT / 

/ l CIRCULARIZE J 

/ # 270 NA*. / DEORBIT 

/ / / 


J 


s 


/ 


/ 


RETURN 4V BUDGET RANGE: (30 PASSES) j 

MINIMUM MAXIMUM 

DEORBIT 5993 FT/SEC 

CIRCULARIZE 133 » 100 N.M. 

TRANSFER TO 100 N.M. - 
DOCKING 

RESERVES 324 


5993 FT /SEC 
380 @ 270 N.M. 
654 

400 

1371 


TOTAL 


6450 FT/SEC 


8800 FT /SEC 


FIGURE 1.2. 0.0-1: COMPARISON OF CONVENTIONAL AND AEROBRAKING TRAJECTORY PROFILES 


D5-17142 



DS- 1 7142 


1.2 (Continued) 

400 feet per second provides for Shuttle/Tug rendezvous and docking. 

Using the aerobraking mode, the return delta velocity requirements may be 
as low as 6450 feet per second (20,950 feet per second for total round 
trip delta velocity) or may be as high as 8800 feet per second for the 
return mode (23,300 feet per second for the total round trip). This 
total delta velocity range is dependent on the Tug circularization altitude 
and the selection of the vehicle to perform rendezvous and docking opera- 
tions. The delta velocity requirements are reduced by 23 to 28 percent 
of that required by conventional trajectory modes. 

The lower delta velocity requirements for the aerobraking return mode will 
result in a smaller propulsive stage than that required to accomplish a 
mission using conventional trajectory flight modes. The use of the aero- 
braking mode is more advantageous in missions requiring payload retrieval 
or round trip payload where the returning vehicle is heavy than in the 
case of a placement payload where the Tug returns light (no payload). 

For the retrieval or round trip payload missions, the Tug propulsive stage 
required with an aerobraking return mode is only 50 to 60 percent as large 
as that required using conventional return modes. For a placement mission 
the aerobraking mode Tug would have to be approximately 80 percent that of 
the conventional return mode. 

Figure 1,2. 0.0-2 Illustrates the advantages of aerobraking over a conven- 
tional trajectory by comparing the round trip payload capability of each 
trajectory mode. In each example the Space Shuttle 65,000 pound cargo 
capability Is used to capacity. The center two bars reflect the com- 
parison of current state-of-the-art Tug design. The conventional Tug 
can place under 1000 pounds in equatorial synchronous orbit as compared 
to 4050 pounds for the aerobraked Tug. The aerobraked Tug's mass fraction 
(including aerobraking kit components) is only 0.833 as compared to 

0. 875 for the conventional stage. 

The first two bars represent the impact of a low cost Tug using unsophis- 
ticated design concepts. The conventional Tug cannot deliver any round 
trip payload to synchronous orbit. The aerobraked Tug can deliver 3000 
pounds with a stage having a mass fraction of 0.85 (the mass fraction 
including aerobraking kit components is 0.819). 

The last two bars represent the impact of the NASA point design. The point 
design can deliver 3000 pounds of round trip payload with a stage having a 
mass fraction of .895 and an Isp of 470 seconds. This stage requires a 
major advancement in the state-of-the-art 0976). An aerobraked stage with 
the same mass fraction (0.895) and with a propellant loading of 50,300 
pounds (so that stage and payload do not exceed 65,000 pounds gross weight) 
will round trip 6500 pounds payload. 

The conclusions that can be drawn from this comparison are: 

1. If monetary constraints do not permit the technology to be developed 
or if the technology does not meet the desired goals* aerobraking 


1-11 


ROUND TRIP PAYLOAD* (1000 OF POUNDS) 


0-1— 


BASIS: TUG WEIGHT PLUS PAYLOAD WEIGHT = 65,000 LBS. 

MASS FRACTIONS INCLUDE MAIN STAGE 
AND RCS PROPELLANTS 


W p = 50.780 # 
W, = 8,980# 

A STG = 8 ‘ 8 * 
W,. |T _ 2,240* 


KIT = 
*A/B 


= 0.819 


l sp = 460 SEC 


}■%:£=» 1 NO PAYLOAD 

[*': ^ x*l CAPABILITY 
{>:<: ^ FOR COMPARABLE 

[ i# £ SvJ MASS F R ACT I ON 
E$0 m CONVENTIONAL 
l&: ce o «3| TUG 




i:x : o m 1 

Br uj zd g:yl 

■ 

- • •• • ■* 


Wp = 50,785* 

W, = 7,927# 

A' sic = 0865 

W K)T = 2,240# 

A' A B = 0 833 
l,„ = 460 SEC. 






li§ 1 

Ifi: cn o 
E*> uj zd :* j 
fev: <C h- m 


i< i 

cn %:J 
:$ cq *:;:f 

cn O 

& UJ => X:i 


tXw/XfSSfiv 



X - 0.875 


W p = 50,300# 

W ( = 5,900# 

A' stg = 0895 

W K|T = 2,280# 

A' a /B * °’ 860 

l sp = 470 SEC 



W p = 55,550 
W, = 6,450 

A' STG = 0 895 
l» = 470 SEC 


TO 

SYNCHRONOUS 
EQUATORIAL 
EARTH ORBIT 


UNSOPHISTICATED 
TUG DESIGN 


STATE -OF-ART 
TUG DESIGN (1971) 


ADVANCED STATE-OF-THE-ART 
TUG DESIGN (1976) 


FIGURE 1.2. 0.0-2: PAYLOAD POTENTIAL FOR SPACE TUG AER0BRAKING 


D5-17142 


D5-17142 


1.2 (Continued) 

will provide a method of achieving the desired payload capability. 

2. If the technology is funded and the technology goals are achieved, 
aerobraking will provide greater payload capability (if required) 
or can deliver multiple payloads at reduced overall mission costs. 

A further advantage of utilizing the aerobraking Tug is that the conven- 
tional trajectory Tug is sensitive to variations in mass fraction, specific 
impulse, and delta velocity. Historical data indicates that inert weights 
tend to increase with the length of the program development time and 
improved definition of the configuration during the development phase. 

This factor coupled with the sensitivity makes a single, large stage 
Space Tug a high design risk concept. The use of the aerobraking mode 
would partially decrease this risk as the specific impulse and mass frac- 
tion sensitivities impact would be reduced. Further, the options as to 
which vehicle (Shuttle or Tug) performs the low earth orbit maneuvers 
offers two additional methods of reducing sensitivities and program 
risks. 

Utilizing the aerobraking return mode will result in longer missions than 
would be encountered with the conventional mode. If sufficient time is 
all owed for aerobraking, this trajectory mode can be accomplished with a 
minimal heating and with little impact on Tug design and structures. 
However, the longer the mission duration results in increased weight of 
the electrical power system, increased reaction control system fuel, and 
with increased weight for the necessary redundancy and reliability 
requirements of the astrionics system. For an aerobraked Space Tug con- 
figuration which maximizes the payload capability, it is necessary to 
define that mission duration in which the combination of decreasing 
structural and thermal penalties, are minimized. At that mission duration, 
the payload is maximized. The first phase of this study therefore defined 
the compromise in return trip time which resulted in the minimum aero- 
braked weight Tug configuration for each of the selected configurations. 

The second phase (add-on activity) minimized the mission duration to 
two (2) passes (approximately 10 hours). This shorter duration reduced 
the astrionics and control penalties. However, the more rapid return 
increased the thermal impact on the Aerobraking Tug. The results of the 
short duration missions on Tug design, performance and payload capability 
are shown in Appendices A & F. The summary section which follows presents 
the results of both the first and second phase activities and analyzses 
these results to define the total study conclusions and recommendations. 


1-13/1-14 


D5-17142 


SECTION TI - SUMMARY 

2.0 CONCLUSIONS AND RECOMMENDATIONS 

This study investigated the feasibility and practicality of the aero- 
braking mode for return trajectories of the Space Tug from geosyn- 
chronous equatorial orbit. Payloads weighing between 3000 and 4000 
pounds can be carried in a round trip mode to and from an equatorial 
geosynchronous orbit using a Space Tug (exclusive of payload) weighing 
approximately 55,000 pounds. The aerobraking Space Tug payload capa- 
bilities for placement or retrieval missions will be approximately 
two-and-one-half times the round trip payload. As shown in Figure 
2. 0.0. 0-1, this is sufficient payload capability to perform 95% of the 
prognosticated round trip geosynchronous missions in a mode wherein 
a single Shuttle flight can deploy and retrieve the Tug and its 
round trip payload. The aerobraking mode may also be applied for re- 
turn from other high energy missions to provide larger payload capa- 
bilities than those possible wi th similar sized Tugs operating with 
conventional trajectory modes. 

The conclusions reached in this study are, due to the limited study 
scope and time available, preliminary and provide trends rather than 
detailed data. However, the results present ample justification for 
the recommendation of further aerobraking study activity and technology 
programs. Many of the conclusions are subject to re-analysis as the 
aerobraking technique level of knowledge becomes comparable to con- 
ventional trajectory techniques and as the on-going studies further 
define the Shuttle. 

The general study conclusions and recommendations are contained in the 
subsequent paragraphs. The conclusions for each study discipline are 
listed at the end of the summary discussion of the particular dis- 
cipline (Sections 2.1.1 through 2.2). The recommended follow-on 
activities presented herein are further discussed in Appendix E. 

General Conclusions - The general conclusions reached in the study are: 

o The aerobraked Tug's payload capability is maximized by missions 
having 25 to 35 atmospheric passages during the aerobraking 
phase. This corresponds to total Tug geosynchronous mission 
time of from 4 to 7 days. A 5 day mission duration is within the 
on-orbit capability of the Shuttle and permits a single Shuttle/ 
Tug to accomplish a mission. 

o A one day return mission from geosynchronous orbit can be accom- 
plished in from one to five passes. However, the thermal and 
pressure environments Increase the structural requirements and 
result in significantly tower payload capability than the longer 
duration, maximum payload missions. 


PAYLOAD WEIGHT <K LBS) PAYLOAD WEIGHT (K LBS) 


ROUND TRIP PAYLOAD 



BASIS: 

• EOS CAPABILITY ~ H-33 ORBIT HR 

• 287 GEOSYNCHRONOUS MISSIONS 

• 30 PASS MISSIONS 

CONCLUSIONS: 

®95% OF ALL GEOSYNCHRONOUS 
MISSIONS CAPTURED 


CUMULATIVE PERCENT OF PAYLOADS 


PAYLOAD PLACEMENT 



CUMULATIVE PERCENT OF PAYLOADS CUMULATIVE PERCENT OF PAYLOADS 




FIGURE 2.0. 0.0-1: GEOSYNCHRONOUS PAYLOAD CAPABILITY OF AEROBRAKED TUG 


05-17142 


D5-17142 


2.0 (Continued) 

o The maximum geosynchronous payload capability of the aerobraked 
tug can be obtained by optimizing the departure/recovery orbits 
and maximizing usage of the Shuttle for Shuttle/Tug interface 
operations. 

o Comparing the required weights for aerobraked and conventional 
trajectory Tugs to accomplish comparable payload geosynchronous 
missions, the aerobraked Tug weight is approximately 55% 
(retrieval), 65% (round trip), or 80% (placement) that of the 
conventional Tug. 

o The aerobraking kits for the Space Tug can be designed so that 
the aerobraked Tug will fit within the Shuttle's cargo bay. 

The aerobraking kits have a negligible scar weight impact on 
the conventional trajectory Tug. When the kits are removed, the 
Tug may be used for its lower energy missions with insignificant 
reduction in performance. 

o Reducing the ballistic coefficient with a large flare or other 
large surface area drag devices will permit lower thermal and 
pressure loads at reentry. Obtaining this large area, however, 
will reduce the weight available for payload and presents many 
design problems with packaging in the Shuttle cargo bay, deploy- 
ment, retraction, astrionics visibility and payload rendezvous 
and docking to the Tug. 

o In general, short duration aerobraking missions will require 
more complex designs of the aerobraking kit elements and will 
require technology advances in materials to increase payload 
capabilities. 

o A radiative heat shield is more desirable than the ablative heat 
shield as it is lower weight, reusable with minimal and/or no 
refurbishment and Is less complex. 

o The atmospheric anomalies may be overcome by trajectory correc- 
tion techniques. The thermal effect was less than 100°F. 

o The solar, lunar, and earth harmonics perturbations significant- 
ly impact the selection of the target perigee altitudes but have 
only minor impacts on thermal, aerodynamics, and control para- 
meters. These effects are generally predictable and can be 
accounted for in pre-mission planning. 

Recommendati ons 


The results of this study are indicative of the aerobraking potential. 
The study did not (1) fully investigate all of the parameters which 
could potentially increase or reduce the weight aerobraking kits, or 
(2) examine sufficient aerobraking configuration options, or (3) de- 


2-3 



D5-17142 


2.0 (Continued) 

fine the operational modes for an optimum performance/cost system. 

The economic advantages of aerobraking due to fewer required Shuttle 
launches, however, are obvious and represent a potential for a major 
reduction in space program costs. These cost savings were not studied 
and should be assessed in future studies. 

More detailed studies should be completed to develop the design and 
operational detail of the aerobraked Space Tug concept to a level 
comparable with that of the Space Tug configurations previously in- 
vestigated or presently under investigation. Such follow-on studies 
should refine and update the Tug configurations considering the 
evolving Shuttle era technology, the total mission model, optimal 
operational modes, Shuttle/Tug/payload interfaces and economic con- 
siderations. 

The major supporting technology programs should include (1) wind 
tunnel testing of aerobraking configuration options, (2) investigation 
of alternative aerobraking kit concepts, (3) further investigation of 
navigation errors and correction techniques to define guidance laws 
and targeting, and (4) further identification of potential atmospheric 
anomalies. 

2.1 SUMMARY OF TECHNICAL STUDY RESULTS 

The primary approach for this study was to determine the round trip 
payload capability as a function of the return trip time (number of 
aerobraking passes) for each of six specified aerobraking adaptions 
(prior Figures 1.1. 0.0-3 and -4) to a baseline Space Tug configura- 
tion (prior Figure 1.1. 0.0-2). A range of trajectory return times 
from 0.4 of a day to 11 days (2 to 60 passes) was analyzed. The im- 
pact of the various return times were related in terms of the weight 
of the additional structures, materials, subsystems and expendables 
required for thermal protection, increased drag, aerodynamic stability, 
guidance, control, and payload protection. The following subsections 
summarize the results of the technical studies conducted. 

2.1.1 Aerodynamics Analysis 

The aerodynamics analysis was conducted to provide the required drag 
data for trajectory analysis, static stability data for control 
analysis and aerodynamic loading data for structural analysis. Drag 
data were developed for each of the six configurations over the con- 
tinuum, slip flow and free molecular flow regions. The velocities 
encountered by the aerobraking Tug will range from- 20,000 to 35,000 
feet per second. As shown in Figures 2,1, 1,0-1 and -2, the drag co- 
efficients (at perigee) varied from 1.3 for the basic (no flare) Tug 
to 44.5 for the large nose flare. At the higher altitudes, the drag 
coefficients varied from 2.75 to 3.5 for the basic Tug to 59 for the 
nose flare Tug. The wide range of drag coefficients investigated 
provided greater than an order of magnitude change in the ballistic 


2-4 


ALT~ Ktt (N.M.) 


FIGURE 2. 1.1. 0-1: 



05-17142 






















D5-17142 


2.1.1 (Continued) 

coefficient (from approximately 72 for the no flare configuration to 
approximately 2 for the large nose flare configuration). Thus, a 
significant change in configuration could be assessed as to its im- 
pact on control, thermal, astrionics and structural requirements and 
upon performance and payload capabilities. 

The center of pressure and normal force coefficients were determined 
as a function of altitude and velocity. These data were used for de- 
fining the control requirements. At perigee, the basic (no flare) 
and the short 60° flared Tugs have their centers of pressure located 
forward of their centers of gravity and therefore are aerodynamically 
unstable. To offset this instability, the reaction control system 
will be required to provide a controlling moment. The 30°, 45° and 
60° and large nose flared configurations have their centers of 
pressure located either at the center of gravity or further aft and 
therefore are stable. The coefficients of normal force were defined 
over the same altitude and velocity range as the drag coefficients and 
were used to define the aerodynamic moments of each of the configura- 
tions. 

The 30° flare is statically stable at 11.25 feet and offers slight 
improvement in its drag coefficient over the basic Tug. The 45° and 
60° flares achieve static stability at approximately the same flare 
slant length (approximately 9 feet). At flare angles under 45°, the 
flare slant length required for stability increases rapidly. 

The local pressure coefficients were defined for each of the six con- 
figurations. These data were then used to determine the local pressure 
loads at the nose, sidewall and, where applicable, on the flares. 

The data show an order of magnitude drop in pressure from the heat 
shield nose to the sidewall. The flare configurations had lower loads 
than the basic (no flare) configuration and increasing loads with 
flare angle. Increasing the mission duration, decreased the pressure 
loads. The basic (no flare) configuration pressure loads were approx- 
imately one psi on the heat shield for the two pass mission and de- 
creased to approximately .1 psi for a 60 pass mission. Similar data 
were developed for the remainder of the configurations. 

The short 60° flare (4.9 1 slant height) has approximately the same 
drag characteristics as the 30° flare but because of its short length, 
the configuration is statically unstable. The 60° flare, however, 
provides a much larger drag coefficient increase over basic flare 
than the 45° flare does (approximately Cn = 6.6 for the 60° flare 
versus approximately C D - 1.9 for the 45“ flare) for the same flare 
slant height of 14 feet. 


D5-17142 


2. 1,1.1 Aerodynamics Conclusions 


o The basic (no flare) and short 60° flared Tugs have their centers 
of pressure located forward of their centers of gravity and are 
unstable. The other flared configurations have their centers of 
pressure equal to or aft of their centers of gravity and are 
stable. 


o The pressure loads on the basic (no flare) Tug decrease an order 
of magnitude from the nose of the heat shield to the cylindrical 
skirt joint and decrease another near order of magnitude down 
the length of the Tug, 


o 


The pressure loads at the end of the steep flare are comparable 
to the heat shield nose loads and significantly impact the weight 
of the flare. 


2.1,2 Configurations 

The prior Figures 1.1. 0.0-2, -3 and -4 showed the conventional tra- 
jectory Tug configuration and the six aerobraking Tug configurations. 
The externally mounted aerobraking kits required to modify the con- 
ventional to the aerobraking trajectory configurations consist of (1) 
an aft heat shield, (2) sidewall insulation, (3) a flare (as applica- 
ble), and (4) a payload/flare adapter. With these kits installed, the 
aerobraked lug will fit within the 15 x 60 feet Shuttle cargo bay. In 
addition, the reaction control system and the astrionics module are 
impacted. These impacts are discussed in Section 2.1.4 and 2.1.6, 
respectively. 

Aft Heat Shields 


Figure 2. 1.2. 0-1 shows the 2:1 ellipsoidal radiative aft heat shield 
concept which was designed to provide protection for the aft end of 
vehicle. It is used on the basic (no flare) configuration, and on the 
30°, 45°, short 60° and 60° flared configurations for mission durations 
of 5-60 passes. The removable cap and its actuation system are also 
used on the large nose flare for the 2 pass mission. The aft heat 
shield is composed of a fixed dome section (mounted to the aft skirt 
of the propulsion module) and a removable cap. The removable cap is 
emplaced over the engine during the aerobraking phase and during 
transport of the Tug within the Shuttle's cargo bay. It is rotated 
outward during normal Tug operation to provide clearance for the main 
engine exhaust during the main engine burns. The cap is operated by 
an electric motor driving two gears and is latched and sealed while 
emplaced. 

The ablative aft heat shield is used on the 2 pass basic (no flare) 
and short 60° flare Tugs. This ablative heat shield is composed of 
a solid dome (no removable cap). It is actuated similar to the 
radiative concept described above but the motor and gears are located 
on the sidewall. 


2-8 






2.1.2 (Continued) 
Sidewall Insulation 


D5-17142 


The sidewall insulation covers the Tug's sidewalls (and payload adapter 
for the no flare and short 60° flare configurations). It is installed 
over the Tug's aluminum micrometeoroid shield and maintains a 400° F 
temperature limit at the micrometeoroid shield surface. The insulation 
consists of a microquartz insulation covered with a titanium or L-605 
metallic outer foil. 

Flare 


The aerodynamic flare kit provides static stability and increased drag 
for the 30°, 45° and 60° flared configurations. The flare consists of 
an inconel frame and facing sheet with titanium support struts and a 
spring/cable actuation system (Figure 2. 1.2. 0-2). The flare is folded 
around the vehicle (total diameter approximately 14.5 feet) until after 
the deorbit burn from geosynchronous orbit. Prior to the first atmos- 
pheric passage, the flare is extended by spring hinges and is supported 
by the struts. After the last atmospheric passage, the cable system 
retracts the struts and collapses the flare against the Tug sidewall 
so that the Tug can fit within the Shuttle’s cargo bay for return to 
earth. 

The short 60° flare as shown in Figure 2. 1.2. 0-3 is similar to the 
larger 30°, 45° and 60° flares in that the same type of panel system 
is used and the same materials may be used. However, the short flare 
allows a simpler retraction/actuation system. This sytem uses threaded 
rods and followers to elevate the struts and flare. A reversible drive 
motor, a drive chain and 12 drive sprockets (one per support strut) are 
used to actuate the rods and followers. 

The large flare as shown in Figure 2. 1.2. 0-4 is unique in that it is 
located forward of the reentry Tug and is combined with the heat shield 
into a continuous forward drag flare. The large panels present ex- 
tremely difficult actuation/ retracti on problems when coupled with the 
desirability to fit the Tug with the folded flare into the Shuttle 
cargo bay. (Appendix F contains data on Light Weight Large Flare.) 

Payload/ FI are Adapter 

The multipurpose integrated payload/flare adapter is used as a payload 
container, flare mounting fixture and a flare actuation system housing 
structure. It is an aluminum stiffened structure with a guide cone 
and guide tubes that assist with the payload docking operations and 
with payload holddown. The flared configurations have a flare actua- 
tion section located within the payload adapter for mounting the cable 
retraction system. The flare and the aluminum skin of the adapter 
provide the thermal protection for the payload. Both the basic (no 
flare) Tug and the short 60° flare Tug configurations have sidewall 
insulation and an Insulated end cap to provide thermal protection for 
the payload. 


2-10 


S1QEVEW0F 
EXTENOED FLARE 



SIDE VIEW OF 
RETRACTED FLARE 


FIGURE 2.1. 2. 0-2: BASELINE FLARE CONCEPT (60° FLARE) 











D5-17142 




D5-17142 


2. 1.2,1 Configuration Conclusions 

o The Tug's aerobraking kit can be designed to fit (while attached 
to the Tug) within the Shuttle's 15 x 60 feet cargo bay constraint. 

o Aft-end reentry is preferable to payload-end reentry because of 
the payload's greater sensitivity to environment. The highest 
temperatures occur at the reentry end of the vehicle. If the 
payload were at the reentry end, it would require a heavy com- 
plex protection system to accommodate variable length payloads. 

2.1.3 Trajectory Analysis 

For each of the configurations, trajectory analyses were conducted 
using the drag coefficient data developed in the aerodynamic analysis. 
The trajectory scheme used the conventional trajectory profile and 
delta velocity to leave the earth departure orbit and to achieve geo- 
synchronous orbit. For the return trajectory, the Tug deorbits using 
an initial target perigee altitude selected for a desired mission 
return time. Subsequent passes lower apogee altitude significantly 
while reducing the perigee altitude slightly. When the desired cir- 
cularization altitude is reached, an engine burn circularizes the 
orbit. 


The return trip time is a direct function of the initial perigee alti- 
tude for any fixed ballistic coefficient (configuration weight in- 
cluding payload divided by the product of the configuration frontal 
area and drag coefficient: W/CpA). 


Figure 2. 1.3. 0-1 shows the relationship of trip time to initial perigee 
altitude for the various configurations. As shown in the figure, the 
initial perigee altitude may be higher for the flared configurations 
for accomplishing the missions in the same return time than that of the 
non-flared configurations because of the higher drag (lower W/C D A) 
values obtained with the flared configurations. These higher perigee 
altitudes will result in lower temperatures and pressure loads than 
will be encountered with the basic (no flare) configuration. 


The impact of atmospheric dispersion was defined for the basic (no 
flare) and 60° flare Tug configurations. Figure 2. 1.3. 0-2 presents 
the Initial perigee altitude required for decay to a 270 NM orbit for 
the 1962 Standard Atmosphere and a constant High Density and constant 
Low Density Atmospheres, The solid curves of the figure are a result 
of re-isolation of the initial perigee altitude required to force the 
trajectory to a final apogee of 270 NM in the presence of the dis- 
persed atmosphere. The data illustrates the range of entry times due 
to atmospheric dispersions if the vehicle flies an uncorrected tra- 
jectory. For example, the flight time of the basic Tug nominal 30 
pass trajectory varies from 3.6 days (19 passes) for the more dense 
atmosphere to 9.6 days (53 passes) for the less dense atmosphere. 

This large range of potential entry time implies the need for a tra- 
jectory correction technique that will significantly reduce this range. 


2-14 . 


rriAL PERIGf 










05-17142 


2.1.3 (Continued) 

Several techniques were briefly investigated including: (1) impulse 

at the time of exit from the atmosphere and (2) impulse at apogee and 
(3) impulse at atmospheric re-entry. It appears that the impulse at 
atmospheric re-entry technique has greater payload potential as it may 
combine navigation correction with atmospheric dispersion correction, 
but this area should be studied in detail in a follow-on study for 
aerobraking guidance scheme analysis. 

The atmospheric dispersion data used in Figure 2. 1.3. 0-2 to impact 
mission performance was developed by the NASA early in this study 
activity and represented a conservative estimate. Updated data for 
atmospheric dispersion versus time was developed by the NASA. This 
latter data when applied to the aerobraking analysis indicated that 
the dispersion effects would be considerably less pronounced (approxi- 
mately half as severe) as that shown in Figure 2. 1.3. 0-2. 

The impact of lunar, solar and earth harmonic perturbations on the 
aerobraked Tug trajectory was assessed. Variable initial perigee 
altitudes will be required for the perturbed environment versus the 
initial perigee altitude computed using a spherical earth. These 
perturbations are (1) generally predictable, (2) operationally signi- 
ficant, and (3) have only minor impacts on the thermal, aerodynamic, 
and control results. The mission planning phase must account for 
these effects by considering the target perigee altitude to be 
selected and the phasing required for Shuttle/Tug rendezvous. 

2. 1.3.1 Trajectory Conclusions 

o The use of aerobraking can reduce the return delta velocity from 
5700 to 8050 ft/sec. 

o For a specified mission duration, a lower ballistic coefficient 
(high drag with flare) will allow higher initial perigee alti- 
tudes which result in lower thermal environments and lower 
pressure loads. 

o The aerobraking Tug is in the sensible atmosphere (600,000 feet 
altitude) for approximately 3 % of the time. 

o Near-constant trip times can be maintained in a dispersed at- 
mosphere by small correction burns prior to each atmospheric 
passage (approximately 200 ft/sec. total for a 30 pass mission). 

o Lunar, solar, and earth harmonics perturbations require different 
initial perigees than for the spherical earth but do not signi- 
ficantly change the thermal, air load, and control parameters. 


2-16 



INITIAL PERIGEE ALTITUDE - 1000 FT 


DISPERSED ATMOSPHERE 
NOMINAL ATMOSPHERE 
(1962 STANDARD) 


i y 


60° FLARE 


A } BASIC TUG 


n PASSES TO LOWER 
160 APOGEE TO 270 N.M. 


o L/D = 0 

® ORBIT INCLINATION = 0 
© ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
e INITIAL APOGEE ALT = SYNCH. ALT. 

» ENTRY ALT. INTO ATMOSPHERE = 600K FT. 
® FINAL APOGEE ALT. = 270 N.M. 

® + SUMMER, HIGH SOLAR ACTIVITY 
HIGHER DENSITY ATMOSPHERE 
9 - WINTER, LOW SOLAR ACTIVITY 
LESS DENSE ATMOSPHERE 


UNCORRECTED INITIAL 
PERIGEE ENTRY TIME 
RANGE FOR A NOMINAL 
THIRTY PASS TRAJECTORY 


19 PASSES 53 PASSES 


0.5 1.0 5.0 10.0 

TIME TO LOWER APOGEE TO 270 N.M. - DAYS 


FIGURE 2.V.3.0-2: SPACE TUG AER0BRAKING RETURN FROM SYNCHRONOUS ORBIT 

(WITH DISPERSED ATMOSPHERE) 


D5-17142 


D5-17142 


2.1.4 Control Analysis 

The control analyses used the center of pressure data and the co- 
efficient of normal force data developed in the aerodynamic studies 
plus the trajectory data to determine the Tug control requirements. 
These analyses were directed specifically to define the requirements 
during aerobraking considering (1) limit cycle requirement, (2) 
the aeromoment requirements, and (3) directional control requirements. 

The requirements for stabilizing the Tug were defined assuming the 
maintenance of a ± 1° angle of attack throughout the sensible atmos- 
phere. The basic (no flare) Tug is not aerodynamically stable and 
required approximately 550 pounds more RCS fuel than the aero- 
dynamically stable flared configurations. The short 60° flare is also 
unstable and requires approximately 200 pounds more RCS fuel to provide 
stability. The other flared Tug configurations are statically stable 
and therefore have significantly lower fuel requirements. Static 
stability only (no dynamic stability) was investigated. 

As shown in Figure 2. 1.4. 0-1, the control propellant requirements in- 
crease with mission duration. The RCS fuel requirements are signifi- 
cantly different for the basic Tug and short 60° flare Tug due to 
their aerodynamic instability. Fuel consumption are tabulated below 
for a 30 pass mission. 

RCS Fuel 

Tug Configuration Consumption (Pounds) 

Basic (no flare) 620 

Short 60° flare 271 

30° flare 82 

45° flare 80 

60° flare 76 

The impact of atmospheric dispersions was assessed for both the basic 
(no flare) and 60° flared Tugs. The flared Tug requires approximate- 
ly 15 pounds more fuel and is relatively insensitive to mission 
duration. The basic (no flare) Tug is sensitive to mission duration 
and requires approximately 75 pounds additional fuel for a 10 pass 
mission, approximately 30 pounds additional fuel for a 30 pass mission 
and approximately 45 pounds additional fuel for a 60 pass mission. 

2. 1.4.1 Control Conclusions 

o The fuel consumption for the unstable basic Tug is approximately 
550 pounds more than that required for the statically stable 
large flared configurations. 

o The short 60° flare provides some improvement in the static 
stability properties over the basic Tug. The RCS fuel require- 
ments are 200 pounds more than the stable configurations. 


2 -18 


RCS PROPELLANT POUNDS 


— NOMINAL CONDITIONS 
■ — — ATMOSPHERIC DENSITY INCREASE 


BASIC 

-{NO FLARE)- 



{ i 

^ BASIC ftjr> FLARE) 










/SHORT 60° FLARE, 


30° FLARE TUG- 


Ck 

2 : 

3 

O 

a- 400- 

j- 

z: 

< 

-i 

£* 300- 


u? 

cc 200- 


ASSUMPTIONS 

• PITCH, YAW, AND ROLL ATTITUDE IS CONTROLLED 
THROUGHOUT PERIOD OF ORBIT CHANGE. 

® RCS SYSTEM IS DESIGNED WITH: 

1) 5° DEADBAND IN PITCH/YAW (1° WITHIN 
ATMOSPHERE); 2.5° IN ROLL 

2) 4-200 LB HIGH PRESSURE THRUSTERS IN 
THE PITCH AND YAW PLANES 

3) 4-20 LB HIGH PRESSURE THRUSTERS IN 
THE ROLL AXIS 

4) PITCH, YAW, AND ROLL THRUSTER MINIMUM 
FIRING TIME IS 30 MILLISECONDS 


45 FLARE TUG 


LARGE NOSE 
FLARE 


60° FLARE TUG ■ 




0 2 10 20 30 40 50 

NUMBER OF EARTH PASSES 


20 30 40 

NUMBER OF EARTH PASSES 


RCS PROPELLANT CONSUMED DURING 
CHANGE OF ORBIT PERIOD 


EFFECT OF ATMOSPHERIC DISPERSIONS ON 
RCS PROPELLANT CONSUMPTION 


FIGURE 2.1. 4.0-1: RCS PROPELLANT CONSUMPTION 


D5-17142 


2-20 


D 


G 


C 



SPACE TUG BASIC CONFIGURATION 


TRAJECTORY 

1 MAXIMUM EQUILIBRIUM 

1 

I 

> 

' 3 

C 


<r~ 

H 

S PASS 

3320 

5175 

1S85 

1120 

97S 

682 

10 PASS 

2990 

2860 

1410 

987 

875 

591 

30 PASS 

2540 

2420 

1166 

797 

720 

451 

60 PASS 

2240 

2140 

1005 

676 

615 

364 



SPACE TOO 30° FI ARC CONFIGURATION 


TRAI1ICTORY 

maximim roJiuyRiiw 

TIMl'FRATORCS (°F) 1 

A 

B 

C 

i) 

U 

F 

5 TASS 

2940 

2800 

1382 

1032 

1272 

1293 

10 PASS 

2630 

2520 

1215 

943 

1104 

1124 

30 PASS 

2100 

2060 

957 

735 

850 

869 

60 PASS 

1880 

1790 

809 

605 

706 

729 



SPACE TOG 45° HARE CONFIGURATION 


TRAJECTORY 

maximum rouiuBniiM 

TfrTliRATURF-S (°F) I 

A 

B 

C 

0 

I: 

F 

5 PASS 

2SS0 

2480 

1195 

927 

1037 

108S 

10 PASS 

2300 

2205 

1040 

800 

874 

919 

30 PASS 

18S6 

1793 

812 

60S 

635 

680 

60 TASS 

1660 

1580 

691 

507 

513 

558 



imiucrouY 

MAXIMUM EC? H LIBRIUM 

TlttPERAlURTS t°F) i 

A 

B 

i 

D 

L 

F 

5 PASS 

2325 

2220 

1048 

206 

943 

970 

10 TASS 

20S0 

1990 

917 

697 

804 

829 

30 PASS 

1740 

1660 

733 

540 

621 

643 

f.O TASS 

1512 

1467 

609 

435 

503 

538 


(T\ 


FIGURE 2 J .5.0-1 COMPARISON OF MAXIMUM EQUILIBRIUM TEMPERATURES 


D5-171 42 





D5-17142 


2. 1.4.1 (Continued) 

o The reaction control system fuel consumption increases with 
mission duration. 

o To reduce the fuel consumption, it is desirable to have a tight 
deadband in the sensible atmosphere and then to widen the dead- 
band in the free space environment. 

o Decreasing the minimum impulse time will decrease the fuel con- 
sumption (30 milliseconds is considered state-of-the-art). 

o The fuel consumption will increase by 10% to account for the 
atmospheric dispersions, 

2.1.5 Thermal Analysis 

Figure 2. 1.5. 0-1 shows the maximum equilibrium temperatures, at 
various points on each of four configurations, as a function of the 
number of passes required for return. These data were developed 
assuming thin metallic radiative surfaces. As shown, the maximum 
temperature decreases with increasing number of passes for the return 
mission. The maximum temperature for each configuration exists at the 
stagnation point of the aft heat shield. 

Note ; The temperatures shown are the surface temperatures of 

thin films. The effect of heat sink was considered for only (see 
Section 4.5) the basic configuration. The heat sink effects could re- 
duce the stagnation temperature approximately 300° F for a 30 pass 
mission if a hot structure is used for the aft heat shield. Because 
of the anticipated thicknesses of the micrometeoroid shielding and 
flare material, the temperature of the Tug sidewall and flare will 
approach the thin film temperatures shown. 

Figure 2. 5. 1.0-2 lists the maximum temperatures encountered on the aero- 
braking Tug configurations investigated in the add-on activity. The 
two pass basic Tug and the two pass short 60° flare Tug used ablative 
heat shields. The 30 pass short 60° flare Tug and the large flared 
Tug configurations used radiative heat shields and radiative heat 
shield/flares, respectively. The basic Tug and the short 60° flare 
Tug configurations will require a TPS for the payload adapter as the 
temperatures exceed the 300° F payload temperature limit. 

Peak heating rates always occur with the first pass and determines the 
maximum equilibrium temperature. The basic Tug's heating rate is 
approximately four times greater than the 60° flared Tug. The last 
pass has lower heating rates but for longer time periods and therefore 
impacts the insulation requirements. The time in the heating environ- 
ment is approximately 220 seconds for the first pass and will Increase 
to approximately 800 seconds for the last pass of a 30 pass mission. 

The heating durations (i.e., time in atmosphere) are relatively in- 
sensitive to configuration. 


2-21 


CONFIGURATION 


TRAJECTORY 




A 

BASIC (NO FLARE) 

2-PASS 

3880 

SHORT 60° FLARE 

2-PASS 

3290 

SHORT 60° FLARE 

30-PASS 

2120 

URGE NOSE FLARE 

2-PASS 

1410 



MAXIMUM EQUILIBRIUM TEMPERATURES 


D5-17142 
























MAXIMUM EQUILIBRIUM TEMPERATURE - °F 


D5-17142 



FIGURE 2. 1.5. 0-3 MAXIMUM EQUILIBRIUM TEMPERATURES FOR BASIC AND 60° FLARE 
CONFIGURATIONS - NOMINAL AND (+) DENSITY ATMOSPHERE 




D5-17142 


2.1.5 (Continued) 

Similar data was developed for the two pass configurations studied in 
the add-on activity. The heating rates follow the same trends as shown 
above. However, the short duration results in considerably higher 
heating rates for the basic (no flare) Tug-123 BTU/ftVsec. (first 
pass of two pass mission) versus 63 BTU/ft2/sec. (first pass of ten 
pass mission). With the use of the flare, the initial re-entry alti- 
tude is higher and therefore lower heating rates are encountered. 

Figure 2. 1.5. 0-3 illustrates the effect of atmospheric dispersions on 
the basic (no flare) and the 60° flare Tug. Only a minor increase in 
temperature occurs on the heat shield (approximately 70 Q F) with the 
"worst case" atmosphere. 

2. 1.5.1 Thermal Conclusions 

o Maximum equilibrium temperatures occur on the first pass of the 
mission and determine the radiative material temperature re- 
quirements. 

o Insulation requirements are determined by the last passes of a 
mission where total heat input is maximum, 

o The flared configurations have lower maximum equilibrium temp- 
eratures than does the basic (no flare) configuration (for equal 
mission durations). 

o Maximum equilibrium temperatures decrease rapidly by Increasing 
the mission duration from 5 to 20 passes. Temperatures decrease 
at a lesser rate beyond 20 passes. 

o The temperatures on the sidewalls and on the flares are approxi- 
mately 25-50% of those on the heat shield nose. 

o The constant High Density Atmosphere increases: 

- The maximum heat shield temperatures less than 100°F 
The maximum flare temperatures by less than 50° F 

- The sidewall Insulation weights less than 25 pounds 

- The number of passes by 4 or 5 to maintain temperatures at 
the level of the Standard Atmosphere, 

2.1.6 Astrionics Analysis 

The astrionics systems analysis was conducted by the IBM Corporation 
for The Boeing Company. The astrionics study investigated navigation 
accuracy analysis, astrionics system configuration, redundancy analy- 
sis, weight impact, radiation impact and new technology and follow-on 
study effort. 


2-24 


D5-17142 


2.1.6 (Continued) 

The navigation analysis was performed to determine the navigation 
sensors required for the aerobraking mission and to define the navi- 
gation accuracies obtained with these sensor combinations. The IMU, 
star tracker, landmark tracker, horizon sensor and laser radar were 
selected as the required navigation sensors. It was determined that 
the same complement of hardware was required for all aerobraking 
mission durations, and that only redundant hardware would be added 
to achieve reliability. 

The results of the navigation analysis indicated that the navigation 
perigee uncertainties were rather insensitive to configurations (no 
flare or flared Tug). A summary of the perigee uncertainty for the 
basic configuration is shown in Figure 2. 1.6, 0-1. This figure shows 
that the maximum 1 <r errors expected for the first pass of 2 to 15 
day missions are less than .35 nm and the steady state values 
are in the order of 0.05 nm. It should also be noted that the navi- 
gation perigee uncertainties are generally smaller for the longer 
duration mission because of the decreased atmospheric perturbations. 

Figure 2. 1.6. 0-2 illustrates the navigation update history. The hori- 
zon sensor is used for the high altitude navigation accuracy estimates. 
At lower altitudes (under 4000 n.m.), the effective operating range 
of the landmark tracker is reached and this more accurate sensor is 
used for navigation updates (1800 to 500 seconds before perigee; 300- 
500 to 1700 seconds after perigee). The star tracker is used to 
provide attitude update for correction of the platform drift of the 
IMU system and to improve navigation update accuracy. After the aero- 
braked Tug has passed through perigee, the Kalman filter is re- 
initialized because the vehicle has attained a new orbit. Landmark 
tracker updates after filter re-initialization reduces the post- 
perigee navigation uncertainties to under one mile (3 sigma). 

Figure 2. 1.6. 0-3 illustrates the position error, delta RCS fuel con- 
sumption and delta velocity corrections as functions of first pass 
correction burn time prior to perigee. The unshaded top portion of 
the figure represents the perigee errors if a correction burn is made 
based on relatively limited information (e.g., after horizon sensor 
updates). A burn at this time would result in a perigee error 
greater than the expected error with no burn (data point on the border 
between the shaded and unshaded portions). The shaded bottom portion 
of the figure represents the perigee errors if a correction burn is 
based on the more accurate landmark tracker Information. Correction 
burns made after at least 500 seconds of landmark tracking result in 
relatively small perigee errors. 

The impacts of repeated passes through the Van Allen radiation belt 
for the aerobraking mission do not appear significant at this time. 
Assuming a vehicle skin thickness of approximately .090 Inches of 
aluminum* no problems appear for times up to 10 days. The 15 day 
analysis indicates that care must be taken in the selection of silicon 


2-25 


I 

PERIGEE PASS NUME 


FIGURE 2.1.G.0-1: RADIAL PERIGEE 




2-27 


LT,ST 

UPDATE 

(AT=5 AND 10 SEC) 



/ HS - HORIZON SENSOR 

REINITIALIZE LT - LANDMARK TRACKER 

KALMAN FI LTER ST - STAR TRACKER 

AT - TIME BETWEEN UPDATES 


FIGURE 2.1. 6.0-2. NAVIGATION UPDATE HISTORY 


D5-17142 



£ 

\ 

l 


ro 

i 

ro 

co 


Cd 

o 

on 

Cl 

ill 


21 

o 

h* 


O 

CL 

m 

m 


O 

a; 

Ui 

CL 

-1 

< 

< 





FIGURE 2.1. 6*0-3: 


ORBITAL CORRECTION BURN* AV RCS FUEL CONSUMPTION, AND RADIAL 
PERIGEE POSITION ERROR 




FUEL CONSUMPTION (LBS) 



D5-17142 


2,1.6 (Continued) 

transistors in the electronics. Also, additional shielding, included 
as part of the component packaging, would decrease the effects of 
radiation. 


The redundancy analysis indicated that additional components and hard- 
ware must be added to the basic astrionic system configuration as the 
mission time increases in order to maintain a 99% mission success 
probability. 


In summary, power and redundancy provide the delta weight impacts re- 
quired for aerobraking. The weight impacts of these items and the 
total weight deltas based on aerobraking mission time are summarized 
in Section 2:1.8, Figure 2. 1.8. 0-4. 


2. 1.6.1 As trionics Conclusions 


o Autonomous navigation in synchronous orbit can limit navigation 
uncertainties to a RSS steady state accuracy of 5 n.m. (la) and 
2 ft/sec. (lo) using a horizon sensor (after 12 hours on orbit). 

o The perigee radial position uncertainties for the 5, 10 and 15 
day missions are relatively insensitive to configuration. The 
perigee uncertainties for the two day mission are slightly 
higher for the basic (no flare) configuration. 

o The autonomous navigation configuration used for this study can 
limit radial perigee uncertainties to less than 0.35 n.m. (la) 
for initial passes and to the region of 0.05 n.m. (la) steady 
state. 

o Navigation updates after perigee and at higher altitudes are re- 
quired to limit uncertainties during those periods and to pre- 
vent excessive perigee uncertainties on subsequent passes. 

o Minor midcourse corrections for navigation errors can reduce 
perigee position errors by an order of magnitude. These 
corrections should be based on the landmark tracker information. 

o Astrionics system weight increases as aerobraking mission dura- 
tion increases. Therefore, minimum weight deltas are obtained 
by minimizing aerobraking mission time. 

o Additional reactant and tanks for electrical power (1-1/2 Ibs/hr) 
and additional components to maintain acceptable astrionic system 
reliability are the prime contributors to weight Increases to the 
astrionic system due to aerobraking. 

o The radiation impact increases as the aerobraking mission duration 
increases. Radiation impacts to electronics by the Van Allen 
radiation belt appear insignificant for missions of 10 days or less. 


2-29 


05-17142 


2.1.7 Materials 

The sift heat shield and flare materials were selected to (1) withstand 
the maximum temperatures and pressure loads expected at a specific 
location on the vehicle and (2) have a minimum inert weight impact. 

The weights of the flares and aft shields shown in the succeeding 
Section 2,1.8 reflect the changes in materials selected as the envir- 
onment varied. For the lower temperatures (500 to 1800°F) Titanium 
6AL-4V, Inconel 718 and Rene' 41 were utilized. TD-nickel -chrome was 
used for the next higher temperature range (up to 2000°F). Temp- 
eratures (above 2000°F) as seen by the basic (no flare) configura- 
tion and by the short duration (5-10 pass) flared configurations' 
heat shields, require advanced state-of-the-art alloys such as 
Fansteels 85 and 60. For less than 5 passes, an ablative insulation 
over a titanium structure was used. The selected ablative was ESA-3560 1 1 A. 

The Tug's sidewall insulation material is Johns Mansville microquartz. 

This insulation has a long life temperature capability to 2000° F 
(melting point 3000°F) and has a low density (3 pounds per cubic foot). 

The insulation thickness is varied as required to maintain the Tug's 
sidewall temperature at 400° F. To cover and protect this insulation 
material, a thin radiative sheet (0.002 inches) of L-605, a Haynes 
cobalt alloy (for higher sidewall temperatures) or titanium (for tem- 
peratures under 800°F) was used. The basic (no flare) or short 60° 
flare configurations' payload adapter also uses this same microquartz/ 
titanium combination for thermal protection of the payload (300°F 
limit). 

The payload/flare adapter is protected from the environment by the 
flare or by the payload insulation. Therefore, aluminum skin and 
structure is utilized for this adapter. 

2. 1.7.1 Material Conclusions 

o State-of-the-art radiative materials may be used for the long 
( ~ 10 days) duration basic (no flare) configurations and for 
the medium ( ~ 3-4 days) duration flared Tug. Shorter mission 
durations will require advanced state-of-the-art radiative 
materials or the use of ablatives. 

o Ablatives are required for the one to four pass missions. They 
were not considered for longer multipass missions, (1) due to 
lack of data on the properties of the recycled ablative and (2) 
due to their high inert weight. 

o The aft heat shield will experience high temperatures which 
will cause unique problems in sealing and actuation. 

o Sidewall temperatures vary significantly down the length of the 
Tug. To minimize weight, a tapered insulation is desirable. 


2-30 


D5-17142 


2. 1.7.1 (Continued) 

o The temperatures experienced by the flares are such that metals 
are required except for long duration missions (above 10 days). 

2,1.8 Weights 

One of the key guidelines in the Initial study activity was to maxi- 
mize the payload capability by minimizing the weight penalty associ- 
ated with performing the aerobraking operations. In the add-on study 
effort, mission duration was minimized and aerobraking kit weight was 
not considered the overriding guideline. For each of the selected con- 
figurations, the weights of the aerobraking components were determined 
as a function of number of passes. Figure 2. 1.8. 0-1 illustrates the 
weight of the radiative and ablative aft heat shield domes as functions 
of the number of passes for each of the configurations. The radiative 
dome materials were changed from tantalum to TD-nickel -chrome as the 
temperatures encountered decrease to 2000° F and lower. The impact of 
the atmospheric density on the weight of aerobraking components was 
computed for the basic (no flare) and 60° flare configurations. The 
atmospheric density had negligible effect in the maximum payload capa- 
bility mission duration regions (25-35 passes). 

For the two pass missions with the basic (no flare) and the 60° short 
flare Tugs, an ablative heat shield was required. The heat shield 
ablative was ESA-3560 IIA mounted atop a titanium support structure. 
Note: The use of ablatives resulted in high heat shield weights. 

Similar type data was developed for the flare configurations and are 
shown in Figure 2. 1.8. 0-2. For each of the flare options, the flare 
weight was determined as a function of the number of passes. At 
approximately 30 passes, the material thickness reuuces to where it 
will be necessary to maintain a minimum thickness for handling rather 
than that required for the pressure loads and thermal environments. 

The atmospheric density effects on inert weights are significant for 
low passes (5-10 passes), but are not significant In the maximum pay- 
load capability (25-35 pass) region. 

The large nose flare is not shown in the figure as it is a unique con- 
cept which integrates the heat shield and flare into a composite 
structure. This concept employs a very large flare (72') diameter as 
compared to the other flares. The weight of this flare resulted in 
zero payload. (See Appendix F for Light Weight Large Flare Concept.) 

The payload adapter weight is different for the baseline (non- 
aerobraked) Tug, the basic (no flare) aerobraked Tug and for the flared 
aerobraked Tug* The trajectory for the non-aerobraked Tug does not ex- 
perience temperatures above 300°F andj therefore, requires only docking 
and holddown fixtures for the payload adapter components. The basic 
(no flare) and large nose flare Tugs* adapters weigh 350 pounds (ex- 
clusive of thermal protection materials). The peyload adapter for the 
30°, 45 p , short 60® and 60° flare Tugs weigh 390 pounds (exclusive of 
thermal protection materials). This additional 40 pounds accounts for 
the supports for the flare actuation system. 


2-31 


D5-17142 



FIGURE 2.1 .8.0-1 : AFT HEAT SHIELD WEIGHT VS. NUMBER OF PASSES 


2-32 



D5-17142 






D5-17142 


2,1.8 (Continued) 

The sidewall insulation weights are shown in Figure 2. 1.8. 0-3. The 
microquartz insulation thickness was varied depending on the thermal 
requirements to maintain a 400° F sidewall and a 300° F payload tempera- 
ture orr the aluminum below the insulation. The sidewall protection 
covers the cylindrical section of the propulsion module, the astrionics 
module sidewall and, where applicable, the payload adapter. For both 
thermal and handling purposes, an outer foil is mounted over the in- 
sultion. For short duration missions where temperature along the Tug 
sidewall are high, L-605 was used. For the longer duration missions, 
the sidewaT temperatures are lower and titanium may be used. 

Figure 2. 1.8. 0-4 illustrates the astrionics module weight change as a 
function of the number of passes. The changes to apply the aero- 
braking kit components did not influence the astrionics weight. How- 
ever, mission duration is significant and will effect the electrical 
power requirements and the redundancy requirements to maintain a 0.99 
reliability. 

The reaction control system propellant consumption generally increases 
with mission duration as shown in prior Figure 2. 1.4. 0-1. The basic 
(no flare) Tug and the short 60° flare Tug require significantly more 
fuel (to maintain aerodynamic stability) than the stable, 30°, 45°, 

60° and large nose flare configurations. Atmospheric dispersions 
further increase the RCS fuel requirements. 

Figure 2. 1.8. 0-5 illustrates the combined weight impact of all the 
aerobraking kit variables. The astrionics and RCS fuel weights in- 
crease with mission duration while the heat shield, flare and thermal 
protection system decrease with mission duration. The minimum inert 
weight and hence the maximum payload is obtained when the mission 
duration is between 25 to 35 passes. 

2. 1.8.1 Weight Conclusions 

o The structural and thermal aerobraking kit weights decrease with 
increases in mission duration. Mission durations with approxi- 
mately 25 to 35 atmospheric passages have the minimum aerobraked 
Tug gross weights (exclusive of payload). 

o Short duration missions have low payload capabilities due to 
large increase in aerobraking kit weights. 

o The use of radiative materials over ablative materials for the 
heat shield is desirable as lower weights are associated with 
radiative thermal systems. 

o The shorter duration missions (5-10 passes) have higher tempera- 
tures and air loads which significantly Impact flare weights. 

o The weight of steep angle and long slant height flares decrease 
rapidly by increasing mission duration beyond 5 passes. 


2-34 


05-17142 



NUMBER OF PASSES 


FIGURE 2. 1,8, 0-3: THERMAL PROTECTION SYSTEM WEIGHT VS. HUMBER OF PASSES 

2-35 


w 

i 

OS 

o> 



FIGURE 2.1. 8.0-4: ASTRIQNICS SYSTEM WEIGHT VS. NUMBER OF PASSES 


D5-17142 


TOTAL TUG WEIGHT AT START AEROBRAKING (LBS. X IIP 


D5-17142 


SHORT 60° FLARE 
BASIC (HO FLARE) 


— — + 3 a ATMOSPHERIC 
DENSITY VARIATION 
NOTE: LARGE NOSE FLARE 
HAS NEGATIVE PAY- 
LOAD AND NOT PLOTTED 


SHORT 60° FLARE 


30 FLARE 



NO FLARE 


SASIC TUG AT START AEROBRAKING 


PROPULSION INERTS 

(POUNDS) 

5,388 

REACTION CONTROL 

480 

ASTRIONICS INERTS 

1,960 

PAYLOAD ADAPTER 

200 

PROPELLANT 

1,690 

TOTAL (WITHOUT 

9,718 


AEROBRAKING) 


HUMBER OF PASSES 


FIGURE 2. 1,8. 0-5: TOTAL TUG WEIGHT VS* NUMBER OF PASSES 









D5-17142 


2. 1,8.1 (Continued) 

o Long duration mission flare thickness is dictated by handling and 
processing constraints rather than by loads or temperatures. 

o The constant High Density Atmosphere 

Has insignificant impact on heat shield weights and on 
long duration mission flare weights 

Has significant impact on short duration mission large 
flare weights (approximately 250 pounds for 5 pass 
60° flare). 

2.2 SUMMARY OF SENSITIVITY ANALYSIS RESULTS 


Figure 2. 2. 0.0-1 shows the representative Shuttle payload capabilities 
used in this study. The Shuttle can place 65,000 pounds into 100 NM/ 
28.5° circular orbit and approximately 60,000 pounds into 200 NM/28.5 0 
orbit. Using the nominal operational mode shown in prior Figure 
1.2. 0.0-1 (100 NM”™^geosynchronous“>^-270 NM 1U0 NM), the 
gross weight of the Tug and its round trip payload is approximately 
58,000 pounds. The paragraphs below identify the sensitivites to ideal 
conditions (Standard Atmosphere - no navigation error) and to realistic 
conditions (atmospheric perturbations and navigation errors). 


2.2.1 Sensitivities Of Payloads And Temperatures For Tug Config 
urations For Standard Atmosphere 


The geosynchronous round trip capability of the aerobraked Tug is 
sensitive to (1) the aerobraking kit weights and to (2) the delta 
velocities associated with the various operational modes, configura- 
tions and environments. The basic (no flare) configuration has a 
lesser aerobraking kit inert weight and a greater delta velocity re- 
quirement than do the flare configurations. When comparing the maximum 
round trip payloads among the no-flare and 30°, 45®, short 60® and 60® 
flared configurations, these two factors (weight and velocity) tend to 
equalize the capabilities of these two configuration types. 


To more nearly optimize the Shuttle/Tug combination, a Tug departure 
and recovery altitude of 200 n.m. was selected, Using this opera- 
tional mode, the gross weight of the Tug and payload is slightly less 
than 60,000 pounds and within the Shuttle's capability. This 200 n.m. 
mode substantially decreases the total Tug mission delta velocity 
budget since the requirement for the transfer to another orbit and sub- 
sequent circularization is deleted (i .e. , 270 n.m. w—w^-ioo n.m.). 
Approximately 1200 pounds of additional round trip payload capability 
is achieved by this partial optimization of the Shuttle/Tug combina- 
tion. 

Figure 2. 2. 1.0-1 shows the round trip payload capabilities of the six 
configurations as a function of the number of passes in a mission. 

The operational mode is the 200 NM Tug departure and recovery mode 
discussed above. The configurations maximize their payload capabtlt- 


2-38 


PAYLOAD - THOUSANDS OF POUNDS 



05-17142 




ROUND TRIP PAYLOAD (THOUSANDS OF POUNDS) 


D5-17142 



NEUTRALLY STABLE 

SHORT 60° FLARE 
1 / 1 

(2000°F) 

M00 o F i 

/ 

V 2000°F 

. 3i i© 

* 

r SHORT £0° 




BASIC (NO FLARE) 

& SHORT 60° FLARE 
(ABLATIVES) 




BASIS: 

© GROSS WEIGHT OF TUG & PAYLOAD AT 
DEPARTURE FROM 200 N.M. LESS THAN 
60,000 POUNDS. 

® MAIN ENGINE CIRCULARIZES TUG 

® DEPARTURE AND RECOVERY OF TUG 
AT 200 N.M. 

© MAXIMUM HEAT SHIELD STEADY-STATE 
EQUILIBRIUM TEMPERATURES 


NOTE: LARGE NOSE FLARE 

HAS NEGATIVE PAYLOAD 



NUMBER OF PASSES 


FIGURE 2. 2.1. 0-1 s SYNCHRONOUS ROUND TRIP PAYLOAD VS. 

NUMBER OF PASSES (200 N.M. RECOVERY) 





D5-17142 


2.2.1 (Continued) 

ties at approximately 30 passes and the payload capability is relative- 
ly insensitive over the mission duration range of 10-60 passes. The 
30° flare configuration has the greatest round trip payload capability 
(4225 pounds) of the six configurations studied in depth. Also shown 
on the figure is an estimate of the 30 pass payload capability of a 
neutrally stable short 60° flare. Because the 30° flare is also neu- 
trally stable, this data point indicates that near-neutral stability 
is a desirable attribute for an aerobraked Tug configuration. 

The differences in payload capability for the flared configurations 
are severely impacted by the flare length which in turn directly im- 
pacts the flare weight. The neutrally stable 30° and short 60° flare 
configurations have relatively small and lightweight flares. Their 
delta velocity requirements are approximately the same as those of 
the heavier and more stable flared configurations. Shortening the 
45° flare to achieve idle same stability as the 30° flare should in- 
crease the payload capability of this shortened flare configuration 
to be comparable to the 30° flare. Further decreasing the flare 
length of the 45° and 60° flares to achieve a comparable ballistic 
coefficient to the 30° flare may result in decreased payloads. TFis 
is typified - by the 30 pass short 60° flare shown on the figure. The 
increased RCS propellant consumption (stability) and the requirement 
to completely insulate the payload more than offset the decrease in 
flare weight. 

Also shown on Figure 2. 2. 1.0-1 are the maximum steady state heat shield 
temperatures seen by the configurations having radiative aft heat 
shields. These steady state temperatures are based on thin wall ana- 
lysis and do not include heat sink effects (see Thermal Analysis, 
Section 2.1.5 above). The 45°, 60° and neutrally stable short 60° 
flared configurations are not maximum payload limited by the use of 
TD-nickel -chrome (2000°F limit). The 30° flare configuration will 
only have a slight degradation in maximum payload capability 
(approx, 100 pounds) by the use of TD-nickel -chrome with a 2000°F 
limit. 

2.2.2 Sensitivities of Payloads and Temperatures for Tug Configura 
tions for Atmosphere Perturbations and Navigation Errors 

The Initial atmosphere perturbations data utilized in the study were 
based on a percentage range (approximately +5055 constant High Density 
Atmosphere to -4055 constant Cow Density Atmosphere) from the 1962 
Standard Atmosphere, The expected variations during the entire a§ro- 
braking mission duration or between individual passes were unknown. 

Figure 2,2. 2.0-1 shows the effects of the High Density Atmosphere on 
the payload capabilities of the basic (no flare) and 60 p flare con- 
figurations, The maximum round trip payload capability of the basic 
(no flare) configuration (approx. 3600 pounds 0 30 passes) is reduced 
less than 1055 from the Standard Atmosphere case (shown as dotted 
lines). The maximum round trip payload capability of the 6Q° flare 
configuration is reduced approximately 1255. Also shown on the figure 


2-41 


ROUND TRIP PAYLOAD {THOUSANDS OF POUNDS) 




NUMBER OF PASSES 


FIGURE 2.2. 2. 0-1 : ROUND TRIP PAYLOAD VS. NUMBER OF PASSES 

(DEPART AND RECOVER AT 200 N.M.) 



D5-T7142 



D5-17142 


2.2.2 (Continued) 

are the maximum heat shield equilibrium temperatures (steady state). 

The effect of the High Density Atmosphere is to require four or five 
additional passes in a mission to maintain the equivalent Standard 
Atmosphere temperature. 

Figure 2. 2. 2. 0-2 shows the round trip payload sensitivities (10 pass 
mission) to atmospheric anomalies and navigational errors. Shown in 
the first two hatched columns of each configuration are the payloads 
and temperatures previously discussed (Figure 2. 2. 2. 0-1). The third 
hatched column of each configuration shows the round trip payloads 
achievable under the combined effects of the High Density Atmosphere 
and the 3 sigma navigational errors (single correction burn at entry- 
after 1300 seconds of landmark tracking). Because the prior payload 
analyses have reserved 400 ft/second delta velocity for navigation 
error corrections, including the navigational errors has a minimum 
Impact on the round trip payload. Even for the relatively short 
duration 10 pass mission, the 3000 pound payload capability is re- 
tained under these "worst" conditions depicted. The temperatures 
shown in Figure 2, 2. 2.0-2 are the heat shield nose stagnation tem- 
peratures. The expected increases are only 110°F for the 60° flare 
configuration and 160°F for the basic (no flare) configuration. 

The atmospheric dispersion analysis conducted was based on a constant 
High Density and a constant Low Density Atmospheric model. This 
model results in a more severe environment than would be expected to 
be encountered. Therefore, the temperatures shown are higher than 
expected. 

A Varying Atmosphere model was received at the time the add-on 
activity commenced. Due to the objectives and scope of the add-on 
activity, only two trajectories were flown using the Varying Atmos- 
phere model. Because the Varying Atmosphere model is less severe, 
its effects on the payload capabilities are less than the High 
Density Atmosphere. For example, the ten pass basic (no flare) con- 
figuration required 180 ft/sec delta velocity to correct the trajectory 
for the High Density Atmosphere. The requirement decreased to 100 ft/ 
sec for the Varying Atmosphere. 

2.2.3 Conventions 1/Aerobraking Tug Performance Comparison 

The groundrules used for the comparisons shown in this section were 
those established for the MSFC Point Design Tug Studies. They differ 
in some respects from those used in the Aerobraking Study. Therefore, 
no direct comparisons with other aerobraking results should be made 
without first rationalizing these differences. 

Figure 2. 2. 3. 0-1 shows the round trip payload capabilities of the con- 
ventional and aerobraked Tugs. The initial MSFC Point Design goal of 
3000 pounds payload { bf - 0.895, Isp - 470 seconds, total usable 
propellant - 55,552 lbs) is shown at the top of the figure. If the 


2-43 


o 

<c 

o 


oo 

o 


0 

01 


>- 
C U. 
Q- O 

CL to 

Si 

»— c 

§3 

^ o 



4 


3 


2 


1 


0 



BAS«C (NO FLARE) 
CONFIGURATION 


60° FLARE 
CONFIGURATION 


2990° F 2080° F 

I • I lii 


BASIS: 

• MAIN ENGINE CIRCULARIZATION 

• DEPART & RECOVER AT 200 N.M. 

• MAXIMUM STEADY-STATE 
HEAT SHIELD EQUILIBRIUM 
TEMPERATURES (°F) 

• 10 PASS MISSION 


STANDARD 1962 ATMOSPHERE 
AND NO NAVIGATIONAL ERRORS 


HIGH DENSITY ATMOSPHERE 
AND NO NAVIGATIONAL ERRORS 


HIGH DENSITY ATMOSPHERE 
AND 3 a NAVIGATIONAL ERRORS 
(RSS CORRECTION 4 V) 


FIGURE 2.2. 2. 0-2: ATMOSPHERIC ANAM0LY AND NAVIGATION ERROR EFFECT ON 

ROUND TRIP PAYLOAD (10 PASS MISSION) 




D5- 17142 



MASS FRACTION ( X ) 



ROUND TRIP PAYLOAD (THOUSANDS OF POUNDS) 


FIGURE 2.2.3. 0-1: ROUND TRIP PAYLOAD CAPABILITIES GF CONVENTIONAL AND AEROBRAKED TUGS 


D5-17142 


D5-17142 


2.2,3 (Continued) 

main engine Isp were to be degraded to 460 seconds, approximately 500 
pounds of payload capability would be lost. If the conventional stage 
mass fraction were degraded (while the propellant loading and Isp are 
held constant), the payload capability will be decreased as shown. 

The conventional Tug used as a "Starting Point" for this aerobraking 
study has a mass fraction of 0.852 without the payload adapter (total 
usable propellant weight = 45,000 pounds). The 3000 pound round trip 
payload could be achieved by the aerobraked 30° flare configuration 
(30 pass mission) with the current stage mass fraction and with an 
uprated main engine having an Isp of 470 seconds. Using the current 
engine (Isp = 460 seconds), the stage would require a mass fraction of 
0.862 to attain the 3000 pound payload capability. If the aerobraked 
Tug stage were designed similar to the Point Design and with a reason- 
able scaling factor to account for the differences in propellant 
loading, the aerobraked stage might have a mass fraction of 0.875. 

This would provide a round trip payload capability of approximately 
4400 pounds . 

In addition to the capability to use state-of-art technology (lower 
mass fractions and Isps), the aerobraked Tug with its 3000 pound pay- 
load has a lower gross weight in the Shuttle. If the Shuttle's pay- 
load were to be reduced by approximately 10 percent from its 65,000 
pounds capability, the aerobraked Tug could more easily withstand the 
transition. 


2.2.4 Sensitivity Conclusions 

o The Tug's propulsion module technology requirements are signifi- 
cantly reduced by aerobraking. 

o The single Shuttle/Tug launch per geosynchronous mission is 
possible for 95% of the missions using Tug aerobraking. 

o Round trip geosynchronous payloads of 3000-4000 pounds are 
achievable by a 45,000 pound propellant aerobraked Tug. 

o Placement and retrieval of geosynchronous payloads of 7000-9000 
pounds are within the capability of the 45,000 pound propellant 
aerobraked Tug. 

o The operational mode used by the aerobraked Tug should utilize 
the Shuttle's capability to deliver and recover in orbits above 
the 100 NM earth orbit. 

o Near-neutral static stability will maximize payloads for flared 
configurations. 


2-46 


05-17142 


2.2.4 (Continued) 

o The round trip payload capability of the aerobraked Tug is 
relatively insensitive to mission times of 10-60 passes in a 
Standard Atmosphere. This payload sensitivity to mission 
duration is somewhat greater in the perturbed atmospheres. 

o The atmosphere variation and navigation error correction burns 
can be combined into one burn, thereby reducing the total delta 
velocity requirement. 

o The application of the aerobraking concept to the Space Tug will 

approximately double the round trip payload capability of 
the advanced technology Space Tug; 

permit the use of a lower mass fraction (if advanced 
technology does not meet goal) and permits the use of a 
lower specific impulse (existing engines) to deliver the 
desired payload to the high energy orbits; 

Still deliver the desired payload even if the Shuttle's 
cargo bay size decreases and a 10% loss in Shuttle 
payload capability occurs. 


2-47/2-48' 


D5-17142 


SECTION III 

GROUNDRULES, GUIDELINES AND ASSUMPTIONS 


3.0 GENERAL 

The Space Tug must interface with all of the elements in the Integrated 
Space Program. This includes: The Space Shuttle* Space Station, Nuclear 

Shuttle, other Space Tugs and variable size and weight payloads. These 
interfaces impose on the Tug (1) dimensional and weight constraints, 

( 2 ) thermal and aerodynamic environmental constraints, (3) handling and 
transportation constraints, and (4) operational mode constraints. The 
impact of the major constraining factors were incorporated Into this study. 
Additional analyses are required to fully assess all the influencing factors. 

The guidelines and assumptions were grouped into three categories: Over- 

all Space Tug, Propulsion Module, and Astrionic Module. Each of these 
categories are discussed in more detail in the following subparagraphs. 

3.1 OVERALL SPACE TUG 

a. The study activity was directed to identify the maximum payload 
capability aeroLraked Space Tug concept, i.e., the concept with 
the least aerobrtking kit weight penalty. 

b. Minimum duration aerobraking re-entry missions {1 to 2 passes) were 
given limited study due to the thermal limits of the state-of-the-art 
radiative materials and reduced aerobraked Tuo payload capability. 

c. The unmanned geosynchronous, round trip mission was used as the 
baseline aerobraking mission. 

d. The geosynchronous missions considered nominal geosynchronous on- 
orbit time (1/2 to two days) and aerobraking return times from one 
pass to 85 passes (1/4 to 15 days). 

e. The baseline (non-aerobraked Tug configuration) was scaled from the 
configuration developed in the Boeing Pre-Phase A Space Tuq study and 
is compatible with a 65,000 pound Shuttle payload capability. 

f. The baseline Tug configuration used in this study was based on 
state-of-the-art technology. The latest proposed Tug modifications 
incorporating advanced engines and other Tug advanced technology were 
not assessed. 

g. The aerobraking kit modifications were designed to be removable and 
to be applied in kit form. The kit elements were to impose minimal 
penalty to the conventional Tug configuration. 

h. The kit elements were sized to fit within the Shuttle cargo bay 
constraints (15 foot diameter by 60 feet long). 


D5-17142 


3.1 (Continued) 

i. The Space Shuttle was assumed to deliver and retrieve the Space Tug 
at a 100 n.m. orbit. The impact of other altitudes was assessed, 

j. The payloads will be protected by the payload adapter. The maximum 
temperature the payload would experience was restricted to 300°F. 

k. The baseline return trajectory mode used was: (1) Deorbit from geo- 

synchronous orbit; (2) use aerodynamic drag to decrease the apogee 
altitude on each pass; (3) circularize at 270 n.m. after aerobraking; 
(4) phase with the Shuttle; (5) deorbit to 100 n.m.; (6) circularize 
at 100 n.m.; and (7) rendezvous with the Shuttle. 

l. The Tug provided all rendezvous and docking delta velocity with the 
payloads. 

m. A spherical earth was assumed for trajectory analysis. The impact 
of solar, lunar and earth harmonics were assessed. 

n. The study investigated the effects of the zero and small angle of 
attack trajectories. 

o. Atmospheric effect above 600,000 feet were assumed negligible. 

p. A 1962 U. S. standard atmosphere was assumed. The impacts of high and 
low density atmospheric perturbations were assessed. 

q. No solar heating or hot gas radiation was assumed. 

3.2 PROPULSION MODULE 

a. The propulsion module was scaled from the 39,800 pounds propellant 
loadina of the Boeing Space Tug Pre-Phase A Study (prior Reference 
1.1. 0.0-1) to a 45,000 pound propellant loading. Affected propul- 
sion module inerts were scaled upwards. 

b. The gaseous LOX/LHo reaction control system used for this study was 
not modified from the previous study. 

c. The upper temperature limit for state-of-the-art radiative materials 
was TD-nickel -chrome with 2000 °F capability. Advanced materials 
were identified for thermal environments exceeding 2000°F. 

d. Selected materials used for the thermal protection system were 
identified by Boeing and NASA and were based on Shuttle era tech- 
nology. No material trade studies were conducted. 


3-2 




D5-17142 


3.3 ASTRIONICS MODULE 

a. The astrionics system proposed by IBM in "Preliminary Definition of 
an Astrionics System for Space Tug Mission Vehicle Payload", Final 
Report, (Reference 3. 3. 0.0-1) was used as the baseline from which 
aerobraking Space Tug astrionics requirements were defined. 

b. The technology used for the Space Tug astrionics systems was Shuttle 
era technology. 

c. The Space Tug was based and maintained on the ground. 

d. The Space Tug astrionic module was designed to minimize the need 
for ground support. 

e. The astrionic module design concept was modular and provided the 
capability of automatic operation. 

f. The astrionic module concept was designed to allow a remove and 
replace maintenance and reconfiguration concept in space and/or on 
the ground, 

g. The astrionic system was designed to be self-sustaining. 

h. The astrionics systems was designed for automatic rendezvous and 
docking operations. 

i. The 1G0 day quiescent mode impact was not considered as a part of 
this aerobraking study, The astrionics system was sized to accom- 
plish the round trip synchronous aerobraking mission. 


3-3/S-4 


D5-17142 


SECTION IV 

PERFORMANCE AND TRADE STUDIES 


4.0 GENERAL 

This section presents the results of the detailed studies of the aerobraking 
aerodynamics, configurations, trajectories, control, thermal, astrionics, 
materials and weights. Shown in parentheses on Figure 4.0. 0.0-1 are the applicable 
subsections for each of the technical discipline results discussed in this 
section. 


4.1 Aerodynamic Analysis 

Aerodynamic data is a required input for the trajectory, loads and controls 
analysis for the space tug aerobraking study. However, in order to proceed 
with the aerodynamic data analysis, the flight regime characteristics and 
the configuration geometry had to be defined. 

In order to identify the flight regimes of interest, parametric trajectory 
data were generated to define the relationship between the initial perigee 
altitude and the total time to decay to an orbit With an apogee of 270 n. mi., 
as a function of the configuration W/CqA (assumed constant for the parametric 
trajectory data) . The results indicated that for decay times between « . 25 
and e 20 days, perigee altitudes ranged between « 220 K ft and 340 K ft 
and perigee velocities between « 22 Kfps and « 34 Kfps for a W/CpA range 
from 10 to 80 psf . Using these preliminary trajectory results, a 2:1 ellipsoid 
heat shield was selected from an aeroheating analysis and the "basic” tug 
configuration was defined as a 2:1 ellipsoid nose/cylinder configuration, 

The aerodynamic analysis required for the aerobraking study is somewhat unique 
due to the flight regimes involved. The preliminary trajectory data indicates 
that the flow field encountered by the space tug vehicle will range from hypersonic, 
slightly rarefied (slip flow) to highly rarefied (free molecule flow). Thus, 
the primary objective of the aerodynamic analysis is to predict the aerodynamic 
characteristics (Cd 0 , Cm , CP/D and Cp) of the space tug vehicle configuration (s) 
in the rarefied flow regimes encountered. 

Both the Apollo type vehicle and Ballistic missile Reentry vehicles encounter 
these flight regimes during their reentry mission profile; however, both 
terminate their missions in the continuum flow regime. The problems of 
aerodynamic heating, airloads, and vehicle dynamics are so much more severe 
in the continuum regime than an the very brief duration, rarefied flow regimes, 
that the latter have received relatively little experimental attention to 
date, particularly for complex configurations. The lack of experimental data 
is compounded by the almost total lack of theoretical or analytical methods 
applicable to aerodynamic characteristic prediction in the slightly and 
moderately rarefied flow regimes. 


4-1 



































D5-17142 


4.1 (Continued) 

In the Free Molecule Flow (FMF) regime the problem of predicting aerodynamic 
characteristics is somewhat different in nature. A large amount of experi- 
mental work has been performed; however, it primarily deals with rather 
simple shapes such as spheres, flat plates, cylinders normal to the flow 
direction, etc. In addition, theoretical techniques and results are 
readily available for many complex as well as simple shapes. The main 
source of uncertainty in the FMF regime forces result from lack of know- 
ledge of the incident molecule/surface interaction phenomena on which 
aerodynamic forces are primarily dependent. However, quantitative aero- 
dynamic force coefficient values may be calculated once the molecule/surface 
interaction phenomena has been defined, 

With the above in mind, the following approach was defined to pre- 
dict the Space Tug aerodynamic characteristics. The best estimate of 
the vehicles aerodynamic characteristics would be calculated in 
the contiuum and FMF regimes. Then, based on available wind tunnel 
data trends and empirical interpolation or ‘'bridging" schemes, a "K" factor 
would be defined which would allow the aerodynamic coefficient at any 
altitude and velocity (h,V) to be expressed as a function of the continuum 
and FMF values. That is, 

x(h*v) = x C0MT + k (y, PM p -x CONT ) 


where 


X is the aerodynamic coefficient 
Xp^p is the free molecule flow value 

\GNT con ^ nuun * value 

K is the "bridginq" parameter = f(h»V) « f (Mo#, Re« ) 

(The Bridging K factor will be discussed in more detail below. The choice 
of h&V for the independent variable results from trajectory analysis con- 
siderations. Presentation of the aerodynamic coefficients, specifically 
drag, as a function of Kjj 6 » or Mo# and R 0W was unsuitable for the trajectory 
analysis.) 

With this general "Bridging" K factor aoproach defined, the detailed aero- 
dynamic analysis was performed. 

The overall problem for the aerodynamics analysis may be divided Into two 
portions: The “basic" or no flare Tug configuration; and the flared Tug 
configuration analyses. Since the aero data required for the basic Tug 
configuration is also required for the flared Tug configuration, the basic 
Tug analysis was considered first. 


• 4*3 


05-17142 


4.1.1 Basic Tug Aerodynamic Analysis 

The basic Tuq configuration is presented in Figure 4. 1.1. 0-1. Prior Figure 
4. 0.0. 0-1 shows that the aerodynamic data required consist of drag data 
for trajectory analysis, static stability data for the controls analysis, 
and aerodynamic loading data for the structural analysis. {Dynamic stability 
data was not required, since the controls analysis did not encompass vehicle 
dynamics). Since the trajectory data was required for both the controls 
and aero heating analysis, the drag data was considered first. 


4.1 .1 .1 Drag Data 


The basic Tug drag characteristics result from pressure drag on the nose 
and skin friction drag on the nose and cylinder sidewalls. The variation 
of these two drag components in the slip and transition regimes are treated 
independently, that is, the bridging analysis, mentioned previously, is 
utilized for the nose pressure drag, while an alternate method is employed 
for the skin friction component. 


The nose (pressure) drag component is assumed to have the form 


C D (h,V) 


CONT 


(C D 


FMF 


CD 

CONT 


) K (h, V) 


where Cd cont Cd php and K(h,V) must all be determined. The first require- 
ment washed define^ the continuum and FMF "limiting" values. 


The continuum drag characteristics for the ellipsoid nose were obtained 
from Reference 4. 1.1. 1-1. As expected, the wind tunnel data agreed well 
with the modified Newtonian theory value. In the Free Molecular Flow limit, 
the following assumptions were made: 

1) Diffuse Molecular reflection 

2) Complete thermal accommodation; ce=l, T w =T f 

3) 


T =T« uniformly over the entire Tuq surface in the altitude 
range of & 550,000 ft (& FMF limit) 


These two assumptions eliminate the need to consider the majority of the 
interesting physical phenomena associated with free molecular flow. 
Justification of these assumptions are: 

A) Conservative drag estimate 

B) Simplicity of diffuse reflection aero analysis compatible with 
scope of aerobraking feasibility study 

C) Second order nature of the effects of T r at larqe molecular 
speed ratios (7£ $ m & 12.5 for the current aerobraking study) 


4*4 


0 2:1 ELLIPSOID NOSE 
0 MAIN PROPULSION SYSTEM GIMBAL POINT 
0 ASTRIONICS MODULE 
0 PAYLOAD ADAPTER 
© PAYLOAD 

DIAMETER = 14 FT. 



X/D - CALIBERS AFT OF NOSE 


FIGURE 4.1. 1.0-1. BASIC (NO FLARE) CONFIGURATION 


D5-17142 


D5-17142 


4. 1.1.1 (Continued) 

The actual FMF drag characteristics for the ellipsoid nose were based on 
an interpolation of wind tunnel data and theory for spheres and flat plate 
as a function of the molecular speed ratio; S<». 

In order to estimate the ellipsoid nose drag characteristics in the slip 
and transition flow regimes, a "bridging" analysis of sphere drag charac- 
teristics was utilized. Sphere (drag) data was chosen to define the 
bridging parameters due to the geometric simularity to the ellipsoid nose, 
and due to the large amount of experimental data available in the rarefield 
flow regimes. In addition to the experimental results, most investigators 
have attempted to "fit" the data with some empirical K factor analogous to 
the bridging parameter, K, employed herein. (Reference 4. 1.1. 1-2 presents 
some suggested "bridging" schemes, which were among those considered for 
this analysis). Most empirical approaches can be manipulated to the form 
equivalent to 

C D = C DC0NT * K ( C °FMF “ Cd C0Nt) 

where K is a function of a variety of parameters such as, (Psw.RfiOSE) ,(M <» , 
R ew > ), (Re<»), (Re behind a normal shock, R e $), (Pw.Ps.Ae* RjiOSE)* t wa11 
to free stream enthalpy ratio), ($u, Soa, Kn^, Te®» Me»), plus a number 
of schemes which have experimentally defined "free constants". 

The following approach was employed to select K. for the ellipsoid nose drag 
bridging. The various empirical methods were converted to a common set of 
independent variables, altitude and velocity (h,V) and a "best estimate" 

K (h,V) determined. The best estimate K is based on Reference 4. 1.1. 1-1, 

ReS approach in the near continuum regime and a method suggested by Willis 
(Reference 4. 1.1. 1-1) in the near FMF regime. A smooth fairing was 
utilized between the two schemes. 

As discussed in Reference 4. 1.1. 1-3 the small percentage of the blunt body 
experimental data which indicate an "overshoot" of the FMF boundary can be 
explained via either incomplete thermal accommodation in the near molecular 
flow regime, or a non diffuse surface reflection phenomena. Thus, with the 
assumption of complete thermal accommodation and diffuse reflection, the 
bridging parameter was limited to values from 0 to 1. Therefore, the slip 
and transition flow nose drag coefficient is bounded by the continuum and 
FMF values. 

In addition, the bast estimate of the K(h,V) was constrained to comply 
with the following contlnuum/slip and transitlon/FMF boundaries: 

SLIP OCCURS 0 Me#/ /“R“ sy .01 (References 4, 1.1. 1-4, -5 and -6) 

FMF OCCURS 0 Am/D RS 10 (References 4.1, 1.1-4, -5, -6 and -7) 

Mote that for T w c T r =T w ,AaA w ) 




-4-6 


D5-17142 


4. 1.1.1 (Continued) 

(The reference length* D, is the nose radius of curvature at the stagnation 
point, which for the 2:1 ellipsoid nose geometry equals the body diameter 
of 14 ft. Atmospheric properties which were not available fromthe 1962 
U. S. Standard Atmosphere were obtained from the 1963 Patrick Air Force 
Base extension to the 1962 USSA. ) 

Applying the best estimate K(h,V) resulting from the sphere analysis to the 
2:1 ellipsoid nose (pressure) drag bridging, allows definition of the nose 
drag as a function of altitude and velocity. A best estimate K(h»V) 
applied to the 2:1 ellipsoid nose is presented in Figure 4. 1.1. 1-1 alonq 
with a comparison of some alternate bridging approaches. 

The skin friction variation between the continuum and free molecular flow regimes 
was not "bridged" by the nose drag K factor. Instead the following approach was 
used. The continuum limit was determined from Reference 4. 1.1. 1-8 and modified to 
account for compressibility effects (Reference 4. 1.1. 1-9), assuming the cylinder 
sidewall was equivalent to a flat plate. THE FMF limit was calculated (References 
4. 1.1. 1-4 and -8) based on the previously stated assumptions. The slip flow effects 
were then estimated via the techniques of References 4. 1,1. 1-5, -10 and -11 and a 
"best estimate" relationship defined which was consistent with the continuum and 
FMF limits. 

Combination of the skin friction drag characteristics with the bridged nose drag 
characteristics allows definition of the basic Tug drag coefficient as a function 
of altitude and velocity as presented in Fiqure 4. 1.1. 1-2. Figure 4.1. 1.1-3 
presents the "original" basic Tug drag coefficient based on a total lenqth of 
60 ft. The final Tug total length of 50.4' draq characteristics were used 
in a trajectory analysis and only a slight variation in trajectory parameters was 
found, due to the small change in the near continuum characteristics combined 
with the low perioee altitudes required for the hi qh W/CqA basic Tua confiauration. 

4. 1.1. 2 Static Stability Data 

The second set of aerodynamic data desired was the static stability characteris- 
tics, and CP/D as a function of altitude and velocity. 

The C Ne( and CP/D contribution of the ellipsoid nose in the continuum reqime was 
based H on the wind tunnel characteristics of Reference 4. 1.1. 1-1, 

In the FMF regime, the nose component &.£* and CP/D were again estimated by an 
interpolation of flat plate and sphere theoretical and wind tunnel as a function 
of the molecular speed ratio, Sm . 

The embedded Newtonian theory (References 4. 1.1. 2-1 and -2) is utilized for the 
determination of the static stability contribution (C« a and CP/D) of the cylinder 
in the continuum flow limit. The embedded Newtonian theory requires a knowledqe 
of the bow shock structure due to the nose and the flow field properties "embedded" 
between the body and bow shock. The following method was employed to estimate the 
bow shack profile and the embedded flow field properties: The ellipsoid nose bow 


4-7 



D5-17142 








20 K FPS 


Ar = 154 FT 


i um 

vmm 


NOTES: 

1) ANGLE OF ATTACK = 0° 

2) 2:1 ELLIPSOID HEAT SHIELD NOSE 

3) D1AM. = 14* 

4) LTH = 60’ 

5) C D MAY BE ASSUMED CONST. 

FOR ALT. BETWEEN 150 KM AND 200 KM 

6) DIFFUSE REFLECTION 

7) T u /T m = 1 (ASSUMED) § 150 KM 


5Q 60 70 80 90 1QQ 110 120 130 140 J50 

ALT - KM 

FIGURE 4.1. 1.1-3: PRELIMINARY (NO FLARE) BASIC TUG DRAG CHARACTERISTICS 


j I 


D5-17142 






D5-17142 


4. 1.1. 2 (Continued) 

shock profile was estimated based on the data and method of References 4. 1.1. 2-1 
and -3. The flow field characteristics (P./Poo) and (qL/qoo) in the embedded flow 
field at a given X/D (cal. aft of the nose) and Y/D cal. normal to the Tug center- 
line axis) were assumed to be proportional to the characteristics of References 
4.1. 1.2-1 and -2 hemisphere-cylinder configuration when normalized with respect 

t0 v/r shock. 

Thus, the simple Newtonian cylinder normal force (with centrifugal correction) of 
Reference 4. 1.1, 2-4 was modified to account for the embedded flow field properties 
in order to estimate the distribution of C Na along the cylinder sidewalls, 
dC N0( /d( X/D) . Numerical integration of the H distribution and its moment resulted 
in both the cylinder C^ a and CP/D in the continuum limit. 

In the FMF limit, the cylinder C NO and CP/D were determined via the method and 
tabulated data of References 4. 1.1. 1-7 and 4. 1.1. 2-5. Again, fully diffuse re- 
flection at complete thermal accommodation is assumed, and T u = To# is also 
assumed uniformly over the entire vehicle. w 

Combining the nose FMF contribution allows the total and CP/D to be 
calculated in the FMF limit as a function of S<». With the continuum and 
free molecular flow limits defined, some method must again be determined 
to "bridge" the characteristics in the slip and transition regimes. As 
suggested in References 4. 1.1. 2-6 and -7 the same bridging relationship 
employed for the drag characteristics have been utilized to bridge the 
normal force coefficient and center of pressure between the continuum and 
FMF limits. As discussed previously, the non-existence of the phenomena 
of "overshoot" is also assumed to apply to C^^ and CP/D). The final basic 
Tug configuration CNa and CP/D are presented in Figure 4. 1.1. 2-1 as a func- 
tion of altitude and velocity. These data are required for the controls 
and RCS propellant requirements analysis. 

4. 1.1. 3 Nose Pressure Distribution 

The aerodynamic data requirements for the basic Tug will be completed with 
the definition of the nose pressure distribution and the corresponding 
maximum loading case. As stated in the discussion of the continuum nose 
drag, modified Newtonian theory agreed well with experiment. Thus the fol- 
lowing approach was utilized for the nose pressure distribution. The local 
coefficient was assumed to have the form of C 0 = C D sin z oi 

H L0CAL P MAX LOCAL. 

/ y± 3 \ L 2 

C D was based on a Y of 1.2 and the relationship, C D =i y +l /V wn: 

K MAX h MAX x 7 x 7+i 

The nose geometry plus a 6° angle of attack was used to determine a LOCAL. 

The effects of bow shock induced cylinder pressures were analyzed based on 
the data of Reference 4. 1.1. 3-1 (checked with References 4.1. 1.3-2, -3 and 
-4), slightly modified for differences in the nose/cylinder geometry. The 
resulting pressure coefficient distribution on the nose and nose/cylinder 
region is presented in Figure 4. 1.1. 3-1. In order to determine the worst 




1/RAD cp/d ~cal. fwd of base 


f 


D5-17142 



ALTITUDE KM 
CENTER OF PRESSURE 



ALTITUDE - KM 

COEFFICIENT OF NORMAL FORCE SLOPE 


FIGURE 4. 1.1. 2-1. BASIC (NO FLARE) TUG STATIC STABILITY CHARACTERISTICS 



4-12 




D5-17142 








D5-17142 


4. 1.1. 3 (Continued) 

case airloads condition the following approach was employed. The trajectory 
characteristics (h,M) at MAX q on each pass were determined from the nominal 
trajectory data. From the bridging analysis, the maximum divergence of the 
aerodynamic coefficients from their continuum value was determined at the 
(h,M) of q MAX and the product K BRI0QE . q max D EFINED* The m * imm P roduct was 
used to define a K* which would allow direct conversion of CP L0CAL to the 

local pressure P, . (A small correction factor is added to account for the 
small ambient pressure). The conversion factor, K A ‘, is also presented in 
Figure 4. 1.1, 3-1 as a function of the number of passes required to achieve 
an apogee of 270 MM. 


The coefficients of local pressure (C p ) and the K«* factors shown in Figure 
4. 1,1. 3-1 were converted to a local pressure profile (in pounds per sauare 
inch) over the heat shield and cylindrical section of the basic configuration. 
Figure 4, 1.1. 3-2 illustrates the rapid decrease in local pressure from the nose 
to the base of the heat shield. This near-order of magnitude decrease is 
followed by a gradual decrease along the Tug's sidewall. For a nominal mission 
time of 30 passes (5 day) for the basic (no flare) Tug configuration* the local 
pressure is 0.12 psi at the heat shield nose; 0.013 at the heat shield/sidewall 
interface; 0.002 at the propul sion/as tri on ics module interface and 0.0017 at the 
astrionic/payload interface. These low pressure loads will not have a significant 
design impact on the configuration aft of the heat shield. 

4.1.2 Flared Tug Configuration Analysis 

The objective of the flared Tug aerodynamic analysis is somewhat broader than 
the basic Tug configuration analysis. Actually, the aerodynamics analysis had 
two specific requirements; First, the configurations had to be defined, and 
then, their aerodynamic characteristics had to be determined. The first objective 
was to define the configurations. Figure 4. 1.2, 0-1 presents simplified versions 
of the technical approach to sizing the flares. 

4, 1.2.1 Flow Field and Flare Sizing Analyses 

A preliminary configuration analysis identified three preliminary flare 
configurations for consideration. The physical constraints placed on the 
configurations were; 

1, Avionics visibility capability, and 

2. Attachment at some main structure location 

In addition to these constraints, the following criteria were defined 
for the flare selection; The "Targe" flare would be defined such that its 
W/ r nA would be approximately an order of magnitude less than the basic 
Tug, and, also, provide a large static stability margin (both calculated 
in the continuum flow regime for expediency). 


4-14 



05-17142 


4. 1.2.1 (Continued) 

The ''small" flare would be defined such that it would be approximately 
neutrally stable in the continuum flow regime. 

The criteria for the third or "intermediate" flare selection will be 
discussed later. 

Having defined the flare selection criteria and constraints, the first 
requirement for the flare sizing study was to determine the flow charac- 
teristics in the continuum flow regime. As discussed above, the basic 
Tug nose bow shock profile was estimated via the methods of References 
4. 1.1. 2-1 and -3, and is presented in Figure 4. 1.2. 1-1. The flow field 
characteristics, Pl/P w q|./q«»» were estimated on the basis of the data 
in References 4,1, T. 2-2 and -3, and are presented in Figure 4. 1.2. 1-2. 

For the current study, the following two groundrules were established 
which seemed justified on the basis of the reduced scope of the current 
feasibility study. First, flow separation effects would not be quanti- 
tatively analyzed. Althouah some degree of flow separation will undoubtably 
exist above some limitinq (small) flare anqle, these effects are expected 
to decrease progressively in the slip flow regime. In addition, various 
geometric modifications might be employed to reduce flow separation and 
the associated problems (for example, a configuration similar to Reference 
4. 1,2. 1-1). 

The second groundrule was that boundary layer would be assumed "small 
enough", such that the flare characteristics in the presence of the boundary 
layer would not differ appreciably from the characteristics defined, 
assuming a boundary layer thickness of zero. 

Both of these groundrules should be given extensive consideration in future 
aerobraking studies. While both appear incompatible with a rigorous 
analytical analysis, they should be appraised on the basis of the following 
considerations: 

1. A detailed flow field analysis was beyond the scope of the aerodynamic 
study. 

2. The expenditure of a large amount of time analyzing flow separation 
and boundary layer characteristics would not be warranted, if, by 
disregarding these effects* the flare concept still appears unfeasible 
due to either weight or thermal protection system requirements. 

3. The optimum perigee altitude for the flared configurations may be 
sufficiently hiqh that the continuum flow analysis may be inapplicable* 
i.e., slip flow effects may dominate the flow field and thus render a 
continuum regime flow separation analysis overly pessimistic. 

4. The continuum flow characteristics are utilized primarily to establish 
the "lower bound" on the flared configuration aerodynamic charac- 
teristics. 


-- 4-15 



D5-17142 


4. 1.2.1 (Continued) 

Thus, with the flow field characteristics and above groundrules defined, 
consideration was directed to the parametric flare effectiveness analysis. 

The concept of embedded Newtonian theory (References 4. 1.1. 2-1, -2 and -3), 
plus the analysis of Reference 4. 1.2. 1-2, are utilized for the parametric 
flare effectiveness analysis. The flare attach point was chosen as the 
payload adapter for the following reasons: 

1. Minimize astrionics module visibility problems (field of view = 90° 

+ (90° -d v °). 

2. Aft flare location will result in larqe flare stabilizing effect. 

3. Payload adapter should qualify as a stronq structural attachment 
point. 

4. Location of flares at same body position will allow maximum extra- 
polation of flare length and semi vertex angle results. 

Figures 4. 1.2. 1-3 and 4. 1.2. 1-4 present theilCd @c*=0 and due to 
the flare as a function of the flare semivertex angle and flare slant 
length. These data plus the following rationale allowed definition of 
the flare configurations. 

4. 1.2. 2 Flare Selection Rationale 

A 60° flare angle was selected for the large flare configuration. A steep 
flare angle is required in order to obtain the maximum 4Cd 0 per-unit flare 
length and per unit flare surface area. One may be able to "live with" or 
"fix" the flow separation problems associated with a 60° flare, but pos- 
sibly not a 90° flare. Thus, the limiting case of a forward facing flat 
plate (90° fl v ) was groundruled out. 

A flare slant length of one caliber (14') was selected, in order to 
insure that the bow shock would not Impinge on the flare at 6° angle of 
attack. As can be seen from Fiqure 4. 1.2. 1-4, the 1 caliber, 60° flare, 
produced a restoring moment of approximately seven times the basic Tuq 
nose/cylinder disturbing moment. (The effects of the flare shock/bow 
shock interaction have hot been considered). Figure 4.1. 2,2-1 presents 
the geometry of all flare configurations analyzed. 

As previously stated, the selection criteria for the small flare was neutral 
stability in the continuum regime. Frgm Figure 4, 1.2. 1-4 one can see that 
a neutrally stable configuration can be obtained (continuum flow, 

Me®~35) with the 60°, 45°, 30°, or 15° configuration, if the flare lengths 
are .625, .65, .805 or 1.315 calibers, respectively. 

Since it appeared that some configuration optimization study would be 
desirable near the studies conclusion, the decision was made to consider 
three different semivertex angle flares. Ideally then, the parametric 


4-16 ' 


LOCAL PRESSURE (POUNDS PER SQUARE INCH) 



' D5-17142 



DISTANCE APT OP HEAT SHIELD NOSE (FEET) 

FIGURE 4. 1.1, 3-2: MAXIMUM LOCAL PRESSURE PROFILE 

(BASIC - NO FLARE) 

- - / 4-17 




BASIC 

TUG 

AERO BATA 


D5-17142 





























Y/Drm CALIBERS PERPENDICULAR TO CENTERLINE AXIS 


NOTES: 

© 2:1 ELLIPSOID NOSS COMMON TO ALL CONFIGURATIONS 
0 MAIN PROPULSION GIMBAL POINT 
© ASTRIONICS MODULE 
© PAYLOAD ADAPTER 
© PAYLOAD 

TUSDSAM. = D R = 14 FT. 


CONFIG. 


SUNT LTH 

LARGE 

60° 

14.0 

INTERM. 

45° 

14.7 

SMALL 

30° 

11.25 

OPT. #1 

15° 

17.4 

opt. n 

30° 

18.4 



X/Dr «* CALIBERS AFT OF NOSE 


FIGURE 4.1 .2.2-1 


FLARED TUG CONFIGURATION (S) GEOMETRY 


D5-17142 


D5-17142 


4. 1.2. 2 (Continued) 

flare data, combined with the trajectory, aerodynamics, weights, and thermal 
analysis might allow at least a qualitative traue study to be performed 
which would include 6 V as a parameter. Thus, the 3Q° flare was chosen 
for the "small" flare configuration. Note that the flare slant length 
required for neutral stability is 2* .805 calibers. 

Optional flare #1 is a 15° semivertex angle which also results in neutral 
stability. Its slant length is 1.315 calibers. The choice of the 30° d v 
for the "small" flare rather than the 15° 6 v is based on the followina: 

1. 4Cp 0 30° flare is approximately 2 times the /JCd 0 for the 15° flare 
in the continuum limit. 

2. The total flare surface area is ££63% lower for the 30° flare. Thus, 
although the TPS weight per unit area will probably increase as dy 
increases (at a given flow field environment), the total TPS weight 
may be less due to the lower surface area, 

3. The TPS required for the 15° flare may not differ appreciably from 
the basic Tug cylinder requirements, and, hence, given the 15° flare 
trajectory characteristics, the TPS required might be estimable 
without a rigorous aero heating analysis. 

A complete set of aerodynamic data is included for the “small" 30° flare. 

The drag characteristics for the 15° flare have also been generated for 
the trajectory analysis in case this data is desired in the final con- 
figuration optimization studies. Both neutrally stable configurations 
are presented in Figure 4. 1.2. 2-1. 

The remaining flare configuration, the "intermediate" flare, was selected 
primarily on the basis of the W/C^A required for trajectory analysis. 

Figure 4. 1.2. 2-2 indicates the nonlinear relationship between time to 
decay, N and W/CpA. Figure 4,1. 2. 2-3 illustrates the Cp and W/CpA ranqe of 
the basic Tug, large and small flare configurations, and the W/CpA and Cp 
of the "intermediate" flare required to investigate the nonlinear N vs. 
W/CpA relationship. As discussed earlier, it was also desired to 
investigate flare semivertex anqle effects with the three flare configura- 
tions. 

Thus, the "intermediate" flare was defined as the following: 45° senii- 
vertex angle, flare slant length of 1.05 calibers. The continuum value 
of W/CpA for the "intermediate" flare configuration issi30 psf. 

Optional flare #2 was defined as a 30° setnivartex angle flare with a 1.24 
caliber slant height. This configuration is close to the W/CpA of the 45° 
flare, and might allow flare length changes to be investigated (at least 
qualitatively) in the final configuration optimization studies. 


4-23 


BALLISTIC COEFFICIENT (W/C D A) - psf 


* 








D5-17142 


4. 1.2. 2 (Continued) 

The "family" of flared Tug configurations is presented in Fiqure 4. 1.2. 2-1. 
The configurations have been selected in order to satisfy the following 
requirements and objectives: 

Requirements: 

ll Avionics visibility 

2) Reasonable structural attachment point 

Objectives: 

1) Investigate an approximate order of magnitude variation in 
8 < W/CpA < 80 psf. 

2) Select configurations such that any non linearities in the time 
to decay vs. W/CpA will be analyzed. 

3) Investigate a range of static stability from approximately 
neutral stability to large static stability. 

4) Investigate flare semi-vertex angle, effects. 

Having defined the flare configurations, the task remained to estimate their 
aerodynamic characteristics, {Cd 0 » C^, CP/D, and Cp) over the entire al- 
titude and velocity flight regime of interest. As with the basic Tug 
configuration, the draq data was the first objective, 

4. 1.2. 3 Flared Tug Drag Characteristics 

Prediction of the flared Tug configuration drag characteristics requires 
combination of the basic Tug nose drag and skin friction draq characteristics 
with the flare pressure drag and skin friction characteristics. The skin 
friction drag component for the flared configurations was assumed to have 
the same altitude and velocity (Me®, Reoo) variation as the basic Tug, 
based on the flared Tug wetted area. The ellipsoid nose pressure drag 
component was assumed unchanged due to the presence of the flare, (Note 
that this is equivalent to assuming that the characteristic dimension of 
the vehicle is the nose characteristic dimension (vehicle diameter), 
for the nose aerodynamic characteristics and flow regime analysis, even 
though the vehicle actually has a much larger maximum dimension, such as 
the flare diameter. While this assumption surely is invalid for very 
large flares or flares close to the nose, it appears reasonable for the 
configurations considered). Thus, having defined the nose and skin friction 
drag contributions throughout the flight regimes, the remaining task is to 
define the flare drag contribution. 

The flare continuum drag contribution estimate is based on the embedded 
Newtonian theory (Reference 4. 1.1. 2-1, -2 and -3), plus the analytical 
approach of Reference 4,1. 2, 1-2, The FMF contribution is obtained via 
the method and data of References 4. 1,1. 1-7 and 4. 1.1. 2-5. The basic 
Tug bridging factor, K (h*V), is correlated to a K (M<x» t Rew)» and re- 
defined for the flare characteristics usinq a reference length equal to 


D5-17142 


4. 1.2. 3 (Continued) 

the flare base diameter. (With the assumption of nose draq independence 
with flare configuration, this is equivalent to saying that the nose 
will reach FMF at a lower altitude (smaller A») than the flare. While the K 
(h,V)' so defined for the flare characteristics bridging analysis is 
somewhat less justifiable than that for the nose bridging analysis, it 
is the only approach which appeared within the scope of the current 
feasibility study. Utilizing this K (h,V) allowed definition of the flare 
component drag characteristics over the (h,V) ranqe of interest. Combininq 
these data with the ellipsoid nose data and the total vehicle wetted area 
skin friction contribution, allowed definition of the flare configuration 
drag characteristics. 

Presented in Figures 4. 1.2. 3-1 and -2 are the drag characteristics for the 
flared Tug (large, intermediate, small, option #1 and optional #2) con- 
figurations. It is noteworthy to compare the variation of Cq of the five 
configurations in the continuum and FMF limits. In the continuum regime, 
the embedded flow field characteristics (primarily the reduced local dynamic 
pressure), dominate the flare drag effectiveness; however, in the FMF limit, 
the drag characteristics are more closely proportional to the flares maxi- 
\ mum cross-sectional area. 

Having completed the flare configuration draq analysis, attention was 
directed to the static stability characteristics. 

4. 1.2. 4 Static Stability Analysis 

The approach to the flared configuration static stability analysis parallels 
the approach to the draq analysis exactly. The contribution of the basic 
Tug nose/cylinder was modified for the shorter effective cylinder lenqth and 
assumed independent of the flare configuration. The same (draq analysis) 
referenced data and bridging factor was utilized to determine tho static 
stability contribution of the flare(s). Combination on the nose/cylinder 
Ch« and CP/D with the flare contribution results in the static stability 
characteristics of the flared Tug confi durations. The large, medium and 
small flare configuration static stability data. Cm « and CP/D, are presented 
in Figures 4. 1.2. 4-1 and -2, respectively. (Recall that only draq data 
was to be defined for the optional flare configurations). 

One rather interesting aspect of the static stability data is the variation 
of CP/D from the continuum limit to the FMF limit. In comparing the basic 
Tug CP/D variation, (Figure 4. 1.1. 2-1), with the flared configurations, the 
following items should be noted: 

1. Each flare configuration is opposing the same basic Tua disturbing 
moment. 

2. For the basic Tug, the CP/D (Figure 4. 1.1. 2-1) variation from 
continuum to FMF is primarily a “planform 11 effect, i.e.» the CP 

moves from the nose region in the continuum reaime towards the vehicle 
"area centroid" in the FMF regime. 


4-26 



ALT- KM (N.M.) 


FIGURE 4,1. 2.3-1: DRPG COEFFICIEN 



05-17142 










FIGURE 4,1 ,2.3-2. DRAG COEFFICIENTS FOR OPTIONAL FLARED CONFIGURATIONS 


D5-17142 



ALT~ KM (N.M.) 


ECIENT OF NORMAL FORCE SLOPE FOR FLARED CONFIGURATIONS 







CP/D - CAL. APT OF NOSH 


BASIS 



REFERENCE AREA (A r ) =154 FT 2 
REFERENCE DIAMETER = 14 FT 2 


t 

CO 

o 


Vco KFPS 




ALT - KM {M.MJ 

FIGURE 4 J. 2.4-2: CENTER OF 


CP/D - CAL. AFT OF NOSE 




FOR FLARED CONFIGURATIONS 


D5-17142 




D5-17142 


4. 1.2.4 (Continued) 

3. For a 30° fl v flare, the CP of the flare is approximately at the 
base of the flare in both the continuum and FMF regime. Thus, the 
planform effect of the basic Tug nose cylinder dominates the CP 
variation with altitude, and the CP again moves from the nose aft. 


4. For the 45 and 60° 3 V flares, the continuum flare CP is ^ 1,0 and 
4.0 calibers aft of the flare base, respectively, in the continuum 
limit. However, in the FMF limit, the CP of both flares moves 
forward to approximately the flare base. Thus, for the large and 
intermediate flare configurations (60® and 45°), the large negative 
flare Cft^'s and associated forward CP shift from continuum to FMF 
dominate the total vehicle CP/D variation with altitude. Hence, 

CP moves forward, towards the nose, as the flow regime changes from 
continuum to FMF. 


In determining the C^ and Cjq of the flared configurations in the continuum 
limit, the effects of flow shielding on the “upper* 1 side of the flare have 
not been included. The reason for this is that neglecting the shielding 
effects will result in a conservative estimate of the continuum stability 
at a considerable savings In time, due to the complexity involved in the 
shielding analysis, 

With the completion of the flared configuration static stability, the 
only remaining aerodynamics analysis task was the definition of the nose 
and flare pressure coefficient distribution. 

4. 1.2. 5 Pressure Coefficient Distribution Data 


The nose and cylinder sidewall pressure coefficient distributions at zero 
angle of attack were determined in exactly the same manner as the basic 
Tug. The distributions for the flares are presented in Figure 4. 1.2, 5-1. 
The distributions are seen to differ only in the K* factor (determined in 
the same manner as discussed in Section 4. 1,1. 3), supplied to convert 
Cp to local pressure In psf. 

The flare pressure coefficient distributions, are based on a similar, 
maximum (Kbridge* 4 max ) product plus the local embedded flow field pro- 
perties determined from the flow field analysis (Pi/P*»and qi/qqs). Both 
the uniform distribution at ce« 0® and the non-uniform distribution at 
ct= 6® were considered since it appeared conceivable that the worst 
possible load case might result from an assymetrical flare load. Thus, 
flow shielding effects are considered in defining the flare pressure 
coefficient distribution. 

Figure 4. 1.2, 5-2 presents the flare pressure coefficient distributions for 
the large, intermediate, and small flare configurations. Definition of 
the pressure coefficient distributions completes the aerodynamic data 
requirements for the Space Tug Aerobraking Study. 


4-31 


. LOCAL PRESSURE COEFFICIENT 1C 


<c 

o 

o 


• ANGLE OF ATTACK = 0° 

• N = NUMBER OF PASSES 

• K n * IS FUNCTION OF N 

o P L = PRESSURE LOCAL <P*f) 
> LOCAL X K n* 


N 

*7 1 

30®. FLARE 45 

8 FLARE 

60® FLARE 

5 

2 7.6 

12.6 

6 

10 

14.9 

6.7 

3.4 

30 

5.7 

2,6 


69 

3.0 

1.5 

mmm 



10 .2 .4 .6 IQ 

X/D - CAL AFT OF NOSE 


FLARE/CYL. JCT. 

10.0 


FIGURE 4.1. 2.5-1: LOCAL PRESSURE COEFFICIENTS OVER HEAT SHIELD 

AND CYLINDER (FLARED CONFIGURATIONS) 


D5-17142 







































D5-17142 


4. 1.2. 5 (Continued) 

The local pressure (C p ) and the K n * factors given in the prior Figure 
4. 1.1. 3-1 and in Figure 4. 1.2. 5-1 were converted as shown in Figure 
4. 1.2. 5-3 to illustrate the pressure loads (in pounds per square inch) 
for the heat shields of the basic (no flare) and 30°, 45° and 60° flared 
Tug configurations. The maximum pressure occurs at the nose for each 
configuration and decreases from 0.47 psi for the basic (no flared) 

Tug to 0.08 psi for the 60° flared Tug. As the number of passes increases 
from 5 to 60, the pressure loads on the shield decrease by approximately 
an order of magnitude. 


The local pressure coefficient (C p ) and the K n * values for the flares 
are shown in Figure 4.1. 2. 5-2. This data was converted to pressure loads 
(pounds per square inch) as shown in Figure 4, 1.2. 5-4. The data pre- 
sented is the "worst case" pressure loads which would occur at the outer 
edge of the flare lower surface (assuming a zero or six degree angle of 
attack). The 45° and 60° flare configurations with a six degree angle 
of attack have approximately the same pressure loads. For a zero angle 
of attack the pressure loads on the 60° flare are approximately 1.6 
times as great as that on the 45° flare. The shorter length and lesser 
angle 30° flare has a significant drop in pressure loads compared to 
either the 45° or 60® flare. The pressure loads on the outer edge of 
the flares represent the "worse" loads conditions on the flare; however, 
the loads are extremely low but do impact the desiqn of the flare for the 
low pass missions (5-10 passes). 


The ‘above figures have identified the pressure levels at specific points 
(heat shield nose, flare outer edge) as a function of the number of 
passes. Figure 4. 1,2. 5-5 illustrates the local pressure loads along the 
entire length of the vehicle for the 30®, 45° and 60® flared configura- 
tions. The 30 pass mission and zero angle of attack was selected as 
representative of the mission variables. The local pressure has a rapid 
decrease over the edges of the heat shield ellipsoid. The 30° flare configura- 
tion has a higher pressure (0.75 psi) on the nose than on the outer flare 
edge (0.0076 psi). The 45° flare has a smaller pressure differential 
between the nose and outer flare edge (0,34 psi to 0.21 psi). The 
pattern is reversed for the 60° flare with the outer flare edge's 
pressure (0.38 psi) greater than that on the nose (0.019 psi). The in- 
crease in pressure along the slant height of the flares is more pronounced 
with the steeper angles. These flare pressure profiles impact the flare 
thickness and supporting structure which contributes to the heavier 
weights of the 60° and 45® flares versus the lighter weight of the 30® 
flare. 


4-34' 


MAXIMUM LOCAL PRESSURE (PSI) 









NUMBER OF PASSES 

FIGURE 4.1. 2,6-3; MAXIMUM LOCAL- PRESSURE ON HEAT SHIELD HOSE 

■V* / 4*35 



MAXIMUM LOCAL PRESSURE (PS1) 


l 


D5-17142 



NUMBER OF PASSES 

FIGURE 4.1. 2,5-4: MAXIMUM LOCAL PRESSURE AT OUTER EDGE OF FLARE 

' ' - 4 - 36 * 


LOCAL PRESSURE (POUNDS PER SQUARE INCH) 


05-17142 













D5-17142 


4.2 AEROBRAKING CONFIGURATION CONCEPTS 

This section describes the aerobraking Space Tug configuration concepts. 
These aerobraking concepts were bounded by the study groundrules to 
aerobraking kits which can be applied to the conventional Space Tug. The 
study duration (four months) coupled with the limited funding did not 
permit analyses of all of the feasible aerobraking kits nor did it permit 
an examination of all of the more practical modifications to the Space Tug 
to enhance its aerobraking capabilities. The study did, however, provide 
insight to aerobraking capabilities of the Space Tug and did identify the 
environmental criteria effecting material selection and design concepts. 
Aerobraking design concepts were investigated in sufficient depth to 
define some reasonable configuration concepts, their functional capabi- 
lities and their weights. 

The key study groundrules which impacted the aerobraking kit design 
concepts are listed below. These were: 

a. The baseline (non-aerobraked Space Tug configuration) would be that 
developed in the Boeing Pre-Phase A Space Tug Study (prior 
Reference 1.1. 0.0-1) and is compatible to the 65,000 pound 
Shuttle payload capability. 

b. The aerobraking kit modifications were designed to be removable 
and were designed to be applied in kit form. 

c. The kit elements imposed minimal weiqht or configuration penalty to 
the conventional Tug configuration, 

d. The kit elements were designed to fit within the Space Shuttle cargo 
bay constraints (15 foot diameter by 60 feet long). 

e. The payload thermal limit was 300°F. The aerobraking payload adapter 
kit was designed to this limit. 

f. The aerobraking kit elements were designed to be reusable, where 
practical, 

g. The weights of the aerobraking elements were minimized in order to 
maximize payload capabilities. The design concepts were, therefore, 
directed to use lightweight as the governing design criteria. 

h. The thermal protection systems applied to the Tug for the aerobraking 
usage would be shuttle era technology. 

i. The state-of-the-art for radiative thermal protection materials was 
assumed to be 2000°F capability with a TD-nickel -chrome alloy. 

j. The astrionics systems applied for the aerobraking would be shuttle 
era technology. 


' ' ’4-38 



D5-17142 


4.2 (Continued) 

The subsequent paragraphs of this section identify the conventional tra- 
jectory Space Tug configuration used as a starting point for the study 
(Section 4.2.1), the aerobraking kit elements (Section 4.2.2) and then 
identifies the aerobraking kit design options and recommendations (Sec- 
tion 4.2.3). 


4.2.1 Conventional Trajectory Space Tug Configuration 

The study groundrules stipulated that the aerobraking kits would be 
applied to the conventional trajectory Space Tug configuration. Figure 
4. 2. 1.0-1 illustrates the conventional Space Tug used as the baseline. 

The Tug consists of a propulsion module, an astrionics module, payload 
adapter and the payload. The Boeing Space Tug design concept used a 
14 foot diameter rather than the 15 foot Shuttle cargo bay diameter 
as a design limit. The one foot diameter difference was allowed to 
account for micrometeoroid shielding, reaction control system nozzle 
clearances and to allow for other proturbances the vehicle may require. 

The propulsion module stage length with the two position nozzle retracted 
(as during delivery to orbit by the Shuttle) is 33.4 feet. After removal 
from the Shuttle, the propulsion module with the nozzle extended is 38.25 
feet. The stage has a LOX/LHo propulsion system with an uprated RL-10A- 
3-8 engine. The reaction control system consists of four sets of four 
thrusters located 90° apart at a vehicle station comparable to the bottom 
of the LH 2 tank. These thrusters are used for pitch and yaw maneuvers. 

The eight roll thrusters are located at a station comparable to the engine 
gimbal point and are spaced 90° apart. The astrionics module is 14 feet 
diameter and four feet high. It is mounted atop the propulsion module 
and below the payload adapter. The combined length of the propulsion 
and astrionics module is 37.5 feet. The remaining length (60 foot long 
Shuttle cargo bay minus 37.5 feet) is 22.5 feet which is available for 
housing the payload adapter and payload. 

The propellant loading in the propulsion module is 45,000 pounds of LOX/LHg. 
The propulsion module inerts weigh 5,868 pounds, qiving a stage weight of 
50,868 pounds and a stage mass fraction of 0,885. The astrionics module 
weighs 2526 pounds (from IBM's previous Space Tug Study - prior Reference 
3. 3. 0.0-1). This module used state-of-the-art technology and v/as updated 
in this study to reflect Shuttle era technology. The advanced technology 
reduced the astrionics module weight to 1960 pounds. The mass fraction 
of the overall Space Tug with the old astrionics module and the new 
astrionics module sre 0.840 and 0.849, respectively. 

To provide the docking and holddown capability of the payload to the 
Space Tug, a payload adapter will be required. This adapter was de- 
signed to be a separate, removable kit. The weight of the adapter Is 
200 pounds. The payload adapter, when used with a flared configuration 
aerobraking Tug, is used as a multipurpose structure: (1) To mount the 

aerobraking flare to the Tug; (2) to provide housing for the fiare 
actuation system; and (3) to provide a thermal protection housing for 
the payload. Because this adapter is a special fixture required for 


4-39 


4-40 


, 

6Q’0" 





ASTRIONICS 

MODULE 



LHj 

TANK 


33'5" 


WEIGHT ESTIMATE (LBS) 


PRE-PHASE A SPACE TUG 
SPACE TUG WITH 

UPRATED 

ASTRIONICS 


e PROPULSION MODULE 

® PROPELLANT 45,000 

• INERTS 5, 868 

• ENGINE 639 

• PROP/MECH 801 


•THERMO/MICRO 573 

• STRUCTURE 2912 

• CONT. & RESID. 943 


50,868 


50,868 



® ASTRIONICS MODULE 
® ELECTRICAL 
• AVIONICS 
« STRUCTURE 
“THERMAL 

* PAYLOAD ADAPTER 

TOTAL SPACE TUG WEIGHT 
MASS FRACTION 


RL-tOA-3-8 ENGINE 
(NOZZLE RETRACTED) 


841 

625 

400 

660 


2,526 1,960 

. 515 

842 
300 
303 

200 200 


53,594 53,028 

0,840 0.849 


FIGURE 4.2.1. 0-1: CONVENTIONAL SPACE TUG CONFIGURATION 


D5-17142 


D5-17142 


4.2.1 (Continued) 

aerobraking, it was considered an aerobraking kit element. 

The conventional Space Tug configuration was based on the previously 
conducted Boeing Pre-Phase A Space Tug Study and on the IBM Astrionics 
Module Study. Both of these studies based their design concepts on 
Saturn era (pre-1970) technology, i.e., state-of-the-art technology 
per the prior study groundrules. In this study* the astrionics module 
was updated to reflect the Shuttle era (1972-1974) technology. The 
propulsion module was not updated as the current studies are in the 
process of updating the Tug technology. It is expected that the pro- 
pulsion module inert weights will be significantly reduced by new high 
performance engine systems, new propuls ion/machani cal systems. Integrated 
astrionics and integrated structures. The engine system for the advanced 
Tug will be specifically designed for earth orbital operations and will 
have lower thrust levels and higher specific Impulses (470 seconds). 

These changes may increase the overall Space Tug mass fraction to 0.860 
or 0.870. The lower inert weights coupled with the improved specific 
impulse will increase the payload capability of the aerobraked Space Tug 
well beyond the payload capabilities shown in the summary and the sensi- 
tivities sections of this report. 

4.2.2 Aerobraking Kit Elements 

The selection of aerobraking kit elements are dependent on the re-entry 
mode used. The Tug could perform the aerobraking re-entry with tumblinq 
or in a controlled planform mode. A tumbling mode was rejected due to 
its Impact on the guidance and navigation systems and upon the need to 
protect larger Tug surface areas from the thermal environments. In the 
controlled planform mode, the Tug may be flown payload end first, side- 
ways or aft end (propulsion module) first. 

The highest temperatures are encountered on the forward end of the 
vehicle. The payload has the lowest temperature capability (300°F) 
of any element of the vehicle. The propulsion module first re-entry 
would have the engine system exposed to the higher thermal environments. 

The engine systems can withstand greater temperatures and have some 
thermal protection already supplied with the conventional Tug concept. 
Therefore, the payload end first trajectory would require more thermal 
protection than the aft end first trajectory. Further, the provision 
for payload thermal protection would have to be of variable length due 
to the differences in length of the potential payloads. 

A sideways controlled mode would require the use of the reaction control 
system continuously to provide static stability. The variable payload 
sizes would change the vehicle configurations from mission to mission 
and would require a mission by mission correction analyses prior to launch. 
The sideways entry also has high sidewall temperatures as this would be 
the forward end of the vehicle. Continual rotation would be required to 
evenly disperse the heat. 


05-17142 


4.2.2 (Continued) 

Therefore* the aft end first mode of controlled entry appears to offer 
the greatest advantage. With this mode, the aft surface presents the 
smallest area to the high temperatures, is partially protected from the 
aerobraking thermal environment by the conventional thermal protection, 
and thus may be protected with the least thermal v/eight penalties of 
any of the re-entry modes. 

In the Initial studies of the aft end re-entry, some thought was given 
to re-entry without the use of an aft heat shield. An examination of 
the temperatures the engine components could withstand indicated that 
large quantities of thermal protection would be required to protect these 
components during aerobraking. Further 0 the complexities of the aero- 
dynamics associated with a protruding nozzle would necessitate that wind 
tunnel type testing be performed to provide realistic aerodynamics input 
data for the control, airloads and trajectory analyses. These problems 
indicated that it would be desirable to provide some aerodynamic and 
thermal protection type aft heat shield to protect the engine system on 
each of the four selected aerobraking configurations investigated. 

For the basic (no flare) configuration, the payload adapter identified 
in the previous Pre-Phase A Space Tug Study could be used. However, an 
extension would be required to totally house the payload in order to 
protect it from the thermal environment. For the configurations employing 
a flare, a payload adapter with additional capabilities for mounting the 
flare and housing the actuation system is required. 

The flared configurations require some aerobraking kit element which will 
Increase the cross sectional area and thus increase the drag. For 
three of the four selected configurations, different angle flares of 
different lengths were investigated. 

The control requirements for the aerobraking Tug can be accomplished 
with the conventional Tug‘s reaction control system with no additional 
thruster and/or support bottles, accumulators, etc. However, the amount 
of reaction control system fuel required will vary from configuration to 
configuration depending on the stability of the configuration, the number 
of passes required and the corrective maneuvers required. 

Hie Tug sidewalls (both the propulsion and astrionics modules) will require 
some additional thermal protection as a result of aerobraking re-entry. 

This can be accomplished with removable sidewall insulation. 

The astrionics systems for the aerobraking trajectory mode will require 
seme additional guidance and navigation sensor systems, Increased 
electrical power capability, and redundant systems for maintenance of 
reliability during long duration missions. The astrionics module systems 
will be discussed in Section 4.5, except for the sidewall thermal pro- 
tection for the astrionics module which is discussed herein. 


4-42 



D5-17142 


4.2.2 (Continued) 

Therefore* the aerobraklna kit elements consist of: 

a. Aft heat shield (Paragraph 4. 2. 2.1) 

b. Sidewall insulation (Paragraph 4.2. 2. 2) 

c. Reaction control systems (Section 4.4) 

d. Astrionics systems (Section 4.6) 

e. Flares (Paragraph 4. 2. 2. 3) 

f. Payload adapters (Paragraph 4. 2.2. 4) 

4. 2. 2.1 Aft Heat Shield 

The aft heat shield is required to protect the engine nozzle and engine 
systems from aerodynamic heating and pressure loads during atmospheric 
braking. This aft protection system must be removable so that during 
the normal Tug operations, the main engine system may be operated. When 
the Tug is loaded into the Shuttle payload bay, the aft heat shield is 
positioned over the engine in the same position as during aerobraking 
operations. After the Tug has been delivered to orbit, the heat shield 
is retracted or removed. 

A large number of heat shield concepts were investigated. The use of 
ablatives was not considered due to the lack of data on the recycle 
capabilities of ablatives. Therefore, by study groundrules, only 
radiative type materials were considered for use as a heat shield. All 
were feasible and differed only in the weight penally associated with 
the heat shield. The metals used in these radiative shields were de- 
pendent upon the maximum temperatures encountered. Rene 41 and TD- 
nickel-chrome were used for the configurations and mission durations 
having the lower temperatures (2000°F or below). The Fansteels were 
used for the configurations and mission durations experiencing the high 
temperatures (above 2000 9 F). 

The selected heat shield used In the performance analyses was the lightest 
weight system. Figure 4, 2. 2. 1-1 illustrates this system. It consists of 
a fixed dome section and a removable cap. The fixed dome is mounted to 
the aft skirt of the propulsion module. The dome section has a one foot 
cylindrical segment (to provide for engine clearance) starting at the aft 
skirt joint. The forward end of the fixed dome portion of the heat shield 
is a 2:1 elliptical contoured dome with a center port opening eight feet 
in diameter. This opening allows for the engine nozzle extension and 
movable gimbal motion during normal Tug operations. The lip on the fixed 
dome port has several 90° turns on its sealing surface to match that on 
the movable cap. 


t 4—43 




/ 


D5-17142 




05-17142 


4, 2.2.1 (.Continued) 

View A-A of Figure 4.2.2.1-T illustrates the latching mechanism. The 
four tie rods are interconnected so that a rotational motion at the key 
pivot point moves the connecting rods which in turn move the latch into 
the desired locked or unlocked position at two points. The details on 
the figure illustrate this mechanism in the locked and unlocked position. 

The movable cap is eight feet in diameter and is tied to the fixed dome 
at two hinged points approximately two feet apart. The hinges allow 
the cap to be rotated approximately 145° outward. The hinges are geared 
to two matching gears connected by a drive shaft and driven by an electric 
motor. 

Figure 4.2.2. 1-2 illustrates an alternative aft heat shield concept. 

This concept has an Internal cap which is rotated out of position during 
normal engine operations. The cap is first retracted several inches to 
lower it below the fixed portion of the dome. Then a gear mechanism 
located on the fixed dome section of the heat shield rotates it 90°. A 
motor located on the cap extends or retracts eight movable arms which 
actuate the latching mechanism. This system is the second lightest 
heat shield concept Investigated, 

Figure 4.2,2. 1-3 illustrates the internally actuated, clam shell concept. 
This concept has a fixed dome heat shield section to the required eight 
foot port opening. The port opening is covered by a two piece clam shell 
during aerobraking. To operate the shell, an electric motor drives a 
gear/pinion mechanism to elevate the clam shell upward. After the shell 
has cleared the fixed dome portion of the aft heat shield sufficiently, 
each side of the clam shell i.s rotated outward approximately 145°. Re- 
traction operations are performed in the reverse order, i.e., first the 
shells are rotated Inward, then they are Towered into the port opening. 

Figure 4. 2. 2,1-4 illustrates the single actuation point, total elliptical 
dome retraction concept. This concept does not use a two piece heat 
shield (fixed dome and movable cao) but rather retracts/emplaces the 
total ellipsoidal portion of the heat shield. When in place on the 
aft end of the vehicle, the aft heat shield is held in place by four 
latches spaced 90° around the vehicle. Heat blocks are provided between 
the aft skirt and cylindrical portions to prevent heat transfer to the 
Tug. Along channels located on either side of the Tug, the heat shield 
actuation system pivot points would be extended aft approximately six 
feet. This would elevate the heat shield away from the vehicle and will 
permit clearance between the shield and nozzle when the shield is rotated 
to the retracted position. This longitudinal movement could be made by a 
small cable/winch/electric motor or by a worm gear/rod system. Upon 
reaching the aft stop, a microswitch would activate an electric motor/gear 
chain to rotate the pivot point, its attached rods, and the entire ellip- 
soidal dome. Another microswitch device would be provided to stop the 
rotation upon touching the Tug's exterior wall. If desired, the pivot 
P£ Tn J c ° u M be drawn forward to its original emplaced position, clearing 
the heat shield from the aft end completely. The main engine nozzle can 


4-47 




D5-17142 



FIGURE 4,2.2»1-4s SINGLE PIVOT POINT-TOTAL ELLIPTICAL DOME RETRACTION CONCEPT 


4*50 



D5-17142 


4.2.2. 1 (Continued) 

then he extended and the engine operated in a normal manner. Although 
the vehicle CG will be asymmetrical, main engine gimballing and RCS trim 
could compensate. 

After deorbit burn and prior to the first perigee pass, the heat shield 
will be elevated, rotated 90°, lowered, emplaced and latched. The method 
is an exact reversal of the steps used to retract. 

Although it is possible that the shield could be retracted after aero- 
braking, it Is not likely because of damage to the external structure 
and possible fusion of the ellipsoidal heat shield to the cylinder. 
Therefore, a jettison backup device must be provided to separate the 
shield from the Tug either at the aft skirt or at the cylinder joints. 

Figure 4.2. 2. 1-5 illustrates the double actuation pivot point, clam 
shell concept. This concept used the movable cap which is divided into 
two nonsymmetric sections. An offset cap joint was used so that the 
maximum stagnation temperature would not occur at the joint, The pivot 
points for the two parts are likewise nonsymmetrical with respect to 
vehicle location. A fixed dome portion of the heat shield extends from 
the Tug aft skirt to the heat shield port opening. 

The operational sequence and requirements are similar to that of previous 
concept. However, this concept has a less complex actuation system. 

Only rotation of the cap sections about the two pivot points is required 
to emplace/retract the cap. The clam shell has the advantage of a more 
symmetrical CG location during normal (non-aerobraked) flight thus re- 
quiring less main engine gimbal or RCS attitude control Impulse. 

Figure 4. 2. 2. 1-6 illustrates the single actuation pivot point, one piece 
cap concept. As with the previous concept, this concept uses the movable 
cap principle. The operational sequence and requirements are basically 
similar to the single point total elliptical dome retraction concept 
substituting the word “cap 1 ’ for "ellipsoid" in the above description. 
Compared to the double pivot point clam shell concept, this concept has 
a more complex actuation system, better cap properties (one-piece), and 
a greater unsynanetrical CG location during normal flight. 

Figure 4,2.2. 1-7 illustrates the single actuation pivot point# clam 
shell concept. This concept could achieve the clam shell opening through 
the use of a single pivot point. It is considered to be the least 
attractive alternate concept investigated. The actuation system com- 
plexity is increased by requiring both translation and rotation. The 
rotation requires either a double-shafted motor (shafts rotating in 
opposite directions) or a clutch arrangement to turn one section in one 
direction followed by an opposite gearing to turn the other section. 

This complexity, coupled with the basic disadvantages of splitting the 
cap, negated the possibility of selecting this concept for the base- 
line. 


* 4-51 



D5-17142 



FIXED ROD CONFIGURATION; 
(TYPICAL 8 PLACES) 


FIGURE 4.2. 2*1 -5: DOUBLE PIVOT POINT-CLAM SHELL CONCEPT 










05-17142 


SIDE VIEW . 
ROTATED 90® 


FIXED GOD-CAP* 
ATTACHMENT 


FIXED SHIELD 
-FIXED ROD- 


POSITION S3 
(RETRACTED) 


NOTEt „ 

END VIEW ROTATED 90 
l 





^ / 



1 


y — 

l 

V 

\ 

\ 

\ 

\ 

\ 

T 

/ r 
/ / 


•POSITION m 
(LONGITUDINAL 
EXTENSION) 


POSITION #1 
(IN PLACE) 


\ L- 





-OVERLAP 
BETWEEN FIXED 
SHIELD & CAP 


■AFT END OF 
AFT SKIRT 


FIGURE 4.2. 2. 1-6: SINGLE PIVOT POINT - ONE PIECE CAP CONCEPT 

4-53 


D5-1 71 42 



POSITION 112 1 PIVOT POINTS SLIDE AFT 3.2* 
TO AFT END OF AFT HEAT 
SHIELD - LIFTING TOTAL CAP 
3.2*. 

POSITION #3t CAP SEGMENTS SEPARATE AND 
ROTATE ABOUT PIVOT POINTS 
TO VEHICLE SIDEWALLS. 


FIXED ROD CONFIGURATION! 



ATTACH POINT 
TO CAP 


FIGURE 4.2.2. 1-7 


SINGLE PIVOT POINT-CLAM SHELL CONCEPT 


D5-17142 


4.2. 2.1 (Continued) 

In addition to the aft heat shield concepts, a forward heat shield mounted 
over the payload was investigated. Figure 4. 2. 2. 1-8 shows the payload 
heat shield operational concept which is identical to that used for the 
propulsion module total dome heat shield retraction system (Figure 
4.2.2 .1 -4) . The pivot points for the shield are extended outward along 
a channel on the Space Tug side wall, then the whole shield is rotated 
clearing the payload. Disadvantages of this concept (and for most con- 
cepts employing heat shields over the payload) are associated v/ith the 
variable payload lengths. Shown in the figure is the forward end of a 
"typical" payload but these payload lengths may be considerably shorter 
or longer. This variable payload length could be compensated for by 
several methods: (1) The three rods supporting the shield on each 
side could be telescoping; (2) the pivot point locations could be made 
more flexible; (3) tailored shields for payload length could be made; 

(4) the shield could be made to fit the longest payload and allow a gap 
for the remainder; or (5) combinations of the above. No analysis has 
been made comparing these methods. 

A six inch overhang is provided for payload airstream protection. This 
fully utilizes the Shuttle's cargo bay diameter constraint. If further 
overhang is required, the ends could be folded alonq the Tug sidewall. 

This latter concept is not shown and would be expected to increase inert 
weight. A 14 foot diameter ring with cross members will provide the 
structural strength and adequate attachment points for the fixed and 
telescoping rods. 

4. 2. 2. 2 Sidewall Thermal Protection System 

The conventional Space Tug load carrying sidewall is designed to with- 
stand temperatures to 300° F. A 400°F limit is used for the micrometeoroid 
shield as the shield does not serve as a load carrying structure. 
Therefore, thermal protection will be required for Space Tug vehicle 
sidewalls when the Tug is operating in an aerobraking mode. The tempera- 
tures on the sidewalls will vary due to (1) distance aft of the re-entry 
heat shield, (2) the aerobraking configuration, and (3) mission duration. 
These factors coupled with the desirability for a removable thermal 
protection system and a minimum Tug weight penalty identified the major 
criteria for selecting the sidewall protection system. In addition to 
protecting the sidewall of the propulsion module, the sidewalls of the 
astrionics module and the payload adapter (basic no flare configuration 
only) must be insulated. 

The basic (no flare) Tug sidewall temperatures (from the heat shield to 
the payload) will vary from 1100°F to 1600°F (one day* mission) down to 
700°F to 1000°F (10 day missions). The flared configurations will not 
experience temperatures as severe. The temperatures will vary from 
800°F to 1050°F (one day mission) down to 450°F to 600°F (10 day mission). 
Figure 4. 7. 2. 0-1 in Section 4.7 tabulates the expected sidewall tempera- 
ture. 


4-55 


D5-17U2 


POSITION #3 
(CLEARING P/L) 


\ 


V - 




POSITION n 
(EXTENDED) 






r?*ZZZ\ 


I _CIRCpCAR_RIMG 
if JNDOFP/L _ 


-POSITION #1 
(IN-PLACE) 


1 

k^ ! 


POSITION Hi 
(CLEAR & READY 
TO INTERCHANGE 
P/L'S) 


S \ \\ 

\\\' 



f TELESCOPING -v 
ARMS RETRACT > 
AFT OF P/L - D.A. 
LINE TO PROVIDE 
CLEARANCE 
(2 PLACES - 180° 
APART) 


PAYLOAD 


DOCKING 

ADAPTER 


ASTRIONICS 

MODULE 



PROPULSION 

MODULE 


FIGURE 4.2.2.1-8. PAYLOAD END HEAT SHIELD CONCEPT 





D5-17142 


4. H. 2. 2 (Continued) 

The material selected for the insulation must have a wide range of 
temperature capability and be light weight, John Mansville microquartz* 
a fiberous material, was selected. It has a low density of three pounds 
per cubic foot and a long life temperature capability of 2000°F. Its 
thermal conductivity varies from 0,36 at 3600°F to 0.91 BTU/hr/ft z /1n./°F 
at 1000°F. This material can be cut to the desired shape and tapered 
down from the greater thicknesses required at the heat shield/sidewall to 
the lesser thickness required at the end of the payload sidewall. It can 
be mounted and bonded atop the aluminum micrometeoroid shield of the 
propulsion module, over the louver doors of the astrionics module and 
atop the aluminum skin of the payload adapter. To prevent damage to 
the microquartz during handling and transportation, a thin foil (0.002 
inch) of titanium or inconel may be used. Figure 4.2.2 .2-1 illustrates 
the sidewall insulation for the 30 pass (5 day) mission using the 60° 
flared aerobraking Space Tug concept. 

4. 2. 2. 3 Aerodynamic Flare 

The basic (no flare) Tug configuration experiences severe thermal 
environments due to its low drag/high ballistic coefficient. One method 
of increasing the drag and thereby reducing the ballistic coefficient is 
through the use of an aerodynamic flare. Several design concepts were 
investigated ranging from silicon rubber inflated bags for low tempera- 
ture environment flares to all metallic flares for high temperature 
environments. 

The selected flare concept is a metallic flare as shown for the 60° 
flare in Figure 4. 2. 2, 3-1. The same concept is applicable to the 30 and 
45° flares and is the lightest weight concept investigated. Because the 
flares will encounter temperatures ranging from approximately 1300°F 
(30° flare, 5 pass mission) down to 540°F (60° flare, 60 pass mission), 
an Inconel 718 was used for the flare skin. 

The baseline aerodynamic flare concept is built from 72 panels. The 
panels are designed so that when they are retracted, they will not 
exceed the 15 foot diameter Shuttle bay limit. These are 36 intermediate 
panels and 18 (each) inner and outer panels. The intermediate panels do 
not extend to the Space Tug outer skin and thus reduce the overlapping 
of panels which will reduce the overall flare weight. 

The flare actuation sequence is as follows. The aerodynamic flare is 
in place and in the retracted position when delivered to orbit by the 
Shuttle. After the Tug has accomplished its mission and has deorbited 
for the aerobraking return, the flare is extended. First the electric 
motor releases the 36 cables connected to the 36 struts. The cable con- 
nects to the two segment struts at the strut mid-point. The flare is 
connected to the Tug wall by spring hinges. Similarly the strut is 
connected to the Tug by spring hinges. Releasing the cables allow 
these spring hinges to elevate the flare and to extend the strut seg- 
ments. The strut will extend until the lower strut segment and the upper 
strut segment are in line and then the two sections lock in place. When 


4-57 



83"fr 




SIDEWALL INSULATION REQUIREMENTS 

LENGTH 

(FEET) 

5 

10 

15 

20 

25 

30 

35 

40 

THICK- 

NESS 

(INCHES) 

0.5 

0.46 

0*42 

0.39 

0*35 

0.32 

0*23 

0.24 


INSULATION - MICROQUARTZ (3 S/FT 2 ) 

OUTER SKIN - TITANIUM OR INCONEL 
CONSTANT THICKNESS 
0.002 INCH 


FIGURE 4. 2.2. 2-1. SIDEWALL INSULATION - 60 DEGREE FLARE SPACE TUG 

CONFIGURATION - 30 PASS (5 DAY MISSION) 


D5-17142 










FRONT VIEW OF 
RETRACTED FLARE 




SIDE VIEW OF 
RETRACTED FLARE 



i r o i ft. 


FURR STRUT -35 REQUIRE* 



retracted pure details 


-15* DIAMETER 
SHUTTLE BAY CC}i$t»ANT 


FIGURE 4. 2.2. 3-1 . BASELINE FLARE CONCEPT 



05-17142 


4. 2. 2. 3 (Continued) 

aerobraking Is complete# the flare is retracted by actuation of the 
electric motor which rewinds the 36 cables on the cable drum. The 
retracting cables collapse the strut segments which draws the flare 
and the struts inward toward the Tug sidewall. The panels are folded 
so that the 18 outer panels formed the collapsed outer flare surface. 

The 18 inner panels are against the Tug sidewalls. The 36 intermediate 
panels are folded between the inner and outer panels. When folded# the 
flare Increases the Tug diameter to 14'6"» still within the 
Shuttle cargo bay limit. 

The 72 panels are connected by piano hinges in an inner panel, inter- 
mediate panel, outer panel# intermediate panel -back- to-inner panel 
pattern as shown in Figure 4. 2. 2. 3-1. The inner and outer panels are 
0.02 inches thick and are reinforced by ribs 0.50 inches thick extending 
the entire slant height length and along both of the piano hinge joints. 
The outer edge is similarly ribbed. The intermediate panels are also 
0.02 inches thick and have 0.50 inch thick ribs along each side and on 
the outer edge. 

The skin thicknesses shown on the figure are based on the "worst case" 
thickness requirements. For the 60 pass# 10 day mission# a thickness of 
only 0.006 inches would be required. As a practical fabrication thick- 
ness# an 0.010 inch thickness was considered the minimum handling and 
processing thickness. 

To provide bending support for the extended panels, 36 support struts# 
evenly spaced at 10° increments are used. The struts are tubular titanium 
1-1/4 inch diameter and 1/32 inch wall thickness. The struts vary in 
diameter from 1 to 1-3/4 inches and from 0.030 to 0.0625 inches in wall 
thickness. The struts are composed of two equal sections per strut. 

The cables join the strut at the joint between the sections. The cables 
then are wound around pulleys located approximately 56 inches down the 
payload adapter (aft of the astrionics module). The cables then follow 
the payload adapter wall until another set of pulleys located at the 
hinge points redirect the cables Inward toward the cable drum located 
in the forward section of the payload adapter. 

Several alternative flare concepts were investigated. Figure 4.2. 2. 3-2 
Illustrates another metallic panel concept (upper portion of figure) and 
an Inflatable bag concept (lower portion of figure). The metallic concept 
uses 10 sets of panels located 36° apart. Each set consists of an Inner 
panel and two outer panels. Each of the outer panels divide into two 
sub-panels. Upon actuation the outmost sub-panels slide outward until 
the subsections extended to the ends of the middle panels. Then the 
middle panels separate until the sub-panels reach the ends of the inner 
panel. Retraction will be performed in the reverse order. 

The ten sets of panels are elevated by 10 spring leafs located at the Tug 
sidewall/flare joint. Twenty support struts# spring actuated* provide 
flare rigidity. Each strut consists of a two piece support section 


4-61 


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D5-17142 



05-17142 


4. 2.2. 3 (Continued) 

which locks the flare In the extended position. Retraction is accom- 
plished by retracting the 20 cables joining the mid-points of the 
support legs. This concept was not used because the overlapping panels 
were heavy and friction may prevent the panels from sliding properly. 

The Inflated bag concept (shown in the lower portion of the figure) 
uses a Kapton or silicon rubber bag to elevate the flare. The bag 
may be used for long duration missions where the temperature on the 
flare would not exceed approximately 700°F. A metal panel with insula- 
tion on the surface facing the bag may be used to cover the bag for high 
temperature use. Helium is recommended for use to inflate the bags as 
helium bottles were already proposed to pressurize the oxygen tank of 
the Tug. By partial inflation, a variable angle flare could be obtained 
for changing the drag and ballistic coefficient. This would permit 
variable mission durations and/or corrections to the trajectory through 
modifying the flare angle. This concept is practical. However, there 
are complex design problems associated with this concept which could 
not be investigated and therefore it was not selected as the baseline 
flare concept. 

The concepts studied for the flare configuration were limited by study 
duration and by funding limitations. There are other flare concepts, 
other drag devices, other locations and other actuation methods which 
may perform equally as well as the baseline flare concept. These options 
should be studied in future aerobraking studies. 

4. 2. 2. 4 Payload Adapter 

The payload adapter for the basic (no flare) configuration is mounted 
atop the astrionics module and serves only to hold down the payload and 
protect it from the thermal environments. The payload adapter for the 
flare configuration aerobraking Space Tug has the additional functions 
of serving as the flare mounting fixture and the housing structure for 
the flare actuation mechanism. Figure 4. 2. 2. 4-1 illustrates the payload 
adapter for the flared configurations* 

The payload adapter would be the same diameter (14 feet) as the propul- 
sion and astrionics modules. The length would be approximately 10 feet 
long. The structure is an aluminum stiffened skin structure. The stiffener 
consists of 36 stringers which also serve as the mounting structure for 
the 36 flare support struts. Three ring frames are employed. The 
forward ring frame is used to bolt the payload adapter to the astrionics 
module. The intermediate ring flare is welded to the 36 stringers at 
the ooint where the flare support struts are mounted (flared configurations 
only). The aft ring frame is used to support the payload guide structure. 

For the flared configurations only, the forward ring frame also is used 
to mount the flare spring hinge and to mount 36 pulleys used to control 
the cables during flare actuation. This area is also reinforced by 
cross beams which are used to mount the electric motor and the cable 


*4-63 



4-64 


INTEGRATED 

PAYLOAD 

ADAPTERS 



representative - 

PAYLOAD . 


STRINGER 


PAYLOAD DOCKING 
GUIDE CONE 


SECTION FOR 
MOUNTING 
ACTUATION 
DEVICES 


RING FRAMES 



FIGURE 4. 2. 2.4-1: PAYLOAD ADAPTER CONCEPT FOR FLARED AEROBRAKING TUG CONCEPTS 


D5-17142 


D5-17142 


4. 2. 2. 4 (Continued) 

drum. The flare actuation system is housed within the first six inches 
of length. 

At the aft end of the payload adapter is a payload guide cone. This is 
a conical frustum 14 feet diameter at the entrance, six inches long 
and 13 feet at the base. Another cylindrical extension six inches long, 
13 feet in diameter completes the guide cone structure. This device is 
used to assist the payload docking operations by overcoming possible 
misalignment between the payload and Tug. Four guide tubes {90° apart) 
extending the length of the payload adapter also help to align the pay- 
load. 

The payload is held in position by the payload holddown fixtures. These 
fixtures are located at four places, 90° apart, five feet from the front 
of the payload adapter. These devices would be solenoid actuated. 

The skin surface of the payload adapter is a thin aluminum sheet. For 
the basic (no flare) configuration, microquartz insulation with an outer 
titanium foil cover provides thermal protection to 300°F for the payload. 
For the flared configurations the temperatures will not be as severe. 

No additional protection other than the aluminum skin is expected to be 
required except for short duration missions. For these missions, some 
protective insulation {microquartz with the titanium foil) may be re- 
quired. 

The base area of the payload adapter for the no flare aerobraking Tug 
configuration will require a protective cover. The cover may be opened 
and closed with an electric motor/gear mechanism or by a spring loaded 
cable mechanism. The cover would be microquartz insulation between an 
aluminum inner cover and a titanium outer cover. This cover may also 
be required for the short mission duration flared Tug configurations. 

The weight of the payload adapter for the basic (no flare) configuration 
is 400 pounds (Including end cover). This is higher than the 390 pounds 
required for the flared configuration. The difference between the no 
flare and flare Tug payload adapter weights is the requirement for the 
crossbeams - 40 pounds added weight - for mounting the actuation 
mechanisms for the flared configurations and the deletion of the end 
cover (flared configuration only) - 50 pounds. This Integrated pay- 
load adapter/flare mounting fixture/flare actuation device housing 
fixture is one of several concepts investigated. No significant weight 
difference among the various concepts investigated were observed. This 
one was selected as the baseline payload adapter configuration. 

4,2.3 Configuration Design Options and Recommendations •, . 

The aerobraking kit elements consist of the aft heat shield, sidewall 
insulation, modified astrionics, flares and payload adapters. Each of 
these kit elements may have alternatives which wilt reduce weight, 
complexity, or may modify operational methods or increase versatility. 


4-65 


D5-17142 


4.2.3 (Continued) 

This study could only examine several representative, practical, low 
weight systems. The kit alternatives should be investigated in more 
depth. 

For example, the aft heat shield could be replaced by an ablative shield, 
by exhaust gas cooling techniques, or by a forward aerodynamic flare. 
Different nose shapes could be used in lieu of the 2:1 ellipsoidal shape 
including hemispherical and blunted conical frustrum shapes. 

The sidewall insulation may be replaced with ablatives, foam insulation, 
or insulations used in conjunction with the micrometeoroid shielding. 
Methods of providing heat blocks, accounting for heat leaks, pumping of 
insulation and insulation shielding techniques require investigation. 

The astrionics module is covered with the same sidewall insulation as 
the propulsion module so the above comments apply to it also. The as- 
trionics guidance and navigation systems are discussed in Section 4.6. 
Methods of reducing weight of the electrical systems, determining the 
reliability and accuracy of existing and new sensor systems require 
investi gat ion. 

The flare concept is only one of many practical concepts. For larger 
flares, the weights would become prohibitive using this concept. Low 
weight alternatives require investigation including open flare structure 
near to the Tug's sidewall, split flares, inflatable bag flares, lift to 
drag flares, and steeper angle or flat plate flares. 

The payload adapter is normally not considered a part of the aerobraking 
kit. However, for the flared configurations its multipurpose made it a 
trade variable. Only limited design activity was undertaken to provide 
some understanding of the weight penalty, payload environmental considera- 
tions on the aerobraking mode and upon aerobraking kit design. More 
detailed analysis of the payload/Space Tug interfaces is required. 

In summary, the configuration options considered for the aerobraking kit 
elements were practical. Low design risk kit elements were selected to 
operate with the existing Boeing Pre-Phase A Space Tug Concept. The 
payload capability obtainable with the aerobraking kit elements were 
identified. The major considerations were directed toward low weight 
as the aerobraking kit weight is traded pound for pound with round trip 
payload. Designs were prepared only to the degree necessary to determine 
inert weights. Follow-on activities should consider the design of 
aerobraking kits that are reusable and can be refurbished. The design 
Impact of aerobraking on the Shuttle/Tug interfaces such as handling and 
transporting and environmental constraints of Shu ttle/Tug/pay load are 
secondary considerations but should be evaluated in future studies. 

The design of the conventional Tug should be investigated to determine 
what design modifications should be considered to provide the best 
compromise Tug design applicable to both aerobraking and non-aerobraking 
applications. 


4-66 


D5-17142 


4.3 TRAJECTORY ANALYSIS 

The baseline mission selected for analysis in this study was 
the geosynchronous round-trip payload mission, since this 
mission imposed the most stringent performance requirements 
on conventional tug systems. In this mission profile the tug 
is deployed by the shuttle in 100 NM, 28° inclination orbit. 

The tug transfers the payload via Hohmann transfer to 
equatorial synchronous orbit with the 28° plane change being 
made at apogee. In synchronous orbit the payload is exchanged 
for an equal weight return payload. At the proper time the 
tug and payload are placed on the aerobraking return ellipse 
by applying the deorbit and 28° plane change AV. During the 
aerobraking return the apogee of the ellipse is reduced to 
270 NM by drag dissipation of the orbital kinetic energy on 
one or more passes through the upper atmosphere. The orbit 
is circularized at 270 NM for phasing with the shuttle at 
100 NM and at the proper time a final Hohmann transfer is made 
to 100 NM for rendezvous with the shuttle. 

Initial trajectory analysis consisted of generating preliminary 
trajectories with constant ballistic coefficients (W/C^A) and 
estimated drag coefficients {Cd) to establish the flight regime 
for the aerodynamic analysis. The drag coefficient was varied 
parametrically to vary the ballistic coefficient from 10 PSF 
to 80 PSF, the expected range for the tug. These trajectories 
established the altitude and velocity range that the expected 
configurations would traverse. Final trajectory analysis 
consisted of generating trajectories using the refined aero- 
dynamic data. These trajectories were then used for thermal, 
astrionics, control, and loads analyses to determine subsystem 
weight penalties associated with the aerobraking concept. 

4.3.1 Analytical Model 

Space Tug trajectories involve flight through two different 
mediums; an atmosphere and a vacuum. As a result, it is 
expedient to simulate each type of flight differently to reduce 
computer time and maintain integration accuracy. The following 
technique was used to generate trajectory data for this study. 

o Flight through the sensible' atmosphere was simulated 
by numerically integrating the equations of motion in 
rectangular coordinates. 

o Vacuum flight was simulated by conic segments where 
the orbital elements are determined based on Kepler's 
laws . 

o The total flight path was constructed by patching 
together the conic trajectory and the integrated 
trajectory as illustrated in Figure 4. 3. 1.0-1. 


4-67 


FIGURE 4.3. 1.0-1 MISSION PROFILE 


-END OF \ 
LAST PASS 

AV* 1 




tSISsFIR : 

ELLIPSE, 48* 


- W 'V 


// ' 


-ATMQSPSIERE ENTRY ALT 
600000 FT (NUMERICAL 
INTEGRATION LIMIT} 


.8 


D5-17742 


4.3.1. 1 Assumptions and Groundrules 


The earth model, tug configuration constants and assumptions 
pertinent to the mission profile shown in Figure 4.3. 1.0-1 are 
given below (References 4. 3. 1,1-1, -2, and -3). 

Mission and Trajectory Assumptions and Constants 

1. Initial Orbit ~ 100 NM circular 

2. Synchronous orbit altitude = 117440496 ft 

(19316 NM) 

3* Intermediate orbit at conclusion of aerobraking = 
270 NM circular 

4. Pinal orbit = 100 NM circular 


5. Mission total AV = 22400 fps 


a. AV to synchronous orbit = 14550 fps 

"(includes 400 fps rendezvous) 

b. AV to start aerobrake descent - 6000 fps 

(descent trajectory assumed to have a 250 KFT 
perigee) 


c. AV at end of aerobraking ~ 1850 fps 

(includes 400 fps rendezvous and 400 fps reserve) 

6. The vacuum perigee at atmosphere entry on a pass is 
the vacuum perigee that results from the decaying 
orbit of the previous pass, i.e., the orbit is 
allowed to decay naturally. No corrections are made 
in perigee altitude. (This applies only to the 
nominal atmosphere trajectories.) 

Tug Constants 


1. Total initial weight of tug as 
deployed by shuttle 

2. Total propellant weight 

3. Entry weight at start aerobrake 
(Based on using the 45000 lbs 
of propellant to perform the 
AV of 22400 fps) 

4. Main engine specific impulse, I sp 

5. Tug reference area, A 

. Lift-to-drag ratio, L/D 
(zero angle of attack) 


= 57740 lbs 
=3 45000 lbs 
« 14430 lbs 

= 460 sec 
= 154 ft 2 


4-69 ' 


6 


0 


05-17142 


4. 3.1.1 (Continued) 

Earth Model Constants and Assumptions 

1. Spherical, rotating earth and atassophesQ 

2. Limit of sensible atmosphere o CG0,GOQ ffe alMfel3© 

3. Nominal atmosphere - 1962 U. 0. Standard 

4. Equatorial radius = 20925722 ft 

5. Gravity constant = 14.076539 X 10^® ft^/eec^ 

6. Earth rotation rate = .729211505 X 10” ^ sa3/oc@ 

4.3.2 Preliminary Trajectory Analysis 

The preliminary trajectory analysis using constant ballistic 
coefficients (w/CpA) produced the data shown in Figure 4.3. 2.0-1. 
This data was generated to establish the basic relationships 
between W/CpA, mission time, and initial perigee altitude* 

These data identify the reasonable tug configuration requirements 
and associated altitude and velocity ranges necessary to develop 
preliminary aerobraking design concepts and determine associated 
aerodynamic characteristics. 

% 

4.3.3- Final Trajectory Analysis 

Aerodynamic analyses (Section 4.1) identified the following 
tug configurations for detailed evaluation in this study. 

1, Basic Tug (no flare) 

2, 30° Flare (short flare) 

3, 45° Flare (Intermediate Flare) 

4; 60° Flare (Large Flare) 

The vacuum perigee altitude of the initial descent trajectory 
from synchronous orbit was varied to change the total number 
of passes (and therefore return time) required for apogee 
decay to 270 NM. (This perigee is the equivalent Kepler orbit 
perigee of the first atmospheric passy it is not the actual 
low point of the trajectory in the atmosphere.) Results for 
the four selected tug aerobraking configurations are presented 
in Figure 4.3, 3. 0-1, Data for an optional configuration (a long 
30° flare) is also shown. The initial vacuum perigee was varied 
for each configuration for entry pass numbers between one 
(.24 days) and 60 (11 days). The relationship between number 


4-70 


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FIGURE 4.3. 3.0-1 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - INITIAL PERIGEE ALTITUDE 


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D5-17142 


4.3.3 (Continued) 

of passes and return time proved to be nearly linear and independent of 
the Tug configuration. This relationship is shown in Figure 4. 3. 3. 0-2. 

The apogee altitude decay is shown as a function of pass number for the 
basic Tug and the 60° flare configurations in Fiqures 4. 3. 3. 0-3 and 

4. 3. 3. 0- 4. The apogee shown is the apogee that occurs at the end of 
each individual pass. Only these two configurations are shown, since 
they represent the limiting extremes in W/CqA. It should be noted that 
there is no appreciable difference in the apoqee decay of these two con- 
figurations. The other configurations have similar apogee decay history. 

The vacuum perigee altitude at the end of each individual pass is pre- 
sented in Figures 4. 3. 3. 0-5 and 4. 3. 3. 0-6 for the basic Tug and the 
60° flare configurations. This perigee is the equivalent vacuum perigee 
at atmospheric exit for the pass, and it is the vacuum perigee of the 

orbit at atmosphere entry on the descent leg of the next pass. These 

curves are presented for the two extreme Tug configurations. 

The maximum dynamic pressure as a function of pass number is presented 
in Figures 4.3. 3. 0-7 through 4. 3. 3. 0-9 for all of the Tug configurations 
analyzed. It should be noted that the highest Q^ax occurs on the last 
pass for all configurations. This is caused by the sharp drop in perigee 

altitude during the last few trajectory passes, as illustrated in Figures 

4.3. 3. 0- 5 and -6. 

Figures 4.3.3.0-10 and -11 show the maximum inertial velocity attained on 
each pass as a function of the number of passes in the mission. The 
Inertial velocities are relatively insensitive to the configurations 
as shown in the figures and as tabulated below. On the longer missions 
(30 and 60 passes), the decrease in maximum velocity per pass is nearly 
linear for the majority of the mission. The latter part of the mission 
has an increasingly larqer drop in maximum velocity. The shorter missions 
(5 and 10 pass) have nearly linear decreases. The first pass has a 
maximum velocity of approximately 34,000 ft/sec for all mission times 
shown with the last pass minimums of approximately 26,700 to 28,800 ft/sec 
for both the basic (no flare) and the 60° flare configurations. The 
maximum inertial velocities shown do not occur at perigee because of the 
braking effects of the atmosphere. Rather, these values occur from a 
few seconds to approximately two minutes prior to perigee dependent on the 
number of mission passes and the individual pass number. For comparison, 


4-73 ■ 


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PASS IUHIEN 


FIGURE 4. 3.3. 0-4 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - APOGEE ALTITUDE DECAY 










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• ENTRY ALTITUDE INTO ATMOSPHERE ■ 400000. FT 

• CONTINUUM FLOW W/CqA • 9.94*11.90 
•FINAL APOGEE ALTITUDE • 1640551. IIS FT 


FIGURE 4. 3. 3. 0-6 SPACE TUG AEROBRAKING RETURN 'FROM SYNCHRONOUS ORBIT - PERIGEE ALTITUDE DECAY 


D5-17142 




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1 !• ENTRt ALTITUDE INTO ATMOSPHERE • 600000 
!• CONTINUUM FLON H/CnA -9.94-11.90 
!#f* F1HAL ArOCEt ALTITUDE • 1640SS1 . 185 FI 


MAXIMUM DYNAMIC PRESStWf- 


SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT 


FIGURE 4. 3. 3. 0-8 













MAXIMUM INERTIAL VELOCITY (THOUSANDS OF FRET PER SECOND) 


D5-17142 



0 10 20 30 40 50 60 

NUM8ER OF PASSES 

FIGURE 4.3.3.0-10. MAXIMUM INERTIAL VELOCITY PER PASS 

BASIC TUG (NO FLARE) CONFIGURATION 


4-82 


MAXIMUM INERTIAL VELOCITY (THOUSANDS OF FEET PER SECOND) 







D5-17142 


\ 


4.3.3 (Continued) 

selected maximum inertial and periaee inertial velocities are shown below: 


BASIC (NO FLARE) TUB CONFIGURATION 


No. of Passes 

Pass 

Maximum Inertial 

Perigee Inertial 

in Mission 

No. 

Velocity (ft/sec) 

Velocity (ft/sec) 

5 

1 

33,895 

33,409 

5 

5 

28,768 

27,650 

10 

1 

33,892 

33,704 

10 

5 

31 ,683 

31,415 

10 

10 

27,849 

27,165 

30 

1 

33,889 

33,831 

30 

15 

31,331 

31 ,259 

30 

30 

26,989 

26,651 

60 

1 

33,885 

33,869 

60 

30 

31,249 

31,227 

60 

60 

26,669 

26,470 


60° 

FLARE TUG CONFIGURATION 


No. of Passes 

Pass 

Maximum Inertial 

Periqee Inertial 

in Mission 

No. 

Velocity (ft/sec) 

Velocity (ft/sec) 

5 

1 

33,854 

33,381 

5 

5 

28,824 

27,738 

10 

1 

33,857 

33,672 

10 

5 

31,709 

31 ,478 

10 

10 

27,888 

27,178 

30 

1 

33,854 

33,814 

30 

15 

31,343 

31,282 

30 

30 

27,000 

26,653 

60 

1 

33,850 

33,833 

60 

30 

31,258 

31,232 

60 

60 

26,667 

26,452 


4-84 


f 


1 


D5-17142 


4.3.3 (Continued) 

Figure 4.3.3.0-12 illustrates the relationship of time spent in a free 
space environment (above 600,000 feet altitude) to the time spent in the 
atmosphere (below 600,000 feet) for the basic (no flare) configuration. 
The upper portion of the figure represents the time per pass (orbit) 
when the Tug is above 600,000 feet. Subsequent to the Initial 18,700 
second coast period from geosynchronous orbit to the first pass re-entry 
and its atmospheric passage, the 30 pass mission Tug will be in a free 
space environment for approximately 34,000 seconds (9.44 hours) prior 
to its second pass re-entry. The succeeding free space environment times 
decrease as the number of passes increases. Between the 29th and 30th 
passes, the Tug's time above 600,000 feet is decreased to 5156 seconds 
(1.43 hours). These long duration free space environment times can be 
regarded as vehicle "cool-down" periods as contrasted to the short 
duration "heating" periods below 600,000 feet. 

The bottom portion of Figure 4.3.3.0-12 represents the time per pass 
spent below 600,000 feet. Using the same example as above, the 30 pass 
mission Tug is below 600,000 feet for 337 seconds on the first pass. 

This time Increases as shown until the longest time (1194 seconds) per 
pass is reached on the 30th (final) pass. For all of the missions ex- 
amined, the total time spent below 600,000 feet is less than 45£ of the 
total aerobraking mission time. 

The velocities shown in Figure 4.3.3.0-10 Indicated acceleration experi- 
enced with atmospheric passage which displaced the maximum velocity from 
the normal conical perigee for the basic (no flare) configuration. This 
is reflected in the differences in times down to and up from perigee. 
Representative times for different passes of different mission durations 
are listed below. 


BASIC (NO FLARE) 


No. of 
Passes 
in Mission 

Pass 

No. 

Total Time Spent 
Below 600,000 ft 
(sec) 

5 

1 

363 

5 

5 

1,017 

10 

1 

348 

10 

5 

422 

10 

10 

1,110 

30 

1 

337 

30 

15 

414 

30 

30 

1 ,194 

60 

1 

325 

60 

30 

407 

60 

60 

1*333 


CONFIGURATION 

Time from 600,000 Time from Perigee 
ft to Perigee to 600,000 Feet 
(sec) (sec) 


176 

187 

320 

697 

168 

180 

208 

214 

374 

736 

168 

169 

206 

208 

480 

714 

160 

165 

200 

207 

S44 

789 




4-85 


TIME (THOUSANDS OF SECONDS) 


D5-17142 


( 



4*86 



D5-17142 


4.3.3 (Continued) 

The column "Time from 600,000 Feet to Perigee" is significant in that 
this value is one factor in the minimum allowable time for a final mid- 
course correction burn prior to re-entry. As such* it impacts the 
navigation and accuracy analysis and the RCS requirements. 

Similarly, Figure 4.3.3.0-13 illustrates the same data for the 60° flared 
configuration. The shape of the curves is similar to that presented for 
the basic (no flare) configuration. Because of the higher perigees flown 
with the flares* slightly less of the total mission time is spent below 
600,000 feet. Representative time divisions between "600,000 Feet to 
Perigee" and "Perigee to 600,000 Feet" are shown below. 


60° FLARED CONFIGURATION 


No. of 
Passes 
in Mission 

Pass 

No. 

Total Time Spent 
Below 600.000 Ft 
(sec) 

Time from 600,000 
Feet to Perigee 
(sec) 

Time from Perigee 
to 600,000 Feet 
(sec) 

5 

1 

344 

168 

176 

S 

5 

956 

296 

660 

10 

1 

330 

160 

170 

10 

5 

397 

192 

205 

10 

10 

1,041 

352 

689 

3Q 

1 

313 

152 

161 

30 

IS 

389 

192 

197 

30 

30 

1,174 

448 

726 

60 

1 

304 

150 

154 

60 

30 

379 

187 

192 

60 

60 

1,252 

512 

740 


4.3.4 Dispersed Atmosphere Effects 

The effect of atmospheric dispersion on the aerobrafcing trajectory of the 
basic Tug and large flare (60°) configurations are presented in this 
section. In addition, two candidate schemas for controlling the tra- 
jectories once a dispersion has been encountered are presented. 

The dispersed atmosphere models (Reference 4. 3, 4. 0-1) used are given in 
Figures 4.3.4. 0-1 and -2, The more dense atmosphere used was the summer, 
high solar activity model of Figure 4.3.4.0-1 (* density). The less dense 
atmosphere used was the winter, low solar activity model of Figure 
4. 3, 4. 0-2 (~ density). These dispersed atmosphere density versus alti- 
tude functions were assumed to be invariant with time during the period 
of an entry trajectory. 


TIMS (THOUSANDS OF SECONDS) 


,® STANDARD ATMOSPHERE. 
© CIRCULARIZE § 270 M.M. 


TIME PER PASS ABOVE £00,000 FEET ALTITUDE 
(TIME FROM EXIT TO RE-ENTRY) 


TOTAL TIME SPENT BELOW £00,000 FEET ALTITUDE! 
5 PASS - 0.72 HOUR OF 21.6 HOURS (3.3%) 

10 PASS - 1.33 HOURS OF 43.2 HOURS (3.1%) 

SO PASS - 3.76 HOURS OF 131 HOURS (2.9%) 

60 PASS - 7.34 HOURS OF 264 HOURS (2.8%) 


TIME PER PASS BELOV/ 600,000 FEET ALTITUDE 
(TIME DURING RE-ENTRY) 


a 10 .20 30 ’ 40 SO 60 

HUMBER OF PASSES 


FIGURE 4.3.3.0-13: COMPARISON OF FREE SPACE ANO RE-ENTRY TIMES (60* FLARE) 







SUMMER 


i 

co 

VO 


s: 

5kS 


o 

H 


-~ J 250? 

! - : r ■■■ : ! — L-^§ 200 KM 



Preliminary Estimates of the mean and range of density 
lover an eleven year period* expressed as a relative 
^ difference t«) from the US Standard Atmosphere 1962* 
P-TTTT for shuttle re-entry in Cape Kennedy* Florida* area 


( The influence of solar 
activity on range of density 
^Increases with altitude and 
?n * becomes of principle 
* w s significance above the 
r 60 to 90 fen altitude range. 


: : : 5 j • • it / 
f d * ositv 


-.60 


;/ 


a' 1 during summer. For use in preliminary engineering 
estimates only. Further use should not be made without 
contacting personnel of tne Aerospace Environment Division. 
The mean constitutes a systematic variation which Is 
> r , ^predictable for a given month and solar activity level* 

* jji A-l - The range constitutes a random variability which is not 

predictable. The Geomagnetic storm variation is for an 
' " ’] «7f extreme event which occurs once or twice per solar cycle 
and is not predictable a few days in advance, it usually 
* occurs with medium or high solar activi t; . The, effects 

/;i do not propogate into the lower altitudes. 


*40 


60 


80 


100 


120 


140 


160 


PERCENT DEPARTURE FROM 1962 U. S. STANDARD DENSITY 


FIGURE 4.3.4.Q-1 


SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - SUMMER 
ATMOSPHERIC DENSITY DISPERSIONS 


D5-17142 


WINTER 


t 

U3 

O 



Preliminary Estimates of the mean and range of density 
over an eleven year period, expressed as a relative 
difference (5) from the US Standard Atmosphere 1962, 
for shuttle re-entry in Cape Kennedy, Florida area 
during winter. For use in preliminary engineering 
estimates only. Further use should not be made without 
contacting personnel of the Aerospace Environment Division. 
The mean constitutes a systematic variation which is 
predictable for a given month and solar activity level. 

The range constitutes a random variability which Is not 
predictable. The Geomagnetic storm variation is for 
an extreme event which occurs once or twice per solar 
cycle and is not predictable a few days in advance. 

It usually occurs with medium or high solar activity. 

The effects do not propogate into the lower altitudes- 

40 60 00 TOO 120 140 160 


nOTEz 


pThe influence of 
jjq |-splar activity on 

• ranee of density 

* increases with altitude 
land becomes of principle 
.significance above the 
?5G to 90 fcr altitude 

*0 range. 

i * * - ! ___ t 


0 

-30 


-60 



1 -I A 'X 
I - f rr 


PERCENT DEPARTURE FROM 1962 U. S. STANDARD DENSITY 

. 

FIGURE 4.3.4. 0-2', SPACE TUG AER0BRAKING RETURN FROM SYNCHRONOUS ORBIT - WINTER 

ATMOSPHERIC DENSITY DISPERSIONS 


D5-17142 


D5-17142 


4.3.4 (Continued) 

Figure 4. 3. 4. 0-3 presents the initial perigee altitude required for decay 
to a 270 NM orbit for the 1962 standard atmosphere and the dispersed 
atmospheres. The solid curves of Figure 4.3. 4.0-3 are a result of re- 
isolation of the initial perigee altitude required to force the trajectory 
to end up at 270 NM in the presence of the dispersed atmosphere. The 
data illustrates the range of entry time due to atmospheric dispersions if 
the vehicle flies an uncorrected trajectory. For example* the flight 
time of the basic Tug nominal 30 pass trajectory varies from 3.6 days (19 
passes) for the more dense atmosphere to 9.6 days (53 passes) for the less 
dense atmosphere. This large range in entry time implies the need for a 
trajectory correction technique that will significantly reduce this range. 

4. 3. 4.1 Trajectory Correction Techniques 

The following analysis pertains to a more dense atmosphere. The same 
logic will apply in the reverse direction to a less dense atmosphere. 

When descent begins* the density of the atmosphere that is to be encountered 
is not known, therefore* descent will be made on a trajectory that is tar- 
geted to the initial perigee altitude dictated by the nominal atmosphere. 

If the atmosphere is more dense than that expected* the trajectory apogee 
will be too low on this pass and all succeeding passes. The result is 
that the Tug will reach the target altitude much too soon. This means 
that adjustments must be made in the trajectory in order to reach 270 NM 
in the required number of passes. To control the trajectory, a target 
parameter that specifies the correct trajectory must be known or cal- 
culable. For the aerobraking trajectories, the apogee decay is practically 
invariant with respect to W/CgA and/or density for a fixed entry pass 
number. As a result, this parameter is chosen as the target parameter. 

In addition to apogee altitude itself* the rate of apogee decay must be 
controlled. This decay rate is dependent entirely (for spherical earth) 
on the amount of aerodynamic drag per pass. Since the ballistic co- 
efficient (W/CgA) function is invariant for a given Tug configuration, 
the only control over drag is the depth to which the vehicle penetrates 
the atmosphere. This depth is controlled by varying the vacuum perigee 
altitude at atmosphere entry. For each atmospheric model there exists 
a unique vacuum perigee pass history that will have both the proper apogee 
altitude and apogee decay rate. Currently* the necessary perigee pass 
history is not calculable In advance* Thus, a scheme for controlling 
apogee altitude will also consist of a search for the correct perigee. 

Adjustments in the tra jectory may be made by applying an impulse either 
near atmosphere exit or at the uncorrected apogee. The first method 
(impulse at exit) forces the ascent trajectory through the target apogee. 
Thus, the increase in the drag losses of the previous pass would be taken 
out by use of a propulsive 4V. The second method (impulse at apogee) 
increases the perigee of the descent trajectory of the next pass so that 
less drag would be encountered on that pass* This causes the apogee to 
decay less, and after a number of passes are corrected in this manner, 
the apogee decay approaches the target apogee. Thus, the increase in 


4-91 


D5-17142 




| m j : ; I, i - % i , % c ■jail Jutt ift i B a l!i ! U BXo ca: SES 85 S3 jii£ t£,i & 5 Hfltil a 






0 L/0 » 0 

e orbit inclination * o 

6 ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT KITH TIME 
© INITIAU APOGEE ALT » 117*440*496 FT_ + , 
© ENTRY ALT INTO ATMOSPHERE - 600JC FT~i 
o FINAL APOGEE AIT * Z70 H*H* 

* * SUMMER, HIGH SOLAR ACTIVITY, 

HIGHER DENSITY ATMOSPHERE 
o - WINTER, LOW SOLAR ACTIVITY 
LESS DENSE ATMOSPHERE 

^ ^ t j rcjtrmi a 





HPSiyHSaSff 



DISPERSED ATMOSPHERE 
HOKIHAL ATHOSPHERE 
0952 STAIIOARD) 



mmm 


Hw* v * ■ i v« * * I i rr—‘i Ifc-H t+*Y ***3 l 

y* y T T pp l.J 

&Sj gsj|§ 












— : PASSES TO LOHEJt Si 
601 APOGEE TO 270 H.H.fiT 



i ^Tt raSu 1 ■ W fwrf tffi sn B 

a 1. 1 t ■ i * it 1 h HriBi iiira r» mi jjaoggnm ■ 


»mii9ii 

.aiaMaiiii 



PASSES «= 


mmmmmm 


UNCORRECTEO INITIAL 
PERIGEE EHTRY TIME 
RANGE POR A NOMINAL 
THIRTY PASS TRAJECTORY 


V :f£Q 

g=r.tS ap^iaBS^3SS9 



r- -JT1-K -r-a-iw-M rxrM I Till 

EtLiWi mm: iifi ipK< 

l|raffl| 

iM&mm m 


ISIS 


TIME TO LOWEST AP OGEE .TO 270 jKM^^ JPAYS.r f 


mm\ 


m 


FIGURE 4 *3. 4*0-3 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT 
TIME TO LOWER APOGEE TO 270 NM 



■uliiiHHH 


SSI 


in 











( 


D5-17142 


4. 3.4.1 (Continued) 

draq losses of the previous pass are taken out by having less drag on 
succeeding passes. 

1 . Impulse At Exit 

The propulsive AV for the impulse at exit is applied at atmosphere exit 
(600,000 feet) and co-1 inear with the exit velocity. The AV magnitude 
is determined by computing the velocity required at that point to cause 
the vehicle to be on an ascent trajectory with the target apogee. The 
resultant trajectory also has a higher perigee. This perigee value is 
not necessarily the perigee altitude required for the dispersed density, 
but the perigee does change in the correct direction. 

An example of this method is presented in Figures 4. 3. 4. 1-1, -2, and 
-3 where the trajectory correction AV, perigee altitude, and apogee 
altitude are shown for the basic Tug, nominal 30 pass trajectory. (The 
total AV shown in Figure 4.3.4. 1-1 is the accumulated trajectory 
correction AV and does not include correcting the final pass to 270 
NM). Figure 4.3.4. 1-2 shows the perigee altitude required for decay to 
a 270 NM orbit in the presence of the dispersed atmosphere (the solid 
line). In the real case, this periqee history is not known in advance. 

Thus, an impulse is applied on each pass to correct the apogee (Figure 

4.3.4. 1-3) until the resultant periqees (circled points) converqe to the 
desired value. 

2. Impulse At Apoqee 

An example of impulse at apoqee is illustrated in Figures 4.3.4. 1-4, 

-5 and -6. At the end of the first pass the apogee is too low. This 
means that the oeriqee must be raised to reduce the rate of apogee decay. 

The amount the perigee must be raised is not known in advance. Thus, 

the perigee must be incremented in upward steps (circled points on 
Figure 4.3.4. 1-5) by applying an impulse at apogee until the target 
apogee has been approached (at Pass #7 on Figure 4. 3. 4. 1-6). At this 
point, enough information has been gained to indicate the periqee (solid 
line, Figure 4. 3.4. 1-5) required for this atmosphere model. By averaging 
the perigee magnitudes over the passes now completed, a periqee quess may 
be obtained that lies very close to the desired periqee. The next pass 
then uses this averaged value of perigee. Since the perigee and apoqee 
values on this pass are not exactly those desired, the next pass will be 
in error, but not nearly so much as Initially. From this point, the 
perigee is again incremented until the target apoqee curve is crossed. 

The perigee is again set to the average value of the series of steps 
made since the last average was determined. This procedure is repeated 
up until two passes from the nominal final pass number. The cutoff is 
made at this point because the perigee curve becomes very nonlinear and 
an average value of periaee would be v°ry much in error. (The total AV 
shown in Figure 4.3.4. 1-4 is the trajectory correction AV and does not 
include correcting the final pass to 270 NM). 




4-93 



ACCUMULATED TRAJECTORY CORRECTION £V % FT/SEC 



05-17142 

























VACUUM PERIGEE ALTITUDE •x. 1000 FT 


IOX IOTO ThC CtNTlMETER 4o 15 
MwMtL A OklK CO. 



31^ 


PERIGEE REQUIRED FOR 30 P/ 


STEPWISE CORRECTED PERIGEC 


UNCORRECTED PERIGEE 


1+11. 

PASSES | -~ 


BASIC TUG 

IMPULSE AT ATMOSPHERE EXIT 
SUMMER, HIGH SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE (+ DENSITY) 
ROTATING SPHERICAL ATMOSPHERE THAT 
IS INVARIANT WITH TIME 
NOMINAL THIRTY PASS TRAJECTORY 
" i' ! ' ! I I .• It! .1 • i '■ ( • ! 

'Tt 


- 4 - - 4 - 


| — F INAL PASS OF 

' ' UNCORRECTED 
TRAJECTORY 



1 • ‘ 

[ -*i * 1 

. 1 ' 

t l . .1 

| 

[ ; . 


! 


[ 1 


0 - - t - ! 

PASS NUMBER 


ALTITUDE AT ATMOSPHERE EXIT 


D5- 17142 






















D5-17142 


BASIC TUG 

IMPULSE AT ATMOSPHERE EXIT 
SUMMER, HIGH SOLAR ACTIVITY, 
HIGHER DENSITY 1962 ATMOSPHERE 
(+ DENSITY) 

ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
NOMINAL THIRTY PASS TRAJECTORY 


TARGET APOGEE 

APOGEE OF CORRECTED TRAJ. 


UNCORRECTED 

TRAJECTORY 


FIGURE 4.3. 4. 1-3 




10 PASS NUMBER • ^ 1 

I t i » .1 1 >n »i. 1“ »*•■•• *» 


SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS 
ORBIT - APOGEE A1TIJUDE AT END OF PASS 


m 


• 









IGUkt 4.3.4. 1-4 


-4- 2 0 -!-*-• r — j — 

PASS NUMBER 

SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS 
CORRECTION AV FOR IMPULSE AT APOGEE 


• BASIC TUG 

• IMPULSE AT APOGEE 

• SUMMER, HIGH SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE (+ DENSITY) 

• ROTATING SPHERICAL ATMOSPHERE THAT 
IS INVARIANT WITH TIME 

• AV DOES NOT INCLUDE FINAL PASS 
CORRECTION 

• NOMINAL THIRTY PASS TRAJECTORY 














LRK.ri ALTITUDE 


iCIlk 4b 



BASIC TUG 

-.9 IMPULSE AT APOGEE 

:« SUMMER, HIGH SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE (+ DENSITY) 

• ROTATING SPHERICAL ATMOSPHERE THAT 
IS INVARIANT WITH TIME 
'• NOMINAL THIRTY PASS TRAJECTORY 

| j 

*THIS IS ThF INCREMENTED PERIGEE AT THE END 
OF THE PASS AFTER THE IMPULSE IS APPLIED. 








CC\TiMLT£H 


D5- 1 7142 


j 


lJ 

l : 

i 


'l 


"‘1 



BASIC TUG 
IMPULSE AT APOGEE 
SUMMER, HIGH SOLAR ACTIVITY, 
HIGHER DENSITY 1962 ATMOSPHERE 
(+ DENSITY) 

ROTATING SPHERICAL 
ATMOSPHERE THAT IS INVARIANT 
WITH TIME 

NOMINAL THIRTY PASS TRAJECTORY 
; I | III 

TARGET APOGEE 

» — APOGEE OF CORRECTED TRAJ. 





hr 

*i 

' 

\ - 


FINAL PASS OF 


UNCORRECTED 

TRAJECTORY 


I I ; I I | I, I PASS NUMBER | j | I I 1 I I | j 

SPACE TUG ALROBRAKING RETURN FROM SYNCHRONOUS ORClf 
- APOGEE ALTITUDE AT END OF PASS 



















f 


D5- 17142 


4.3.4. 1 (Continued) 

The trajectory correction JV's for impulse at exit run hiqher than 300 
ft/sec for both the basic Tug and large flare configurations. Resulting 
final pass errors in apogee altitude are within 20UK ft for the basic 
Tug and within 1M ft for the large flare Tua. The trajectories have no 
error in the number of passes. 

The trajectory correction dVs for impulse at apoqee are less than 
20 ft/sec, but the errors in the final apoqee altitude and pass number 
are more significant. This is illustrated in Figures 4. 3. 4. 1-7 and -8 
where the errors in apoqee altitude and pass number are given for basic 
Tug and larae flare respectively. 

The AV* s mentioned in the above two paragraphs are the 4 V's for tra- 
jectory correction before reaching the final pass. They do not include 
the AV ' s to correct the final apoqee altitude error. The total increase 
in the 4V requirements due to atmospheric uncertainties are qiven in 
Figures 4.3.4.1-P and -10. These AV's are the total increase in the ^Vs 
(trajectory and final pass correction) required to end up at 270 over 
that required for the unperturbed trajectories. Either of the two tar- 
geting schemes can control Tug aerobraking trajectories in the presence 
of large, unpredictable atmospheric density dispersions. The impulse 
at exit scheme will require 300 to 800 ft/sec correction AV • Improve- 
ments In this scheme to effect better convergence on final apoqee altitude 
will not reduce the 4V requirements significantly because of the inherently 
large impulses necessary at atmospheric exit to make minor trajectory 
changes. Except for the 10 pass trajectory, requirements of the 
impulse at apogee scheme are generally comparable to or less than impulse 
at exit requirements. Impulse at apogee scheme AV requirements are 
mostly 4V necessary to correct the final apogee altitude error, hence 
improvements in this scheme to obtain better convergence on the desired 
final apogee altitude will significantly reduce these 4V requirements. 

Thus the 400 fps 4V reserved for trajectory correction in the orioinal 
mission 4V requirement, will be sufficient assuminq further development 
of impulse at apogee targeting. 

The maximum dynamic pressure In the presence of the dispersed atmosphere 
is compared in Flqur^s 4.3.4.1-11 through -16 for the impul'^ at exit 
target scheme and In Flqures 4.3.4.1-17 throuqh 4.3.4.1-22 for the impulse 
at apoqee target sc hem?. 

The loads (Section 4.1) and thermal (Section 4.5) analysis was performed 
on the first pass of the 10, 30 and 60 pass dispersed atmosphere trajectories 
of the basic and large flare tugs (the first pass events are identical for 
both targeting schemes). A controls analysis (Section 4.4) was performed 
using the trajectories of the impulse at apoqee targeting scheme. 


4-100 




mr 


D5 - 1 7142 


■?p" r n ""p -y r ~y"j 11 — "ny 


; 1 | 

-4‘— -.4 “—I— - ~u 

4 i ! !• 


~t — i~ - 


• BASIC TUG 

• ROTATING SPHERICAL ATMOSPHERE ... 
THAT IS INVARIANT WITH TIME 

-ri-i- -+i 4 ■ W'Hf ; 

IMPULSE AT A PO G_EJL 

Q 5 PASS NOMINAL ? " j~ ~ T 

O 10 PASS NOMINAL 

□ 30 PASS NOMINAL 

V 60 PASS NOMINAL i ! r ' 


1 o- 

ilitli: 


m 




■ i • 





f — |— ' '~t~ t*— ‘ - f* 1 ' * -j- — 

| ! i 1 ! i ; j. i 

: i I LI t i i 1 1-1 1 1 1 ... 

N N - NOMINAL NUMBER OF PASSES TO 270 N , M . 

FOR STANDARD ATMOSPHERE 
N A - ACTUAL NUMBER OF PASSES FOR 
DISPERSED ATMOSPHERE 


SUMMER, HIGH SOLAR 
ACTIVITY HIGHER 
DENSITY 1962 ATMOSPHERE 

WINTER, LOW SOLAR 
ACTIVITY LESS DENSE 
1962 ATMOSPHERE 


t-4 u. 


•! i 



■: o H- 


.1 ...14 


— i — i. 



i .— l i- i- i . -i -4-4-4 - 

RANGE OF APOGEE ERRORS, 
IMPULSE AT EXIT 

i“ IT! TTTT7.TTT 


... 

i-r- 1 


i ! 


. , . . . . ■ - - • . - v ■ 


i I- - 1 4 Ui 4-.4~.l~L4.. FINAL PASS ERROR (N A - N N ) 

FIGURE 4. 3.4. 1-7 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT 
- ■ : - 1 .1; '- FINAL PASS AND APOGEE ERROR TOR IMPULSE AT APOGEE 


1 . 

• ! •• 'i 


4-101 


142 ’ 


5 PASS NO 
10 PASS NO 
30 PASS NO 


NOMINAL NUMBER OF PASSES TO 270 N.M 
FOR STANDARD ATMOSPHERE 
ACTUAL NUMBER OF PASSES FOR 
DISPERSED ATMOSPHERE 



! 

. | 



4; 

i 



SUMMER, HIGH SOLAR 
ACTIVITY HIGHER 
DENSITY 1962 ATMOSPHERE 

WINTER, LOW SOLAR 
ACTIVITY LESS 
DENSE 1962 ATMOSPHERE 


: oi4”-- — r-'-rrf- 


rrvr 


•4 ’-r 4 

L _[ ' ' | ' 'T J 

r ! ' • . " r7 T 1 I* 


I- 


. 4:4 i 

I III. Ill 


•I I 


>3 


RANGE OF APOGEE ERRORS, 
IMPULSE AT EXIT 


I L 


I . I 

, 


•2 : 3 


i.«*i L+LJ — 4 — U- final pass error (n a - n n ) | 

FIGURE 4.3. 4. 1-8. SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT 
L I ' 11- - FINAL PASS AND APOGEE ERROR FOR IMPULSE AT APOGEE 


4-102 ’ 












K +XZ logarithmic 46 5493 

£Z 2 CTC.ti * »' C- V<- .ON> »*ci ■* «• » 


» IfStB < O 


J I 


— 7— ~ ~ — — 

— 7 . 2000 -1 — 

~P - ! 


IMPULSE AT EXIT 
ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
+ DENSITY - SUMMER, HIGH 
SOLAR ACTIVITY HIGHER 
DENSITY 1962 ATMOSPHERE 
- DENSITY - WINTER, LOW 
SOLAR ACTIVITY, LESS 
DENSE 1962 ATMOSPHERE 










, frci-fr 




















1C ( 


? Hi. C..NIII<4C1 LM 


I -r : 
T--“1 



i .^r t 




a: — l 20 

z> 

co r: i 

m ; 

uj* — ! — T“ 

cr 

to. - r i 


! : 

ill 



~ **• • 4- 

i r 





m 

r* r- 


-di 


BASIC TUG 
IMPULSE AT EXIT 

NOMINAL - 1962 STANDARD ATMOSPHERE 
+ DENSITY - SUMMER, HIGH SOLAR ACTIVITY 
HIGHER DENSITY 1962 ATMOSPHERE 
- DENSITY - WINTER, LOW SOLAR ACTIVITY 
LESS DENSE ATMOSPHERE 
ROTATING SPHERICAL ATMOSPHERE THAT IS 
INVARIANT WITH TIME 
NOMINAL TEN PASS TRAJECTORY 


-U-- 


~t* t- 

- 14- 14 


rq- 

,.Ld 


r-rzrm: 


45, O 
1 J — 



+~ - i 


HT 




■ 


j 



: i 


i 

I I 


j : t 


i s -\- 


[ ! 

J_i. i 


+ DENSITY 
NOMINAL 
- DENSITY 


-j — I- 


! i !r : - J T 

, , rT : 


1*1. 


t h 


FIGURE 4.3.4. 


SPACE TUG 
PRESSURE 


AEROBRAi 
FOR BAS 1 1 


r i i i i i « 

3 4 • i|r 5 - r ; 6 7 b • • • 9 • ‘ 

1 ...j. PASS NUMBER :i ; \ j L— L 

NG RETURN FROM SYNCHRONOUS ORBIT - MAXIMUM DYNAMIC -! 
TUG IMPULSE AT EXIT — 



BASIC TUG 
IMPULSE AT EXIT 

NOMINAL - 1962 STANDARD ATMOSPHERE 
+ DENSITY - SUMMER, HIGH SOLAR 
ACTIVITY, HIGHER DENSITY 
1962 ATMOSPHERE 
- DENSITY - WINTER, LOW SOLAR 
ACTIVITY, LESS DENSE 
1962 ATMOSPHERE 
ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
NOMINAL THIRTY PASS TRAJECTORY 



















PRESSUR 


BASIC TUG 
IMPULSE AT EXIT 

NOMINAL - 1962 STANDARD ATMOSPHERE 
+ DENSITY - SUMMER, HIGH SOLAR 
ACTIVITY, HIGHER DENSITY 1962 
ATMOSPHERE 

- DENSITY - WINTER, LOW SOLAR 
ACTIVITY, LESS DENSE 1962 
ATMOSPHERE 

ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
NOMINAL SIXTY PASS TRAJECTORY 




1 : 

1 * ‘ i _i 1_ 

, 

— 


i ‘ : i * i 1 1 


+ DENSITY 
NOMINAL 
- DENSITY 


PASS NUMBER 


IGURE 4.3.4.1-13 SPACE TUG AEROBRAKING RETURN FROM SYNCHROS US ORBIT 
DYNAMIC PRESSURE FOR BASIC TUG IMPULSE AT EXIT 


MAXIMUM 













I 

O 

CD 




1 ... - 

1 

1 

'% | . J 

- 4 r 

i 


; ; > ' 


L: ! 

rrr ! . . 

. J 





L_J !i 


• LARGE FLARE 

• IMPULSE AT EXIT 

• NOMINAL - 1962 STANDARD 
ATMOSPHERE 

• + DENSITY - SUMMER, HIGH 
SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE 

• - DENSITY - WINTER, LOW SOLAR 
ACTIVITY, LESS DENSE 1962 
ATMOSPHERE 

ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
NOMINAL TEN PASS TRAJECTORY 


t/i . 


s - 2 - 

1 




— — 

t~ ~ 

-_ 4 _ 



N- 

“““ V 

sT 

X - 

<■ 

- 1 

/ 




$ 

; ! 


N 

NN— + DENSITY 

r — x ‘ — 

L 


/ 


J ; 

i I : 

lll: . 

NOMINAL 



/ 

i ■ i . i . 

• ! 


• F 

- DENSITY 

[ 


VTT 

i 







1 

A 

r 

i 

' • ’ ! • 

'.'■tr 



. ; r r_J 

nil - 1 

* ' i / 

U ’■ i : 


I Ll-l . 


FT ! . 

i 

l : ... : -i.J 

I j 

/ 

i | 


j ' i ■[ j 


! ■ 

- 1 


! : ? . j 

i- * . 

i j 

-j r — r— 1 

I 

*1 “ r "l 1 

. : ! 




_ 

i - ■ 



i ! i v 


4 1. 5 r ! 

..PASS NUMBER! 


FIGURE 4.3.4.1-14 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - 

MAXIMUM DYNAMIC PRESSURE FOR LARGE FLARE TUG IMPULSE AT EXIT : — 



MAXIMUM DYNAMIC PRESSURE 



















ESSURE - PSF 


i 




LARGE FLARE 

IMPULSE AT EXIT P^: 

NOMINAL - 1962 STANDARD ATMOSPHERE 
+ DENSITY - SUMMER, HIGH SOLAR fr*^7r 
ACTIVITY HIGHER DENSITY 1 962 
ATMOSPHERE fe 

- DENSITY - WINTER, LOW SOLAR L 

ACTIVITY, LESS DENSE 1 962 pu 

ATMOSPHERE f : v| 

ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME (r~;- 

NOMINAL SIXTY PASS TRAJECTORY 




j : *m. :! T:: L : 

i i: i ! - 




:i ! 




0 44f-4 1 0 4— 4-4 201444 30 444 4044-4 504-4-j 60 4 44::::[: 4:4: •; i :.: : 4fe4: 
j- • » - =: i : -i -^1 - 1- ! 1 ■ r P /kc NUMBER :r 4 

■FIGURE 4.3.4.1-16: SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - MAXIMUM 
j:4f DYNAMIC PRESSURE FOR LARGE FLARE TUG IMPULSE AT EXIT 


.J: | I -: 1 

44 30«444r 40-4-4-4 50 4 


4 60 44 













PRESSU 


BASIC TUG • ~ DENSITY - WINTER, LOW 

IMPULSE AT APOGEE SOLAR ACTIVITY, LESS 

NOMINAL - 1962 STANDARD t DENSt ATMOSPHERE 
ATMOSPHERE # • NOMINAL TEN PASS TRAJECTORY 

/density . SUMMER, HIGFFF^^- : 

SOLAR ACTIVITY, HIGHER : . 1 i { ! •= 1 j i ' 

DENSITY 1 962 ATMOSPHERE !j 1 7 I ! M 






?r*r] 



s * 

T“TT f 


J 


+ DENSITY 
NOMINAL 
- DENSITY 


-L..:. 


PASS NUMBER 


FIGURE 4.3.4.1-17. SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT 

OVNAMTC PRFSSIIRF Ff)P RASTC TUG IMPULSE AT APOGZE 


- MAXIMUM J 














BASIC TUG 
IMPULSE AT APOGEE 
NOMINAL - 1962 STANDARD 
ATMOSPHERE 

+ DENSITY - SUMMER, HIGH 
SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE 
- DENSITY - WINTER, LOW 
SOLAR ACTIVITY, LESS DENSE 
1962 ATMOSPHERE 
NOMINAL THIRTY PASS TRAJ. 


DENSITY 

NOMINAL 

DENSITY 


FIGURE 4.3.4.1-18 


- f - — | — r-T-tr— t~-Tr r: h- 

: r — 1 1 — U pass number j 

SPACE TUG AE ROB RAKI NG RETURN FROM SYNCHRONOUS ORBIT 
DYNAMIC PRESSURE FOR BASIC TUG IMPULSE AT APOGEE 


r~ 


- MAXIMUM 





Cf«T MtrtH 4b 1 


BASIC TUG 
IMPULSE AT APOGEE 
NOMINAL - 1962 STANDARD 
ATMOSPHERE 

+ DENSITY - SUMMER, HIG 
SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE 
- DENSITY - WINTER, LOW 
SOLAR ACTIVITY, LESS 
DENSE 1962 ATMOSPHERE 
ROTATING SPHERICAL 
ATMOSPHERE THAT IS 
INVARIANT WITH TIME 
NOMINAL SIXTY PASS 



+ DENSITY 
NOMINAL 
- DENSITY 







4-114 


LARGE FLARE 
IMPULSE AT APOGEE 
NOMINAL - 1962 STANOARD 
ATMOSPHERE 

+ DENSITY - SUMMER, HIGH SOLAR 

ACTIVITY, HIGHER DENSITY 

1962 ATMOSPHERE 

- DENSITY - WINTER, LOW SOLAR 

ACTIVITY, LESS DENSE 1962 

ATMOSPHERE 

NOMINAL TEN PASS 




h H- 



} . : | : | PASS NUMBER! \ ; « j ! 

FIGURE 4.3.4.1-20 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - 

DYNAMIC PRESSURE FOR LARGE FIARF TMPIli SF fit fionrrr 


MAXIMUM 


I 








LARGE FLARE 
IMPULSE AT APOGEE 
NOMINAL - 1962 STANDARD 
ATMOSPHERE 

+ DENSITY - SUMMER, HIGH 
SOLAR ACTIVITY, HIGHER 
DENSITY 1962 ATMOSPHERE 
- DENSITY - WINTER, LOW 
SOLAR ACTIVITY, LESS 
DENSE 1962 ATMOSPHERE 
ROTATING SPHERICAL ATMOSPHERE 
THAT IS INVARIANT WITH TIME 
NOMINAL THIRTY PASS 


- DENSITY 
NOMINAL 
+ DENSITY 


I : ; „ PASS NUMBER J 

FIGURE 4.3.4.1-21 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT 

DYNAMIC PRESSURE FOR LARGE FLARE IMPULSE AT APOGEE 


- MAXIMUM 


D5-17142 








MAXIMUM DYNAMIC PRESSURE 


LARGE FLARE 
IMPULSE AT APOGEE 
NOMINAL - 1962 STANDARD 
ATMOSPHERE 

+ DENSITY - SUMMER, HIG 
SOLAR ACTIVITY, HIGHER 
DENSITY ATMOSPHERE 
- DENSITY - WINTER, LOW 
SOLAR ACTIVITY, LESS 
DENSE 1962 ATMOSPHERE 
ROTATING SPHERICAL 
ATMOSPHERE THAT IS 
INVARIANT WITH TIME 
NOMINAL SIXTY PASS 





+ DENSITY 
NOMINAL 
- DENSITY 


‘FIGURE 4.3.4.1-22 


SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - MAXIMUM 
DYNAMIC PRESSURE FOR LARGE FLARE TUG IMPULSE AT APOGEE 









D5-17142 


4. 3. 4. 2 Dispersed Atmosphere Effects on Inertial Velocity 

Prior Figure 4.3.4. 1-3 illustrated the mission duration ranqe that could 
result from atmospheric dispersions. The high density atmosphere can 
significantly shorten tne mission duration. Similarly, the low density 
atmosphere can siqnificantly increase the mission duration. The mission 
duration is a direct function of the velocity (kinetic energy) decrease on 
each pass. This subsection discusses the impact of atmospheric perturba- 
tions on the first pass inertial velocities. The first pass inertial 
velocity is independent of the targeting techniques discussed in Subsection 
4.3.4. 1 above and is a determining factor for the maximum equilibrium 
temperatures experienced (Section 4.5). 

As discussed in Subsection 4.3.3 above, the orbit's maximum inertial 
velocity is attained at some time prior to perigee because of the atmos- 
pheric braking. Figure 4. 3. 4. 2-1 shows the trends of the elapsed time 
between the orbit's maximum Inertial velocity point and the perigee for 
the first pass of the mission. The greater velocity (kinetic energy) 
loss per pass for the shorter duration missions is reflected in the rela- 
tively long elapsed time periods for the 5 and 10 pass missions. The 
longer duration missions, because of the higher perigee altitudes and less 
velocity loss per pass, have shorter elapsed times. The elapsed time is 
significantly lower for the low density atmosphere because of the smaller 
decelerations experienced. 

Figure 4. 3. 4. 2-2 shows the sensitivity of the basic (no flare) configura- 
tion's first pass inertial velocities to atmospheric effects. The first 
pass maximum inertial velocity is nearly insensitive to mission duration 
and/or atmosphere. This relatively constant velocity is the result of 
the nearly identical elliptical free space trajectories flown from geo- 
synchronous orbit to the first atmospheric entry. The first pass periaee 
inertial velocity is sensitive to the atmospheric state. For a 5 pass 
mission, the perturbed atmosphere changes the first pass periqee velocity 
by 200 ft/sec from that of the standard atmosphere. This first pass 
perigee velocity change is less for lonqer duration missions. The impact 
of these relatively small perigee velocity changes is reflected in siq- 
nificantly longer or shorter mission durations as seen in prior Fiaure 
4. 3. 4. 1-3. 


Figure 4. 3. 4. 2-3 shows similar data for the 60° flare configuration. 

This configuration has the same trends noted in Figure 4. 3. 4. 2-2 for the 
no flare configuration. 

The difference between the first pass maximum and periqee velocities shown 
in Figures 4. 3. 4. 2-2 and -3 is only an indicator of the velocity loss 
during the pass. There will be some losses prior to attaininq the maximum' 
inertial velocity. Between periqee and exit at 600,000 feet altitude, 
another major velocity loss occurs. Also, in a free space environment, 
the periqee Inertial velocity is the maximum inertial velocity alono the 
trajectory and would be somewhat qreater than that shown as the maximum 
in either fiqure. Th^s additional factor represents another velocity loss 
not shown in the figures. 


4-117 


TIME (SECONDS) 



NUMBER OF PASSES 


FIGURE 4. 3. 4. 2-1. ELAPSED TIME TREND FROM MAXIMUM INERTIAL VELOCITY TO PERIGEE 

(FIRST PASS) 


D5-17142 






INERTIAL VELOCITY (THOUSANDS OF FEET/SECOND) 


HIGH DENSITY ATMOSPHERE 


— O— STANDARD ATMOSPHERE 
LOW DENSITY ATMOSPHERE 



FIGURE 4. 3.4. 2-2. SENSITIVITY OF FIRST PASS INERTIAL VELOCITY TO ATMOSPHERE 

BASIC TUG (NO FLARE) CONFIGURATION 


D5-17142 


INERTIAL VELOCITY (THOUSANDS OF FEET/SECOND) 


HIGH DENSITY ATMOSPHERE 


— O — STANDARD ATMOSPHERE 
LOW DENSITY ATMOSPHERE 



FIGURE 4. 3.4. 2-3. SENSITIVITY OF FIRST PASS INERTIAL VELOCITY TO ATMOSPHERE 

(60° FLARE CONFIGURATION) 


05-17142 


4.3.5 


Lunar, Solar, and Earth Harmonics Perturbations 


Because of the larqe amount of computer time necessary to simulate the 
effects on the trajectorie: of earth gravity harmonics and Lunar and 
Solar Perturbatiors, these effects are not included in the trajectory 
data presented in previous sections. However, to assess the impact of 
these perturbations on the study results, some trajectories were simulated 
including these perturbations The most pronounced effect of the pertur- 
bation is a reduction in the semi-major axis of the orbit so that by the 
time the Tuq reaches the entry altitude of 600,000 feet the periqee 
altitude is less than what it would have been in the absence of the per- 
turbinq accelerations. This is illustrated in Fiqure 4. 3. 5.0-1 where the 
behavior of periqee altitude with true anomaly in the presence of gravity 
perturbations is compared with the non-perturbed behavior. In the 
absence of gravity perturbations the periqee altitude is constant until 
the atmosphere is encountered at a true anomaly of about -15°. Draq 
perturbations then decrease periqee altitude slightly and there is no 
further effect on periqee altitude for the remainder of the orbit. Gravity 
perturbations cause periqee altitude to decrease on the descending part of 
the orbit until at periqee the periqee altitude is almost 10,000 feet 
lower than the non-perturbed perigee. The cumulative result of this 
effect on the first and succeeding passes is that the nominal 10 pass 
trajectory re-enters and impacts the earth on the sixth pass. This means 
that the initial perigee altitude (as calculated at apoqee and ignoring 
the non-spherical qravity perturbations) required to produce apogee decay 
to 270 NM in a given number of passes must be greater than that shown in 
the previous data (Figure 4. 3. 3. 0-1). This effect is shown in Figure 
4. 3. 5.0-2 for five, ten, and thirty pass basic Tuq trajectories. 

Obviously then, gravit/ perturbations must be considered in the targetinq 
scheme analysis. It i; felt, however, that the principal of the tarqeting 
schemes discussed in Section 4.3.4 can still be used in the presence of 
gravity perturbations with the major effect being modification of the 
targeting parameters, apoqee altitude and decay rate, to include the 
perturbation effects. 

Impact of qravity perturbations on heatinq and loads environment 
encountered during the time in the atmosphere can be assessed by comparinq 
the altitude time histories of the perturbed ana spherical earth trajec- 
tories. This comparison is made in Fiaure 4. 3. 5. 0-3 for the portion of 
the trajectory below 600,000 feet on the first pass of a 10 pass trajectory. 
Similarity of the curves indicates that heatina, loads, and control data 
presented in the followina sections will not be chanqed siqnif icantly by 
perturbation effects on the trajectory. 

Other perturbation effects on the orbital parameters, while not significant 
from a performance standpoint, will be operationally important. For 
example, the line of nodes will regress about 8° for the thirty pass 
trajectory while the argument of perigee advances 9° and the orbit plane 
inclination decreases half a deqree. These perturbations must therefore 
be considered for proper timing and phasing of the mission. 



000 FT 



ORBIT INCLINATION = 28° 

1962 STANDARD ATMOSPHERE 
ENTRY ALTITUDE - 600,000 FT 
J 2 = 1082.7 X 10-6 
J 3 = -2.56 X 10-° 

J 4 = -1.58 X 10 - 6 
DATE: 1 JULY, 1976 


TRUE ANOMALY a DEG 


ING RETURN FROM SYNCHRONOUS ORBIT 


FIGURE 4. 3. 5. 0-1 


SPACE TUG AEROBRAK 
VACUUM PERIGEE HISTORY OF THE FIRST PASS OF A TEN PASS TRAJECTORY 
























TO 

koo« 


% DAYS 


4 50.0 


VCLES X 70 DIVISIONS 

Keup-rrc a esser c 


PASS NUMBER 


SPHERICAL EARTH 


rn PERTURBED 


FIGURE 4. 3. 5.0-2 


ZlL TIME TO LOWE 

--"T. : I 1 ... I . 

SPACE TUG AEROBRAKING RETURN FROM 
TIME TO LOWER APOGEE TO 270 N.M. 


LOWER APOGEE TO 270 N.l 

1 L_J L_. I A 1 L _i 

SYNCHRONOUS ORBIT 


ORBIT INCLINATION = 28° 

1962 STANDARD ATMOSPHERE 
ENTRY ALTITUDE = 600,000 FT 
J 2 = 1082.7 X 10-6 
J 3 = -2.56 X 10 ‘ 6 
J 4 = -1.58 X 10 * 6 
DATE: 1 JULY, 1976 


05-17142 













































ORBIT INCLINATION = 28^ 
1962 STANDARD ATMOSPHERE 
ENTRY ALTITUDE = 600,000 
J 2 = 1082.7 X 10“ b 
J 3 = -2.56 X 10-6 
J 4 = -1.58 X lO- 6 
DATE: 1 JULY, 1976 


SPHERICAL 

EARTH 


TIME FROM ENTRY AT 600 KFT - SECONDS 


FIGURE 4. 3. 5. 0-3 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - 

ALTITUDE TIME HISTORY OF THE FIRST PASS OF A TEN PASS TRAJECTORY 


D5-17142 









D5- 17142 


\ 


4.4 CJCNTFDL ANALYSIS 

Attitude control in pitch, yaw, and roll is accomplished with the PCS 
system both outside and within the atmosphere. The PCS system also performs 
trajectory corrections, vehicle maneuvers, and docking. This analysis gives 
propellant consumption needed for the PCS system to maintain attitude con- 
trol. Trade studies cm critical parameters are also given. 

4.4.1 Control Methods 


The total PCS propellant consumption required for attitude control is calculated 
by sunning the consumption due to each mode of operation. The modes of 
operation are: 


1. Attitude control by limit-cycling between system firing limits 
(deadband) in pitdi, yaw, and roll. 

2. Aero- moment control within the atmosphere for the unstable basic 
Tug. 

3. Aero-moment control within the atmosphere for the stable flared 
tugs. 

4. Directional control, keeping the tug aligned along the velocity 
vector throughout the period of orbit change. 

The fuel consumption calculations are as follows: 


1. Consumption due to limit-cycling is given by the following equation: 

PROPELLANT WEIGH? = 57. 3T T £N 2 F 2 T 2 

” 4 1~ D Is7~ 


where: 



total time of entry 
radius of vehicle 
nunber of thrusters 
thrust of one engine 
minimum firing time 
vehicle inertia 
specific impulse 
system deadband in degrees 


2. Consumption due to an unstable aero-moment: 


The aero-mcment is calculated by approximating Q and time of aero 
influence as follows: 


4-125 


D5-17142 


4.4.1 (Continued) 


AIKO 

MOMENT 



Aero-nonent at 1° angle of attack: AM = C^ QA (Xqp-Xqq) (1/57.3) 

where: A = vehicle cros sectional, area 

Q = dynamic pressure 
Xq> = center of pressure 
Xq_. = center of gravity 

- coefficient of aerodynamic normal force 

CONTROL MOMENT (CM) = NIF 

2. To determine the bum time necessary to control the tug during 
maximum aero-moment, assume vrorse-case aero influence as shown in 
Figure 4. 4. 1.0-1 


TYPICAL LIMIT-CYCLE OUTSIDE 
ATMOSPHERE 


LIMIT-CYCLE UNOIR HORSE CASE 
INFLUENCE OF AERO-NONENT 



THRUSTER FIRES 


VEHICLE 

CONTROL 


IS UNOER THE INFLUENCE OF TNE 
- AERO MOMENT DIFFERENCE 


THRUSTER SHUTOFF 


Figure 4. 4. 1.0-1 PCS System Limit-Cycle Under Worse Case 
Influence of Aero-Mcment 

From Figure 4. 4. 1.0-1, after thruster burn ends, the only moment actinq on 
vehicle is the aero-moment. To reignite thrusters, 0 must become 
equal the negative of the value at thruster shutoff. 

Patio of thruster bum time (Ter) to cycle-time (Tcy C i e ) • 

T 

BT _ AM _ average aero -moment for al l passes 
T CYCLE W control moment of ECS 

Engine bum time due to aero-mcment = [time under aero-moment 
influence for all passes] [ t bt ] 

T 

CYCLE 


4.4.1 


(Continued) 


05-17142 


Toted propellant = [engine bum time due to aerc-mcment] [NF] 

I SP 


3. Consumption due to a stable aero-mcment : 


As tlie space tug enters the atmosphere, the aero-mcment increases 
causing the tug oscillations to converge and increase in frequency 
as shown in Figure 4. 4. 1.0-2. After passing max q, the Tug oscilla- 
tions diverge. The minimum 0 for the tug oscillations at max q 
depends on the aero-mcment and on the deadband when the tug enters 
the atmosphere, or the initial conditions at start of the damping 
influence (increasing aero-mcment) . F rom Figure 4. 4. 1.0-2, 0 decreases 
to +.25 degree at max q. For a 5 degree deadband, 0 decreases to 
approximately + 1.0 degree. A configuration less stable than the 
60 degree flare tug will give a larger value of 

The amount of the deadband in or outside the atmosphere depends 
on guidance requirements. However, if small oscillations 
are desired at max q a small deadband such as 1.0°, 0.5°, 0.3°, etc., 
could be set by the PCS system prior to entering the atmosphere . 

4. Consumption required for directional control 

The tug must rotate 360° for each earth pass to maintain an align- 
ment with the velocity vector. A conservative propellant consumption 
of 0.3 lbs/pass is used for this operation . Ihis will give a rotation 
rate of 0.5 degrees/seocnd, which is an order of magnitude greater 
than the maximum rate for the 60th pass of a 60 pass trajectory. 

4.4.2 Atmospheric Effects on Controls 

There are two types of attitude control within the atmosphere for the 

different tug configurations: 

1. For the aerodynami cal ly unstable basic tug, the PCS system must 

control the aero-mcment causing greater fuel consumption than for a 
stable tug. The fuel consumption is dependent on the system 
dead) and, since the aero-mcment increases as the angle-of-attack 
increases. The analysis uses a + 1.0° deadband. Effects of using 
a 0.5° deadband are shown in Figure 4. 4. 2. 0-1. 


2. For the aerodynami cal ly stable flared tugs, the aero-mcment will 
cause the tug pitch/yaw attitude deadbands to converge. This 
reduces RCS fuel consumption by up to 550 pounds from the basic 
tug configuration. 



4-128 


0.14 



ATTITUDE RATE 

U) 0.0 

DEG/SEC 

- 0.02 


NOTE: 

RCS SYSTEM DEADBAND =1.0° 


0.12 


0.10-4 


RCS SYSTEM 
FIRING LINES 


0.08-4 


0.06 H 


— r 


TUG ENTERS ATMOSPHERE 


0.04 


0.02 


-0.04 


-0.06 


-0.08 


- 0.10 


- 0.12 


LARGEST VALUE OF $ 
OCCURS AT MAX Q 

1 r 


-0.14 


0.4 


- 0.8 


- 0.6 


- 0.2 0.0 

ATTITUDE {$) - DEGREES 

FIGURE 4.4.I.0-2 PHASE PLANE TRAJECTORY SHOWING TUG OSCILLATIONS FROM TIME TUG ENTERS 

ATMOSPHERE UNTIL IT PASSES MAX 0 


D5-17142 


RCS 

PROP ELI ANT , 
LBS 150 


RcS SYSTEM DEADBAND =1.0' 
WITHIN THE ATMOSPHERE 


RCS SYSTEM DEADBAND = 0.5° 
WITHIN THE ATMOSPHERE 


•NOMINAL SYSTEM CHARACTERISTICS 
■DEADBAND VARIED 


NOTE: 

YAW AXIS CONSUMPTION IS THE SAME AS PITCH 


NUMBER OF EARTH PASSES 


FIGURE 4.4.2.0-I PITCH AXIS PROPELLANT CONSUMPTION DUE TO THE AERO MOMENT FOR 

THE BASIC TUG 



D5-17142 




05-17142 


A 


4.4.3 Astrinnics Effects on Control Requirements 

The RCS system must oontrol the vehicle attitude within a deadband such 
that guidance sensors will operate accurately. Deadbands suitable for the 
operation of the guidanoe sensors are: 

1. Horizon sensor - + 10° pitch/yaw and +2.5 roll 

2. Larklnark tracker - +3.75° pitch, yaw, and roll 

3. Star Tracker - +5° pitch, yaw, and roll 

These, or smaller deadbands, are required only during periods of sensor 
operation. The analysis uses +5° for pitch/yaw outside the atmosphere, 
where the sensors are used. This will give conservative values for the 
total change of orbit period, since a deadband larger than 5° can be used 
when sensors are not operating. 


4.4.4 Control Options 

The RCo propellant consumption due to limit-cycling is highly sensitive 
to ndnimum thruster pulse width and system deadband. An attempt has 
been made to choose values for these parameters which minimize propellant 
consumption and remains within practical design limits. 

The RCS system as given in the Boeing Pxe-Phase A Space Tug Study Report, 

(prior Reference 1.1. 0.0-1) was used as the basis for this study. 


RCS System Characteristics : 

Pitch, Yaw, and Poll Minimum Firing Time 

Pitch and Yaw Thrust 

Pitch and Yaw Thrusters 

Roll Thrust 

Roll Thrusters 

Pitch and Yaw Deadband 

Roll Deadband 
ISP 


30 Milliseconds 
200 Ibs/Engine 
4/Plane 
20 Ibs/Engine 
8 * 

5° CUtside the 
Atmosphere 
2.5° 

400 Sec 


*Only four roll thrusters are used to oontrol roll attitude. 



4-130 .. 



D5 - 1 7 1 42 


4.4.5 Results 

Consumption values for the different modes of RCS operation, except for 
the rotational mode (0.3 lbs/pass), are given in Figures 4. 4. 2. 0-1 and 
4. 4. 5.0-1 through 4. 4. 5. 0-3 for the basic and 60° flare Tugs. For pitch 
and yaw limit-cycle control only one set of curves are given for the 
basic and 60° flare Tugs. This is due to the vehicles having little 
difference in pitch and yaw moments of inertia and consequently little 
difference in consumption. Variations account for different values of 
system deadband and minimum pulse width. Propellant consumption values 
for the different Tug configurations are shown below. (Note that these 
requirements do not include that required for navigational error or 
atmospheric dispersions effects.) 


CONSUMPTION DIFFERENCES AMONG THE TUG CONFIGURATIONS 
FOR A 30 PASS MISSION 


TUG 

CONFIGURATION CONSUMPTION (LBS) 


Basic (no flare) 620.0 
60° Flare 76.0 
45° Flare 80.0 
30° Flare 82.0 


Total consumption for vehicle attitude control for the four configurations 
vs. number of earth passes are given in Figures 4. 4. 5.0-4 and 4. 4. 5. 0-5. 

The effects of varying the atmospheric density based on trajectories 
given in Section 4.3 are shown in Figures 4. 4. 5. 0-6 and 4. 4. 5. 0-7 for 
the basic and 60° flare Tugs. The consumption values for the atmospheric 
density variation account for a delta velocity to correct trajectory 
errors caused by the off-nominal condition. A few trajectories with the 
density variations qave a different number of earth passes than the nominal 
trajectory. This produced a large increase in consumption for the basic 
Tug 10 pass case (Figure 4. 4. 5.0-5) where the trajectory qave 11 passes for 
the change of orbit period. However, in general, atmospheric density 
variations could increase RCS propellant requirements approximately 10%. 

An RCS GO 2 /GH 2 system was assumed in this study. This system Is still in 
development. It is possible that system characteristics will change or 
that another system with lower specific impulse will be chosen, either of 
which will impact RCS propellant weights. 


4-131 



4-132 


1 I I 1 I I I I I l I 

NOTE: 

YAW AXIS CONSUMPTION IS THE SAME AS PITCH 


30 MILLISECOND MINIMUM PULSE 
1 DEGREE DEADBAND 


50 MILLISECOND MINIMUM PULSE 
5 DEGREE DEADBAND 


30 MILLISECOND MINIMUM PULSE 
5 DEGREE DEADBAND 


10 MILLISECOND MINIMUM PULSE 
5 DEGREE DEADBAND 


NUMBER OF EARTH PASSES 


FIGURE 4.4. 5. OH PITCH AXIS PROPELLANT CONSUMPTION DUE TO LIMIT CYCLING FOR THE 

BASIC AND 60° FLARE TUGS 



D5-17142 




4-133 


RCS 

PROPELLANT 

LBS 


52 

48 

44 

40 

36 

32 

28 

24 

20 

16 

12 

8 

4 

0 


NOMINAL SYSTEM CHARACTERISTICS 

DEADBAND VARIATION 










— 

— 1 

■MIN 

IMUM 

PUL 

SE V 

ARIA 

TION 


















' 





5( 

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> 


















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* 

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JLSE 

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M 

BAND 








m ^ _ 




— 

— 

t 



10 


20 


30 40 ,50 60 

NUMBER OF EARTH PASSES 


70 


80 


90 


FIGURE 4.4.5.0-2 ROLL AXIS PROPELLANT CONSUMPTION DUE TO LIMIT CYCLING FOR THE BASIC TUG 


D5 - 1 71 42 





























1 









■ 



Mil 
.5 D 

.LIS 

:gre 

ECON 
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4D8A 

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(-30 Mil 
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W V * ■ 

): 60 70 80 90 


PASSES 

J DUE TO LIMIT CYCLE^FORThE SPACE ‘ 


D5-17142 




600 


RCS 

PROPELLANT 

LBS 


500 


400 


300 


200 


100 


0 




■ 


■ 

n 



- — 



■ 

■ 

■ 

■ 

■ 

■ 

■ 

■ 

B 

-1 




9 

u 






n 

A 

CCIIMC 

>TTfiN< 


rn 

rn 


■ 



m 


■ 

m 




h i iunj 

'ITCH, YAW, AND ROLL ATTITUDE IS CONTROLLED 
HROUGHOUT PERIOD OF ORBIT CHANGE. 
tCS SYSTEM IS DESIGNED WITH: 

) 5° DEADBAND IN PITCH/YAW (1° WITHIN 

ATMOSPHERE); 2.5° IN ROLL 
*) 4-200 LB HIGH PRESSURE THRUSTERS IN 

THE PITCH AND YAW PLANES 
t) 4-20 LB HIGH PRESfURE THRUSTERS IN 
THE ROLL AXIS 

) PITCH, YAW, AND ROLL THRUSTER MINIMUM 

B 

B 



91 



m 

■ 

■ 



■ 


■ 


■ 





■ 





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■ 

■ 

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■ 

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■ 

■ 

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■ 

■ 

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— 

— 

□ 

H 

1 

i 

1 3 

□ 



0 1 0 20 30 40 50 60 70 80 90 


NUMBER OF EARTH PASSES 


RCS PROPELLANT CONSUMED FOR THE BASIC TUG DURING CHANGE OF ORBIT PERIOD 


FIGURE 4.4.5.0-4 


05-17142 



NOTE: 30°, 45°, AND 60° FLARE TUGS ARE AERODYNAMI CALLY STABLE 


I 






III! 

30 3 FLARE TUG 


■45° FLARE TUG 

i i r i 

60° FLARE TUG 


ASSUMPTIONS 

PITCH, YAW, AND ROLL ATTITUDE IS CONTROLLED 
THROUGHOUT PERIOD OF ORBIT CHANGE. 

RCS SYSTEM IS DESIGNED WITH: 

1) 5° DEADBAND IN PITCH/YAW (1° WITHIN 
ATMOSPHERE); 2.5° IN ROLL 

2) 4-200 LB HIGH PRESSURE THRUSTERS IN 
THE PITCH AND YAW PLANES 

3) 4-20 LB HIGH PRESSURE THRUSTERS IN 
THE ROLL AXIS 

4) PITCH, YAW, AND ROLL THRUSTER MINIMUM 
FIRING TIME IS 30 MILLISECONDS 


NUMBER OF EARTH PASSES 

FIGURE 4.4.5.0-5 RCS PROPELLANT CONSUMED FOR TUGS WITH FLARE DURING CHANGE OF ORBIT PERIOD 


D5-17142 



4-137 




8J0-K 


700 


600- 


| RCS 
PROPELLANT 500 
i LBS 


400 


300 


10 


1 1 1 1 1 1 1 1 

NOMINAL CONDITIONS 

ATMOSPHERIC DENSITY INCREASE 




■ 

■ 






NO 1 

i 

‘E: 

THE 5 

PASS 

CAS 

I WAS 

NO! 

i 1 

ANALYZED 

1 

— i 

m. 



■ 

■ 

n 

■ 


■ 



— 











■ 

— , 



■ 








r> — ; 








: 


■ 

■ 

■ 








— ■ 

’ 

— 




1 



■ 

■ 

■ 

■ 

■ 












■ 

'I 

■ 

■ 

■ 

■ 

■ 

■ 



/ 










ASSl 

IMPTI 

)NS 






r 



* 




i 

> PITCH, YAW, AND ROLL ATTITUDE IS CONTROLLED. 

THROUGHOUT PERIOD OF ORBIT CHANGE. 

• RCS SYSTEM IS DESIGNED WITH: 

1) 5° DEADBAND IN PITCH/YAW (1° WITHIN 
ATMOSPHERE); 2.5° IN ROLL 

2) 4-200 LB HIGH PRESSURE THRUSTERS IN 
THE PITCH AND YAW PLANES 

3) 4-20 LB HIGH PRESSURE THRUSTERS IN 
THE ROLL AXIS 

4) PITCH, YAW, AND ROLL THRUSTER MINIMUM 









< 



















> 















i 




FIRING TI 

1 

ME IS 

; 30 i 

BILLISECONDS 

1111 


20 30 40 

NUMBER OF EARTH PASSES 


50 


60 


70 


FIGURE 4.4.5.0-6 EFFECTS OF ATMOSPHERIC DENSITY VARIATION ON RCS PROPELLANT CONSUMPTION 

FOR THE BASIC TUG 


D5-1714? 



4-138 






200 


RCS 

PROPELLANT TOO 
POUNDS 



1 

NOMINAL CONDITIONS' 
ATMOSPHERIC DENSITY INCREASE 




" 








NOT 

T 

• • 

HE 5 ' 

pass' 

case' 

WAS 

NOT r 

naly; 

:ed 







































— 














— 


























* 

y 
y > 

r 
















y 

y 

y 


i 

AS! 

JUMPT. 

IONS 



- 

L~ 








y 

y 

r 

• PITCH, YAW, AND ROLL ATTITUDE IS CONTROLLED. 







/ 

y 



THROUGHOUT PERIOD OF ORBIT CHANGE. 

• RCS SYSTEM IS DESIGNED WITH: 

1) 5° DEADBAND IN PITCH/YAW (1° WITHIN 
ATMOSPHERE); 2.5° IN ROLL 

2) 4-200 LB HIGH PRESSURE THRUSTERS IN 
THE PITCH AND YAW PLANES 

3) 4-20 LB HIGH PRESSURE THRUSTERS IN 
THE ROLL AXIS 

4) PITCH, YAW, AND ROLL THRUSTER MINIMUM 




/ 

y 

y 

y 






y 

y 

y^ 






y 

0 


i 






y 









FIRING TIME IS 30 

till 

MILL ISECC 

1 

NDS 



1 

1 

0 1 

1 

0 2 

i 

0 3 

n 

0 4 

i i i 

0 50 60 5 

'o 


NUMBER OF EARTH PASSES 


FIGURE 4.4.5.0-7 EFFECTS OF ATMOSPHERIC DENSITY VARIATION ON RCS PROPELLANT CONSUMPTION 

FOR THE 60° FLARE TUG 


D5-17142 



D5-17142 


4.5 THERMAL ANALYSIS 

A detailed thermal analysis of the Space Tug configuration during the aero- 
braking mode should encompass all areas which affect thermal design. The 
limitations imposed by the period of performance allowed for this study, 
however, made it necessary to restrict the thermal analysis to the 
determination of heating rates, equilibrium temperatures and insulation 
requirements, with only a cursory analysis of heat -leak, boiloff and payload 
base area thermal environment. 

4.5.1 Analytical Methods 

Convective heating rates were calculated wi^h the Boeing Convective Heating 
and Ablation Computer Program (CHAP). The ''62 U.S. Standard Atmosphere is 
an integral part of CHAP and was used in the analysis for all calculations 
based on a nominal atmosphere. Convective heating rates based on the 
atmospheric density (+) variation are consistent with the density variations 
used to calculate trajectory dispersions in Section 4.3. 

Equilibrium temperatures based on convective heating rates for both nominal 
and a +3o density atmosphere were determined from the relation, 

T eq = 


where, T eq = equilibrium temperature 
q = heating rate 

o = Stefan -Boltzmann constant 

e s surface emissivity 


An analysis of flow conditions at the base of the Space Tuq payload section 
for the basic (no-flare) configuration was conducted to determine heatinq 
rates and equilibrium temperatures at the payload base area. The analysis 
utilized test data from References 4. 5. 1.0-1 and -2 for separated flow 
behind a base. In order for the flow along a streamline within the mixinq 
layer to be able to overcome the pressure rise throuqh the reattachment 
zone and pass downstream, its total pressure must be greater than the 
terminal static pressure at the end of the reattachment zone. If the total 
pressure is lower than the terminal static pressure, the flow is reversed 
and flows back toward the base. For analytical purposes the maximum base 
temperatures were assumed to occur at the point of impinqement of the 
reversed flow. Empirical relations from Reference 4. 5. 1.0-2 for flow in 
a separated reqion were used ';o determine heatinq rates. Equilibrium 
temperatures were determined from relations given above. 


4-139 


D5-17142 


A 


4.5.1 (Continued) 

Thermal analysis of the Tuq sidewall thermal protection system (TPS) was 
accomplished with a one-dimensional heat conduction analysis utilizing 
finite difference techniques to determine thermal gradients in the side- 
wall insulation blanket. 


4. 5. 1.1 Assumptions 

In order to provide a thermal analysis of the four Space Tuq aerobrakinq 
configurations within the time allocated for this study, certain assump- 
tions were necessary. These assumptions are given below: 

1. Zero angle of attack during each perigee pass through the 
atmosphere. 

2. Equilibrium temperatures based on a surface emissivity of 
0.90. 

3. Heating rates based on a constant wall temperature of 80°F. 

4. Micrometeoroid shield temperature should not exceed 400°F. 

5. Space Tuq thermal protection system will be a re-radiation 
system. Ablative materials will not be considered. 


4.5.2 Heating Rates 


Convective heatinq rates based on aerobrakinq trajectories of 5, 10, 30 
and 60 perigee passes were determined for the Space Tug basic configura- 
tion (no flare) and the 30°, 45° and 60° flare configurations. Maximum 
heatinq rates as a function of the number of periqee passes are presented 
in Figures 4. 5. 2. 0-1 through 4. 5. 2. 0-4 for the four configurations. A 
comparison of heatinq rates for the four configurations is shown in 
Figure 4. 5. 2. 0-5. A typical heating rate distribution as a function of 
perigee pass time for the 1.8 day (10 pass) aerobrakinq trajectory is 
shown in Figures 4. 5. 2. 0-6 through 4. 5. 2. 0-9 for the four confiqurations. 

The effect of atmospheric density variations on the convective heating 
rates is shown in Figure 4.5.2.0-10. The atmospheric density variations 
used in the determination of the heatinq rates are discussed in Section 
4.3. 


4.5.3 Equi 1 ibrium Temperatures 

Maximum equilibrium temperatures for the Space Tug basic configuration 
and the 30°, 45° and 60° flare configurations are presented in Fiqures 
4. 5. 3. 0-1 through 4. 5. 3. 0-4, respectively. A comparison of equilibrium 
temperatures for the four confiqurations is presented in Fiqure 4. 5. 3. 0-5. 


4-140 



D5-17142 



FIGURE 4. 5. 2. 0-1 MAXIMUM HEATING RATES - BASIC CONFIGURATION 


4-141 


MAXIMUM I EATING RATE - BTU/FT^SEC 


D5-17142 



FIGURE 4. 5. 2. 0-2 MAXIMUM HEATING RATES - 30° FLARE CONFIGURATION 


4-142 



D5-17142 



FIGURE 4. 5. 2. 0-3 MAXIMUM I EATING RATES - 45° FLARE CONFIGURATION 




4-143 



MAXIMUM HEATING RATE - BTU/FT 2 SEC 


D5- 17142 



FIGURE 4. 5. 2.0-4 MAXIMUM HEATING RATES - 60° FLARE CONFIGURATION 


4-144 



MAXIMUM HEATING RATE - BTU/FT^SEC 


D5-17142 



0 10 20 30 40 50 60 

NUMBER OF PERIGEE PASSES 


FIGURE 4. 5. 2. 0-5 COMPARISON OF MAXIMUM HEATING RATES 


4-145 









HEATING RATI; - RTU/FI^SEC 



FIGURE 4. 5. 2. 0-6 HEATING RATE DISTRIBUTION - 10-PASS BASIC CONFIGURATION 


D5-17142 



HEATING RATE - BTU/FT2SEC 



FIGURE 4. 5. 2. 0-7 HEATING RATE DISTRIBUTION - 10-PASS 30° FLA] 


































\ 





v 




> 

X 







500 600 


5 


CONFIGURATION 


HEATING RATE - BTU/FT2SEC 



0 100 200 300 400 500 600 700 800 


TIME - SECONDS 


FIGURE 4. 5* 2. 0-8 HEATING RATE DISTRIBUTION - 10-PASS 45° FLARE CONFIGURATION 


D5-17142 






HEATING RATE - BTU/FT^SEC 



FIGURE 4. 5. 2. 0-9 HEATING RATE DISTRIBUTION - 10-PASS 60° FLARE CONFIGURATION 


D5-17142 


MAXIMUM HEATING RATE - BTU/FT 2 SEC 


D5-17142 


NOMINAL ATMOSPHERE 

+ DENSITY ATMOSPHERE 


A 




FIGURE 4.5.2.0-10 MAXIMUM HEATING RATES FOR BASIC AND 60° FLARE CONFIGURATIONS 

NOMINAL AND (+) DENSITY ATM0SP1ERE 




MAXIMUM EQUILIBRIUM TEMPERATURE 


D5-17142 



FIGURE 4.5. 3. 0-1 MAXIMUM EQUILIBRIUM TEMPERATURES - BASIC CONFIGURATION 


4-151 



FIGURE 4.5. 3.0- 


MAXIMUM EQUILIBRIUM TEMPERATURE - °F 


D5-17142 



FIGURE 4. 5. 3. 0-3 MAXIMUM EQUILIBRIUM TEMPERATURES - 45° FLARE CONFIGURATION 


N 


4-153 




4-155 



SPACE TUG BASIC CONFIGURATION 


TRAJECTORY 

[ MAXIMLH EQUILIBRIUM 

TBIPERATURES f 0 FJ 1 

A 

T 


T> 


H 

S PASS 


3175 

1585 

1120 

975 

682 

10 PASS 

2990 

2860 

1410 

987 

875 

591 

30 PASS 

2540 

2420 

1166 

797 

720 

451 

SO PASS 


2140 

100S 

676 

615 

364 



SPACE TUG 45° JTjVRE CONFIGURATION 


TOUECIOttY 

MAXIMUM EQUILIBRIUM TEMPERATURES C°F) 

A 

B 

C 

D 

I: 

F ! 

5 PASS 

2590 

2480 

1195 

927 

1037 

1085 

10 PASS 

2300 

2205 

1040 

800 

874 

919 

30 PASS 

1886 

1793 

812 

60S 

635 

680 

60 PASS 

1660 

1580 

691 

507 

513 

558 


SPACE TUG 30° HARE CONFIGURATION 



MAXIMUM EQUILIBRIUM 

TIM’ERATURES f°F) ! 

A 


C 


E 

F 

S PASS 

2940 


1382 

1082 

1272 

1293 

10 PASS 

2630 

2520 

1215 

943 

1104 

1124 

30 PASS 

2160 

2060 

957 

733 

850 

869 

(ill PASS 

1880 

1790 

809 


706 

729 



TRAJECTORY 

MAXLMIN EQUILIBRIUM 

TEMPERATURES (°F) 

A 

B 


D 

E 

F 

5 PASS 

2325 

2220 

1048 

806 

943 

970 

10 TASS 

20 BO 

1990 

917 

697 

804 

829 

30 PASS 

1740 

1660 

733 

540 

621 

648 

60 TASS 

1512 

1467 

609 

439 

503 

538 


FIGURE 4.5.3. 0-5 COMPARISON OF MAXIMUM EQUILIBRIUM TEMPERATURES 


D5-17142 





















D5-17142 


\ 


Equilibrium temperatures based on atmospheric density variations are 
presented in Figure 4. 5. 3. 0-6 and compared with nominal atmospheric 
density based equilibrium temperatures. 

4.5.4 Thermal Protection System Requirements 

The thermal protection system (TPS) proposed for the Space Tug aerobraking 
configurations is a re-radiation system utilizing a passive approach to 
thermal control. Temperatures behind the re-radiation shield are controlled 
by layers of insulation either directly attached to the re -radiation shield 
or remote, such as local insulation of temperature critical components or 
structure. 

4. 5. 4.1 Heat Shield Dome 

The Space Tlig heat shield dome is a re-radiative structure capable of performing 
within the thermal environment imposed on it during the aerobraking mode. 

Materials considered for various temperature ranges are discussed in Section 4.7. 
Hie dome is designed as a hot structure and must be thermally isolated at all 
support and contact points. Temperature critical components and structure behind 
the dome will be subjected to radiation from the backside of the dome, consequently 
local thermal protection in the form of foil-backed microquartz insulation should 
be provided in this area. 

4 . 5 . 4 . 2 Flare Structure 

The flare configurations are designed to allow re -radiation thermal control 
during the aerobraking mode. There is no requirement for insulation on the 
back surface of the flare structure. All supporting structure and attach points 
are assumed to be thermally isolated. 

4 . 5 . 4 . 3 Sidewal 1 Area 

The Space Tug sidewall area, Astrionics Module and payload section (basic no-flare 
configuration only) will require insulation in order to prevent the micrometeoroid 
shield from exceeding 400°F. A passive system utilizing a 0.002 inch Titanium 
outer skin with a layer of microquartz insulation attached was analyzed for 
thermal adequacy. A sketch showing a cross-section of the Tug sidewall is 
shown in Figure 4. 5. 4. 3-1. The basic configuration (no flare) will require 
insulation along the Tug sidewall, Astrionics Module, payload sidewall and payload 
base area. The flare configurations will require insulation along the Tug 
sidewall and Astrionics Module only. The payload section is assumed to be 
thermally shielded by the flare structure. Minimum insulation thicknesses for 
the basic configuration and the 30°, 45° and 60° flare configurations are 
presented in Figures 4. 5. 4. 3-2 through 4. 5. 4. 3-5, respectively. 












D5- 17142 


NOTES: 1. ALL DIMENSIONS ARE INCHES 

2. CROSS-SECTION IS TYPICAL 
FOR LH 2 TANK WALL 



1.41 MULTILAYER 
INSULATION 


0.080 ALUMINUM 


FIGURE 4. 5. 4. 3-1 SPACE TUG SIDEWALL CROSS-SECTION 



INSULATION THICKNESS - INCHES 


D5-17142 



FIGURE 4. 5. 4. 3-2 SPACE TUG SIDEWALL INSULATION UlICKNESS - 

BASIC CONFIGURATION 


4-159 



INSULATION THICKNESS - INCHES 


D5-17142 



FIGURE 4. 5. 4. 3-3 SPACE TUG SIDEWALL INSULATION THICKNESS - 

30° FLARE CONFIGURATION 


4-160 







INSULATION THICKNESS - INCHES 


D5- 17142 



FIGURE 4. 5, 4. 3-4 SPACE TUG SIDEWALL INSULATION THICKNESS - 

45° FLARE CONFIGURATION 


4-161 


INSULATION THICKNESS - INCHES 


D5-17142 



FIGURE 4. 5. 4. 3-5 


SPACE TUG SIDEWALL INSULATION THICKNESS - 
60° FLARE CONFIGURATION 



D5-171 42 


4. 5. 4. 4 Thermal Protection System Weights 

Thermal protection system (TPS) weights for the basic configuration and 
the 30°, 45° and 60° flare configurations (Tug sidewall, Astrionics 
Module and payload section only) are presented in Figure 4. 5. 4. 4-1. 

These weights are based on the minimum insulation thickness defined in 
Section 4. 5. 4. 3 and include the weight of the 0.002 inch titanium outer 
skin. The effect of increased heating due to atmospheric density varia- 
tions on the TPS weights is also shown in Figure 4. 5. 4. 4-1. The dome 
heat shield (including local insulation requirements) and flare re- 
radiation shield weights are presented in Section 4.8. 

The thermal analysis presented and discussed herein is considered 
adequate for the Space Tug aerobraking feasibility study. There are 
areas, however, that should be investigated in mere detail in order to 
better define thermal performance of the aerobraking configurations. One 
of these areas is the thermal protection system. A detailed analysis 
of the thermal protection system should be performed in order to optimize 
system performance and weight, including evaluation of alternate system 
concepts. In conjunction with a more detailed thermal protection system 
analysis, heat-leak analyses are needed in order to define internal 
structural temperatures and propellant boil off. 

4.6 ASTRIONICS ANALYSIS 

The Space Tuq Aerobraking Astrionics System Study was performed by the 
International Business Machines Corporation (IBM), Electronic Systems 
Center, under contract to The Boeinq Company, and for the Marshall Space 
Flight Center, NASA Contract NAS8-27501. 

This astrionics system study is an extension of a previous Space Tug 
analysis performed jointly by the IBM and The Boeing Company. The 
astrionics study results were separately reported in "Astrionic System 
Optimization and Modular Astrionics for NASA Missions after 1974 - 
Preliminary Definition of Astrionic System for Space' Tug Mission Vehicle 
Payload (MVP)", (prior Reference 3.3.0.0-1). 

4.6.1 Introduction 

The objective of the astrionic system analysis task of the study was to 
investigate the navigation techniques and accuracies, system redundancy, 
power-weight-impacts, radiation impacts and resultant physical charac- 
teristics of the astrionic system areas where new technology is required 
and areas where follow-on study effort is desirable were identified. 

The approach to this study effort is given in Figure 4. 6. 1.0-1. Boeing 
data and the study groundrules were inputs to the study. The baseline 
astrionic configuration from IBM Document Number 69-K44-0006H was refined 
to provide the baseline configuration. Navigation analysis was performed 
for the 2, 5, 10 and 15 days missions to determine sensors for aerobraking 
and the magnitudes of navigation uncertainties (i.e., statistical standard 
deviations expected) for the various sensor combinations, A redundancy/ 
reliability analysis was performed to determine baseline (6C-hour, defined 


4-163 


£ 


NEiGirr 


D5-17142 










4-165 


O' 




OUTPUT 


• ASTRIONIC SYSTEM 
DEFINITION 

- NAVIGATION SENSORS 
REQUIRED 

- SHUTTLE ERA 
TECHNOLOGY 

- SYNC ORBIT ’ 
MISSION 

• WEIGHT DELTAS TO 
BASELINE CONFIGURATION 

- MISSION LENGTH 

- REDUNDANCY 

- ACCURACY 
REQUIREMENTS 

e UPDATE CAPABILITY 
OF ONBOARD 
NAVIGATION SENSORS 

• RADIATION IMPACTS 
ON ASTRIONICS 

• NEW TECHNOLOGY 
I IMPLICATIONS 


FIGURE 4.6.1 .0-1 ASTRIONIC STUDY APPROACH 


D5-17142 


D5-17142 




i 


4.6.1 (Continued) 

in prior Reference 3. 3. 0.0-1) astrionic module weights and weight increases 
(deltas) to this baseline for a final system definition. A radiation 
impact study was performed to determine if repeated passes through the Van 
Allen radiation belt would present problems. Study outputs are as indicated 
on Figure 4.6. 1 .0-1 . 

Section 4.6.2 of this report defines the study groundrules and guidelines 
and Section 4.6.3 gives a summary of the study results and conclusions. 
Sections 4.6.4 through 4.6.8 present the detailed analyses performed 
for the navigation, configuration, redundancy, power and radiation areas 
during this study. Sections 4.6.9 and 4.6.10 summarize the recommended 
future study effort to enhance the present study. Appendices B, C and D 
are included for related study data. 

4.6.2 Astrionic System Study Groundrules and Guidelines 

The following groundrules and guidelines were used for this study: 

o The Space Shuttle will be used to deliver and retrieve the 
Space Tuq from low earth orbit. 

o The basic Tug configuration will have minimal changes to 
apply aerobraking components. 

o The geosynchronous payload round trip mission was used as 
the baseline mission to conduct the aerobraking analysis. 

o The aerobraking mode examined in this study was limited to 
unmanned missions. 

o The astrionics system proposed by IBM in “Preliminary Defi- 
nition of an Astrionics System for Space Tuq Mission Vehicle 
Payload", Final Report, was used as the initial baseline 
from which aerobraking astrionics requirements will be 
defined. 

o The technology used for the Tug astrionic system was Shuttle 
era technology. 

o The Tug was based and maintained on the ground for this 
aerobraking study. 

o The Space tug astrionic module was designed to minimize the 
need for ground support during the active mission. 

o The astrionic module was of modular design to provide the 
capability of automatic operation. 

o The astrionic system was designed to be self-sustaining. 

o The astrionic system was designed to use Shuttle related 

components where possible. 

o The astrionic system was designed to be capable of automatic 
rendezvous and docking operations. 

o The 180-day quiescent mode impact was not considered as a 
part of this aerobraking study. 




I 


4-166 



D5-17142 


4.6.3 Astrionic System Analysis Summary and Observations 

The major items for consideration in the astrionic system analysis were 
the weight delta penalties associated with aerobraklnq and the analysis 
of navigation uncertainties. A summary of the results is included in 
this section. The detailed results for the astrionic system study ire 
presented in the following sections and in the appendices, 

4. 6. 3.1 Astrionic System Aerobraking Weight Deltas 

To provide a basis for the weight delta analysis, the basic lug conf igura= 
tion (see prior Reference 3. 3. 0.0-1) was updated and weight penalties 
associated with long-duration space-based modes were deleted. The 
resulting astrionic system configuration for a basic 60^hour mission 
with aerobraking weighed approximately 1960 pounds. Four areas were 
evaluated for weight delta impacts: Power; redundancy; radiation; and 

navigation sensors. 

The power analysis showed that for a 1 kw nominal astrionic system power 
load, with power generated by fuel cells, approximately 1,5 pounds of 
H 2 /O 2 and tanks are required for each hour of aerobraking mission time 
required. 

The navigation sensor analysis indicated that the same complement of 
equipment is required for all mission durations and that no equipment 
other than redundant components is added or deleted as the mission time 
increases. Therefore, no weight delta is imposed by navigation. 

The radiation impact analysis also indicated that no weight deltas are 
added due to increased aerobraking mission time. 

The redundancy analysis indicated that additional components and hardware 
must be added to the basic astrionic system configuration as the mission 
time increases to maintain a 99% mission success probability. 

In summary, power and redundancy provide the delta weight impacts required 
for aerobraking. The weight impacts of these items and the total weight 
deltas based on aerobraking mission time are summarized in Figure 4. 6. 3. 1-1. 

4. -6. 3. 2 Navigation Analysis 

The navigation analysis was performed to determine the navigation sensors 
required for the aerobraking mission and to define the navigation accuracies 
obtained with these sensor combinations. The IMU, star tracker, landmark 
tracker, horizon sensor and laser radar were selected as the required 
navigation sensors. It was determined that the same complement of 
hardware was required for all aerobraking mission durations, and that 
only redundant hardware would be added to achieve reliability. 

An IMU is required to provide an inertial reference. Because of platform 
drift over extended periods of time, a star tracker is required for 
attitude update and, in addition enhances the accuracy of navigation updates. 


4-167 



TIME 

WEIGHT 

(HRS) 

(LBS) 

48 

120 

90 

232 

120 

325 

.180 

495 

240 

639 

360 

1042 



TOTAL DELTA 
WEIGHT 


POWER 


REDUNDANCY 


NOTES: (1) NAVIGATION SENSORS AND RADIATION 

DO NOT PROVIDE A DELTA WEIGHT IMPACT 
DUE TO AEROBRAKING. 

(2) BASIC ASTR IONIC MODULE 
WEIGHT IS 1960 FOUNDS 


) (96) (144) (192) (240) (288) . 

AEROBRAKING MISSION TIME 

FIGURE 4. 6. 3. 1-1. ASTRIONIC SYSTEM AEROBRAKING WEIGHT DELTA SUMMARY 











D5-17142 


4. 6. 3. 2 (Continued) 

The landmark tracker is required for precision navigation updates and, 
since the effective range of the landmark tracker is limited to low 
altitudes, a horizon sensor is required for navigation updates at high 
altitudes. The laser radar is required for automatic rendezvous and 
docking but was not considered in the navigation analysis. 

The results of the navigation analysis indicated that the navigation 
perigee uncertainties were rather insensitive to configurations (basic 
or flared Tug). A summary of the perigee uncertainties for the basic 
configuration is shown in Figure 4. 6.3. 2-1. This figure shows that the 
maximum Iff uncertainty (first pass of mission) expected for the 2 to 
15 day missions (10 to 85 passes) are less than .35 NM and the steady 
state values are in the order of 0.05 NM. It should also be noted that 
the navigation perigee uncertainties are generally smaller for the longer 
duration mission because of the decreased atmospheric perturbations. 

4.6. 3.3 Radiation Impacts Analysis 

The impacts of repeated passes through the Van Allen radiation belt for 
the aerobraking mission do not appear significant at this time. Assuming 
a vehicle skin thickness of approximately .090 inches of aluminum, no 
problems appear for times up to 10 days. The 15 day analysis indicates 
that care must be taken in the selection of silicon transistors in the 
electronics. Also, additional shielding, included as part of the com- 
ponent packaging, would decrease the effects of radiation. 

4. 6. 3. 4 Observations and Conclusions 

The following salient observations were made durinq the study. It 
should be noted that these observations are for the astrionic system 
only. 

o Additional reactant and tanks for power and additional 

components to maintain acceptable astrionic system reliability 
are the prime contributors to weight increases to the astrionic 
system due to aerobraking. 

o Astrionic system weight increases as aerobraking mission duration 
increases. Therefore, minimum weight deltas are obtained by 
minimizing aerobraking mission time. 

o Long duration space basing of the Space Tug astrionic system 
produces significant weight penalties for the astrionic 
module because of additional thermal conditioning, shielding 
and power requirements. 

o The autonomous navigation configuration used for this study 
can limit radial perigee uncertainties to less than 0.35 NM 
(Iff) for initial passes and to the reqion of 0,05 NM ( 1 cr ) 
steady state. 

o Uncertainties are generally greater for the initial perigee 
passes of shorter duration missions because of increased at- 
mospheric density and more sensitivity to orbital parameters. 


4-169 






D5-17142 


4.6. 3.4 (Continued) 

o In general, better navigation accuracy is obtained for lonqer 
duration missions on comparable perigee passes, 
o Autonomous navigation in synchronous orbit can limit naviga- 
tion uncertainties to a RSS steady state accuracy of 5 NM 
(la) and 2 ft/sec (la) using a horizon sensor, 
o the perigee radial position uncertainties for the 5, 10 and 
15 day missions for both the non-flared and flared fug con- 
figurations are essentially the same. The perigee uncer- 
i tainties for the 2 day mission are slightly higher for the 

non-flared Tug configuration. 

o Navigation updates after perigee and at higher altitudes 
are required to limit uncertainties during those periods and 
to prevent excessive perigee uncertainties on subsequent 
passes. 

o Radiation impacts to electronics by the Van Allen radiation 
belt appear insignificant. The impact increases as the aero- 
braking mission duration increases. 

4.6.4 Navigation Accuracy Analysis 

^ 4.6. 4.1 Introduction 

This section describes the selection of navigation sensors required, the 
navigation accuracy attainable, trajectory correction burn impacts, and 
sensitivities associated with navigation imposed by the aerobrakinq 
maneuver. Navigation sensor selection was based on the following criteria: 

o Navigation sensor availability within the Shuttle era. 

o Consideration of navigation sensor accuracy, 
o Minimizing vehicle delta velocity correction requirements to 
4 reduce RCS fuel consumption. 

The navigation scheme selected for the Space Tug employs a Kalman filter 
v (see Appendix B for brief description of Kalman filtering) to process 

recursively navigation sensor measurement data to obtain an optimal 
estimate of the vehicle "state". Selected combinations of navigation 
components including Inertial Measuring Unit (IMU), star-tracker, horizon 
sensor, sun sensor, and landmark tracker were considered as candidates 
for the Tug mission. 

Data inputs to the study were: 

o Preliminary trajectory data on the aerobraking maneuver from 
The Boeing Company, 

o Vehicle configuration parameters from The Boeing Company, 
o 1962 Standard Atmosphere Density Model, 
o Navigation sensor definition by IBM. 
o Preliminary estimates of the mean and range of density of 
the earth's atmosphere over an 11-year period (obtained 
from The Boeing Company). 


4-171 



D5-17142 


4. 6. 4.1 (Continued) 

The navigation analysis assumed that the perigee was allowed to decay 
at the expected rate, with a constant perigee not maintained. Navigation 
uncertainties were calculated for these orbits and the uncertainties 
shown in this section are all 1 <t. No correction burns were factored 
into the navigation analyses except those discussed in Section 4. 6.4. 6. 

4. 6. 4. 2 Navigation Sensors 

The hardware elements that were considered as candidates for the Space 
Tug aerobraking navigation subsystem are as follows, with detailed 
descriptions given in Appendix D. 

o IMU - Typified by Kearfott KT-70 

The IMU consists of two two-degree-of-f reedom gyros for 
attitude reference and three orthogonal accelerometers 
for velocity increment measurement. 

o Star Tracker - Typified by ITT AEROBEE 150A 

The Star Tracker is a strapped down optical sensor using 
electronic gimballing to determine star positions within an 
eight-degree field-of-view (FOV). 

o Sun Sensor - Typified by Ball Brothers Design 

The sun sensor is a strapped down optical sensor that de- 
termines the location of the sun vector relative to vehicle 
axes, . 

o Landmark Tracker - Typified by the Westinghouse Design 

This optical sensor measures tracking angles to earth features 
such as islands and lakes. 

o Horizon Scanner - Typified by the Lockheed Edge Tracker 

(under development) 

This horizon sensor is an infrared radiometer that scans 
the earth's horizon to determine the vehicle's local vertical. 

A summary of the preceding navigation sensor accuracy characteristics is 
shown in Figure 4. 6. 4. 2-1. 

It should be noted that a ground network was not considered for this 
study due to the desirability of having autonomous navigation. However, 
previous study experience indicates that if a ground network, such as 
MSFN, were available, its update capability would be comparable to that 
described for the landmark tracker/star tracker autonomous sensor 
combination. 


D5-17142 


IMU (KEARFOTT KT-70) 

o 0.02 DEG/HR DRIFT 

o .003% SCALE FACTOR 


LANDMARK TRACKER (WESTINGHOUSE) 

o TOO ARC SEC ANGULAR MEASUREMENT UNCERTAINTY 


HORIZON SENSOR (LOCKHEED) 

0 216 ARC SEC HORIZON AND ATTITUDE UNCERTAINTY 


STAR TRACKER (ITT) 

o 36 ARC SEC ALIGNMENT 


SUN SENSOR (BALL BROTHERS) 

o 120 ARC SEC ANGULAR MEASUREMENT UNCERTAINTY 


Figure 4. 6. 4. 2-1. Typical Navigation Sensor Characteristics 


D5-17142 


\ 


4. 6. 4. 3 Analytical Tools 

The following computer programs were used throughout the analysis: 

o Autonomous Navigation Simulation (ANS) 

The Autonomous Navigation Simulation Program simulates 
autonomous navigation along a Keplerian orbit as performed by 
a Kalman filter using data from selected combinations of 
horizon sensor, star tracker, landmark tracking telescope, 
radar altimeter and range finder (laser). A detailed des- 
cription of this program is contained in Appendix C. 

o Inertial Platform Error Program (IPEP) 

The Inertial Platform Error Program utilizes the normalized 
integral technique developed by George R. Pitman, Jr. (author 
of INERTIAL GUIDANCE) to propogate platform hardware errors 
during the powered phases of flight. 

o Tug Integration Program 

This program is a six-dimensional orbital simulator that 
generates accurate orbital trajectories. The simulation is 
in single precision and utilizes a Runge Cutta 6th order 
integration routine, a PRA 63 or 1962 standard reference 
atmospheric model and a gravitational potential model of an 
oblate spheroid. The program calculates and prints the orbital 
trajectory parameters. A detailed description of this program 
is contained in Appendix C. 


4. 6. 4.4 Synchronous Orbit Navigation Uncertainty Analysis 

Previous studies performed by IBM (Astrionic System Study for Saturn S-II 
Expendable Second Stage - see Reference 4. 6. 4. 4-1) analyzed the navigation 
accuracy attainable using various navigation sensor configurations during 
a 100 NM parking orbit. 

The navigation position and velocity uncertainties for the 100 NM parking 
orbit are as follows: 


Navigational Uncertainties 


Radial (R) 320 ft 
Tangential (T) 370 ft 
Normal (N) 360 ft 


Velocitv 


* 

Radial (R) . 

Tangential . (T) 
Normal ‘ (N) 


' 4-174 


0.53 ft/sec 
0.37 ft/sec 
0.42 ft/sec 


D5-17142 


4.6. 4.4 (Continued) 

The 100 NM parking orbit uncertainties and the errors generated during the 
burn to synchronous altitude were propagated to synchronous orbit altitude. 
These uncertainties were combined with the final orbit insertion uncer- 
tainties to obtain a Root-Sum-Square (RSS) uncertainty for synchronous orbit 
insertion. These uncertainties are as follows: 


Position 

Navigation Uncertainti 

Radial (R) 

54,000 ft 

Tangential (T) 

52,000 ft 

Normal (N) 

3,300 ft 

Velocity 

Radial (R) . 

7.2 ft/sec 

Tanqenti al . (T) 

2.0 ft/sec 

Normal (N) 

1 .0 ft/sec 


These values were used as the initial uncertainties for synchronous orbit 
coast. Accuracies for a spectrum of typical sensors were input to the 
ANS program and a navigation uncertainty analysis performed. The results 
are shown in Figure 4. 6, 4. 4-1. These results indicate that for the accura- 
cies derived from the horizon sensor/star tracker combination ("worst case" 
accuracy), an RSS position accuracy of about 5 NM (la) and 2 ft/sec (la) 
can be achieved after 12 hours in orbit. 

4. 6. 4. 5 Aerobraking Descent Orbit Navigation Accuracy Analysis 

This section analyzes the navigation accuracies attainable using the pre- 
viously described navigation sensors during the aerobraking descent orbit 
mission phase. Two Space Tug vehicle configurations supplied by The Boeing 
Company (no flare and 60° flare configurations) and four different mission 
durations (2, 5, 10 and 15 days) were used to parametrically analyze the 
aerobraking navigation problem. In addition, a brief discussion of a 
one pass deorbit mission is included. 

The aerobraking mission phase begins at the initiation of the synchronous 
orbit deboost burn. The delta velocity required to target for a perigee 
of approximately 50 NM and perform a plane change of 28° is approximately 
6000 ft/sec. The IPEP program was used to determine the navigation 
uncertainties that accumulated during the deboost burn. These uncertain- 
ties were then RSS'ed with the navigation uncertainties obtained from the 
Synchronous Orbit navigation uncertainty analysis (Section 4. 6.4.4). The 



RSS’ POSITION UNCERTAINTY (NM) 



D5-17142 






D5-17142 


4. 6. 4. 5 (Continued) 

resultant uncertainties at the 

Position 

R 

T 

N 

Velocity 

R 

T 

N 


end of the deboost burn were: 
Navigation Errors (Iff) 

20,000 ft 

20.000 ft 
3,300 ft 

3.3 ft/sec 
.66 ft/sec 

1 .0 ft/sec 


These RSS uncertainties were used as the initial navigation uncertainties 
for the onboard navigation program following the synchronous deboost burn. 

A typical aerobraking mission navigation sensor observation schedule to 
provide a base for the navigation analysis was required and it was 
determined using the following criteria: 

o Observations would be made before and after perigee to refine 
the old trajectory for possible midcourse correction burns 
and define the new trajectory after passing through the atmos- 
phere . 

o Observations would not be taken while passing through the 
atmosphere. 

o Time was allotted to perform delta velocity correction burns 
and attitude maneuvers prior to atmospheric entry. 

o The duration of the navigation sightings was limited to the 
navigation instrument operational range. 

o Low deadband and prolonged attitude stabilization require- 
ments should be minimized. 

If a sun sensor were used, the attainment of the attitude requirements 
for "fine" sensing of the sun vector would require prohibitive amounts 
of RCS fuel. In addition, the sun could be occulated by the earth or 
moon thus preventing its use as a navigation sensor. This is parti- 
cularly true in the vicinity of the earth where the sightings are designed 
to occur. Because of these considerations and the update criteria, the 
sun sensor was eliminated from consideration as a candidate navigation 
sensor. 


4-177 



I 


D5-17142 


4. 6. 4. 5 (Continued) 

Investigation of the navigation accuracy attainable using only a horizon 
sensor for navigation updates resulted in navigation uncertainties of 
+_ 9 IW on the first pass through perigee. Such inaccuracies at perigee 
were unacceptable for thermal reasons. Therefore, the horizon sensor 
was eliminated as the prime navigation update sensor. However, it was 
later found to be required to limit uncertainties near apogee and was 
added to the navigation subsystem. 

Thus, the landmark tracker was determined to be the best available sensor 
for obtaining navigation updates. However, the operational range of the 
landmark tracker is limited to the order of 4,000 NM. An update profile 
was selected to assure acceptable operation of the landmark tracker and 
to provide time for midcourse corrections or any required maneuver after 
the completion of navigation updating. The landmark tracker observa- 
tions began aoproximately 1800 seconds prior to perigee (to assure tracker's 
maximum range limit not exceeded) and were stopped approximately 500 sec- 
onds prior to perigee (to al low time for RCS midcourse corrections prior 
to encountering the sensible atmosphere). 

An analysis of the two-day (ten-orbit) aerobraking mission was performed 
using the previously described observation schedule for the landmark 
tracker. This analysis revealed that both the apogee and perigee uncer- 
tainties grew unbounded until the navigation uncertainties during the 
third perigee were larger than typical trajectory errors which would cause 
the vehicle to re-enter. Therefore these uncertainties were unacceptable. 
Investigation revealed that the Kalman filter had to be reinitialized 
after exiting from the earth's atmosphere on each pass because the vehicle 
has then attained a new orbit. Since the Kalman filter maintains know- 
ledge of the past orbit, this knowledge prevents proper convergence of 
the filter and the filter must be reinitialized each orbit after leaving 
the atmosphere. 

After incorporating filter reinitialization on each pass, further analysis 
revealed that navigation uncertainties grew prohibitively during the inter- 
val when no landmark tracker measurements were made. To limit the growth 
of navigation uncertainties during this interval, landmark tracker 
observations were added from approximately 300 seconds after perigee to 
approximately 1700 seconds after perigee. In addition, horizon sensor 
measurements were included in the filter computations from apogee until 
the reception of landmark tracking measurements at approximately 1800 
seconds prior to perigee. The final navigation sensor measurement update 
history developed from the previous analysis is shown in Figure 4. 6. 4. 5-1. 

Basic (Mo Flare) Tug Navigation Analysis 

The initial navigation analysis was performed for the basic (no flare) 
Boeing Tug configuration. Analyses were performed for the 2, 5, 10 and 15 
day mission durations. Results for the 11 day (60 passes) mission were 
considered typical of the results for all mission durations considered 
and are shown in detail to indicate the type of analysis performed during 
this study. All cases considered are summarized in this section. 


4-178 


LT,ST 

UPDATE 

(AT=5 AND 10 SEC) 



AT - TIME BETWEEN UPDATES 


FIGURE 4. 6. 4 . 5-1 ‘NAVIGATION UPDATE HISTORY 


D5-17142 




D5-17142 


4. 6.4. 5 (Continued) 

Using the previously defined navigation update history* the results of 
navigation analysis for the first five orbits of the 11 day (60 passes) 
mission are presented in Figure 4. 6. 4. 5-2. The figure depicts the RSS 
position uncertainty as a function of time from deboost initiation. 

As the figure shows, uncertainties build up prior to apogee, or until 
the horizon sensor can begin converging the uncertainties, after which 
the horizon sensor updates then reduce the uncertainties to the region 
of 5 NM as the vehicle approaches perigee. The landmark tracker then 
quickly reduces the uncertainties and limits them prior to and following 
perigee. After the range of the landmark tracker is exceeded, the uncer- 
tainties grow until the process described above is repeated. 

A plot of the RSS position uncertainty at perigee for the 11 day (60 pass) 
mission navigation accuracy analysis using the basic Tug is shown in 
Figure 4.6.4. 5-3. The RSS position at perigee on the first pass is 0.24 
NM decreasing to less than 0.075 NM (1 <t) steady state. 

The basic Tug aerobraking navigation analysis was expanded to include 
the 2, 5, 10 and 15 day mission. 

Figure 4.6, 4.5-4 provides a comparison of the RSS apogee position un- 
certainties achieved for all the reference missions. The major conclusions 
that can be derived from observations of the data presented on the graph 
are: 

o The second apogee navigation uncertainty is the largest for 
most missions. 

o After the second apogee, the apogee navigation uncertainties 
decrease rather rapidly to a fairly constant value. 

o Initial apogee uncertainties reduce as the mission duration 
increases. 

o The final apogee uncertainties for all missions are less than 
1 NM. 

Figure 4. 6. 4. 5-5 provides a comparison of the perigee radial position 
uncertainties achieved for all the basic Tug reference missions. The 
major observations derived from the figure are: 

o The perigee uncertainties of the 2 day mission are consistently 
highe than those for the longer duration mission because of 
increased atmospheric drag at lower altitudes. 

o The 5, 10 and 15 day periqee position uncertainties are essen- 
tially those achieved by the landmark tracker because of 
reduced atmospheric drag orbital perturbation at higher 
altitudes. Because of this, the difference in perigee position 


4-180 


RSS POSITION UNCERTAINTY (NM) 



FIGURE 4.6.4. 5-2 NAVIGATION POSITION UNCERTAINTIES (10 DAY - 60 PASS) 




NOTE: UNCERTAINTIES ARE 1 (J 


1 

1 

2 

3 

PERIGEE NUMBER 



2 DAY (10 PASSES) 


5 DAY (30 PASSES) 


— 10 DAY (60 PASSES) 


15 DAY (80 PASSES) 


(1 ) 2 DAY-K ALM AN FI LTER 
REINITIALIZATION AT 
PERIGEE +300 SECONDS 

(2) 5, 10 AND 15 DAY-KALMAN 
FILTER REINITIALIZED AT 
PERIGEE +500 SECONDS 



D5-17142 





i: 


PERIGEE NUMBER 


21 


FIGURE 4.6.4 


5-5. RADIAL PERIGEE POSITION UNCERTAINTIES (E 


2 DAY (10 PASSES) 


_ 5 DAY (30 PASSES) 

_ 10 DAY (60 PASSES) 

Q 15 DAY (80 PASSES) 


(1) 2 DAY-KALMAN FILTER REINITIALIZED 
AT PERIGEE -i-300 SECONDS 

(2) 5, 10 AND 15 DAY-KALMAN FILTER 
REINITIALIZED AT PERIGEE 45C0 
SECONDS 

(3) UNCERTAINTIES ARE 1<7 



4. 6. 4. 5 


(Continued) 


uncertainties is negligible after approximately 10 passes 
regardless of mission duration. 

o The perigee uncertainties for the longer duration missions 
appear to stabilize around 0.04 NM after about 10 orbits. 

o During the final passes of the longer missions when the vehicle 
is in the atmosphere longer, the perigee uncertainties increase 
slightly. 

In summary, the apogee and perigee uncertainties are greater during the 
first few passes and reduce to a minimum at approximately 10 orbits. The 
magnitudes of the final apogee uncertainties are small, permitting precise 
navigation at the time of orbit circularization. 

Flared Tug Configuration Navigation Analysis 

The initial conditions, mission timelines and navigation sensor configura- 
tions of the 60° flare Tug configuration analysis are identical to those 
used for the basic Tug (no flare) configuration navigation analysis, 
thus allowing the same rate of decay at higher altitudes. 

The flared Tug aerobraking navigation analysis results are depicted in 
Figures 4. 6. 4. 5-6 and 4.6.4. 5-7. The major observations derived from 
the figures are: 

o The apoqee and perigee uncertainties of the 2 day mission are 
significantly lower for the flared configuration than for the 
basic (no flare) Tug configuration due to the decreased 
atmospheric density and orbital state vector sensitivities 
at higher altitudes. 

o The perigee errors for both the basic and flared Tug con- 
figurations for the longer duration missions are essentially 
the same. 

o The apogee errors for the longer duration missions for the 
flared configuration decay slower. 

One Pass (1/4 Day) Mission Aerobraking Impacts 

The 1/4 day (1 orbit) mission was not considered in the basic navigation 
analysis. However, a "quick look" assessment was made to determine some 
of the impacts expected if a one orbit aerobraking deboost mission were 
considered. Using the previously discussed navigation update history 
and navigation sensors, the navigation accuracy achieved for the one 
pass mission is shown in Fiqure 4. 6. 4. 5-8. 



RSS APOGEE POSITION UNCERTAINTY |NM) 



FIGURE 4.6. 4. 5-6. RSS APOGEE POSITION UNCERTAI 


i 


2 DAY (10 PASSES) 

5 DAY (30 PASSES) 
10 DAY (60 PASSES) 


NOTES: (1) 2 DAY-KALMAN FILTER 

REINITIALIZATION AT 
PERIGEE 4300 SECONDS 


(2) 5 AND 10 DAY-KALMAN 
FILTER REINITIALIZATION 
AT PERIGEE +500 SECONDS 

(3) UNCERTAINTIES ARE 1(7 



(FLARED TUG CONFIGURATION) 


D5-17142 



2 DAY 110 PASSES) 
5 DAY {30 PASSES) 

10 DAY (60 PASSES) 


NOTES: (1) 

2 DAY-KALMAN FILTER 
REINITIALIZATION AT 

. 

PERIGEE +300 SECONDS 

(2) 

5 AND 10 DAY-KALMAN 
FILTER REINITIALIZATION 
AT PERIGEE +500 SECONDS 

13) 

UNCERTAINTIES ARE 1 <7 

> s 


10 DAY 


2 DAY 


11 


PERIGEE NUMBER 

FIGURE 4.6.4. 5-7. RADIAL PERIGEE POSITION UNCERTAINTIES (FLARED TUG CONFIGURATION) 







D5 - 1 71 42 


4. 6. 4. 5 (Continued) 

The navigation accuracies from deorbit initiation to the end of landmark 
tracking are similar to those of the other missions (2, 5, 10 and 15 
days) during the first pass. The radial position error of .35 NM is 
near that achieved at the first pass perigee of the 2 day (10 passes) 
mission. However, the growth of navigation uncertainties is very great 
after perigee. As a result, at filter reinitialization, the navigation 
uncertainty is approximately 18 NM which is too large for the Kalman 
filter to handle using landmark tracker updates (see Reference 4.6.4. 5-1 ) . 
Therefore little improvement in the estimate of the orbit is achieved. 
After landmark tracking, the uncertainties again grow to a value of 15 NM 
at apogee (270 NM). Had horizon sensor updates been used instead of 
the landmark tracker after perigee, a navigation accuracy of about 5 NM 
could have been achieved at apogee. This is because the horizon sensor 
has a greater region of filter convergence than the landmark tracker. 
Therefore, if a one pass mission is contemplated, horizon updates should 
be used soon after perigee to limit the navigation uncertainties to 5 NM 
at apogee. 

Another problem associated with the one pass mission is that the sensi- 
tivity of deboost velocity error to final second pass apogee altitude 
(Figure 4. 6. 4. 5-9) is very large (approximately 10,000 NM/ft/sec). As 
the figure shows, small plus or minus ft/sec variations cause large 
variations in the apogee obtained following the initial perigee. In 
addition, a one pass mission requires a lower initial perigee altitude 
than the longer duration missions. Therefore, without any midcourse 
corrections, the chances of re-entering at the first perigee are great. 

Any midcourse corrections would require a high degree of accuracy to 
obtain the desired perigee. Figure 4. 6. 4, 5-9 indicates that a .25 ft/sec 
error in deboost velocity could result in the vehicle re-entering on 
the first pass. Even the deboost velocity tailoff uncertainty (.5 ft/sec) 
is greater than deboost velocity error necessary to re-enter on the first 
pass. 

The results of a preliminary study for the one pass mission indicate: 

o Accurate midcourse corrections are required prior to the 
first perigee. Even with these corrections, large apogee 
uncertainties are expected. 

o The large orbital sensitivities and large expected error imply 
that obtaining an apogee of 270 NM following the first perigee 
will be difficult. Therefore, delta velocity requirements to 
obtain a 270 NM circular orbit can be expected to be larger 
than those required for a longer duration mission in order to 
overcome these expected errors. 

If the one pass mission is selected for further study, additional effort 
should be expended in the following areas to aid in the determination of 


4-189 


CHANGE IN TARGETED DEBOOST VELOCITY (FT/SEC) 



D5- 17142 





I 


D5-17742 


4.6.4. 5 (Continued) 

feasibility of the one pass approach. 

o Navigation accuracy and associated midcourse correction 
capability to accurately hit initial target perigee. 

o Problems associated with uncertainties in obtaining a 270 NM 
apogee at the end of the first pass and solutions to the iden- 
tified problem areas. 

4. 6. 4. 6 Trajectory Correction Burn Uncertainty Analysis 

Guidance, navigation, atmospheric and burn errors cause the actual orbit 
of the aerobralcing mission to deviate from the desired orbit. Thus, it 
is necessary to provide propulsive correction impulses to force tne 
vehicle back to the desired profile. This can be achieved by two means: 

o Apply an impulse to place the vehicle in the original trajectory. 

o Apply an impulse to achieve a new orbit decay profile that 

achieves the desired end conditions, i.e., perigee altitude. 

This section discusses the navigation uncertainties (1) for selected 
times during navigation updating, (2) for the error magnitudes expected 
using a ballistic trajectory, and (3) for the velocity increments 
necessary to correct for errors encountered. The above data is discussed 
for the first orbit, with selected discussions of the second orbit case. 

The analysis of the first two orbits represents the "worst case" situation 
because the greatest uncertainties in the atmosphere and navigation occur 
during these orbits. Both the knowledge of the actually achieved orbits 
and the knowledge of the earth's atmosphere improve with successive 
orbits. 

Initial Deboost Trajectory Burn Analysis (5 Day - 30 Orbit Mission) 

>i The major trajectory dispersion parameters of the initial deboost orbit 

are: 

o The deboost velocity increment uncertainty at synchronous orbit. 

o The initial state vector uncertainty in synchronous orbit just 

prior to the deboost burn. 

o The Kalman filter determined state vector uncertainties at 
the instant of application of the velocity correction burns. 

o The velocity correction burn guidance and navigation uncer- 
tainties . 

A deboost burn navigation and guidance analysis was performed for a 5 day 


4-191 


6 


D5-17142 


\ 


4. 6. 4.6 (Continued) 

(30 orbit) mission as follows: 

o The initial state uncertainties prior to the deboost burn 
were determined in the synchronous orbit naviqation analysis 
previously discussed. 

o The deorbit burn uncertainties were computed using a desired 
burn profile as input to the IPEP program. 

The results of this analysis are presented in Figure 4. 6. 4. 6-1. The 
total navigation uncertainties in position and velocity due to the deboost 
burn are shown in the figure. These uncertainties when propagated to 
perigee will result in a perigee radial position uncertainty (1 cr ) of 
approximately 10,340 ft (1.72 KM). 

For analysis purposes, it was assumed that a trajectory correction would 
be desired prior to the first perigee. Therefore, the velocity correction 
burn delta velocity and associated fuel consumption were computed at 
different portions of the deboost trajectory. Specifically, computations 
were made: 

o at the conclusion of horizon sensor updates 

o 500 seconds after initiation of the landmark tracker observa- 

tions 

o after termination of the landmark tracker observations 

The results of this analysis are presented in Fiqure 4. 6. 4. 6-2. Shown 
on the same figure are the radial position uncertainties at perigee that 
result from (1) no delta velocity burn correction, and (2) delta velocity 
burn corrections applied to the previously specified times. Both sets 
of data were derived from the initial deboost uncertainties (see Figure 
4.6.4.6-1), the navigation uncertainties determined during the descent, 
and the burn uncertainties. 

Salient features of Figure 4. 6. 4. 6-2 are: 

o It is not advantageous to provide a corrective burn prior to 
the use of the landmark tracker sensor. 


o Fuel consumption can be minimized if the corrective burn is 
performed as soon as the navigation data indicates an 
acceptable perigee can be obtained. 

Second Orbit Burn Correction Analysis (5 day - 30 Orbit Mission) 

The major trajectory dispersion parameters of the second orbit are: 

o The statistical uncertainties of the initial perigee pass. 


4-192 



NAVIGATION 

POSITION 

UNCERTAINTY 


NAVIGATION 

VELOCITY 

UNCERTAINTY 


RADIAL 
POSITION 
AT PERIGEE 
UNCERTAINTY 


PARAMETER 

State Uncertainties 


(feet) 


(feet/sec) 


(feet) 


R 

T 

N 

$ 

X 

N 


20 ,000 
20,000 
3,300 


6594.3 

109.5 


3.3 

0.66 

1.0 


60.7 

6225 


Inertial Navigation 
Uncertainties 


R 

T 

H 

H 

T 

N 


93 

10 

93 


30 


1 .38 
0.16 
1.38 


254 

1510 


Tail off Uncertainty 


RSS 


28,476.4 


0.5 


4720 



10,339.2 


DTE: Uncertainties are lcr . 


Figure 4. 6. 4. 6-1. Deboost Burn Uncertainty Analysis 












RADIAL PERIGEE POSITION UNCERTAINTY (NM) 



FUEL 

CONSUMPTION 


AV 

CORRECTION - 


HS - HORIZON SENSOR 
LT - LANDMARK TRACKER 

NOTE: UNCERTAINTIES ARE 1 U 


AFTER HS 
UPDATES 


BURN 

CORRECTION 

UNDESIRABLE 


■ — FUEL 


PERIGEE 

POSITION 

UNCERTAI- 

NTY 


NO BURN 

CORRECTION XvXvX /XwXwXvXv 

, m » « 4-*4*44«4 • 4 t « • f » * 4 *4 ♦ • * * 

+ + • * 4 4 4 -■ • a 4-4«-»4+*>* 

4 * * • p + • » ■ a * v* *44-4 • 4--»4«4«a-* 

• « ■•••»*4»t + »'**»***«#** r-v b i r* a # 

BURN 

CORRECTION : 

x : xT:T:T : T rTx T:TxTx desirable 

■ # * f * • *•■•*****#**** • * * T — — 

4444 **■•*••***#+44# ■»*«!* ■•*■*« «*«■»•«> 
•«#44#44*^* a 4444 1 44*4 

i«»**»a*r41«4*t***« » * ■* * 1 + *»•»••■ * • « 4* * • 

• •4»44**4*r44«4*4 * • » • * * I * *••«*»•*«••*** 

■*»*■••**♦• • * • 444 

1300 SEC AFTER $&$ 

x-x : : : x->x start of lt updates-:*# 




• * • •'*% i ♦ « 
• • * 4 m ft i* * ft 

**"»*■«**".*« V*'’* 


* * • ft •!• ».*«**»**»•«»■* * ♦ * * • * * * * * <,♦ **.•.**. 

* 4 #« vl9lf9l9-'«a9ar#a9»«»a94i 4*4* i ■ 444444* 
V.%* 4 ■ 4 4 4 4 4 • 4 11* f - 4 4 4 • • * # • P/ 


500 SEC AFTER v# *>X 

START OF LT UPDATES;#:™ 

" ft ■ *<• •>•••■>#■*■•••* * * a- a^* a * • * 1 *** 


TIME FROM PERIGEE (SEC) 


FIGURE 4. 6. 4. 6-2. ORBITAL CORRECTION BURN, AV RCS FUEL CONSUMPTION, AND RADIAL 

PERIGEE POSITION UNCERTAINTY 


FUEL COMSUMPTSON (LBS) 




D5-17142 


4. 6. 4. 6 (Continued) 

o The atmospheric uncertainties of the first pass through the 
atmosphere (as supplied by TBC). 

o The navigation uncertainties during the newly achieved orbit. 

o The trajectory correction burn uncertainties prior to the 
second perigee pass. 

A ballistic, trajectory propagation of first pass perigee uncertainties 
(atmospheric and state vector) was performed. Because of the relative 
magnitude of the propagated trajectory perturbations and the desirability 
of performing corrective burns at apogee, the second pass perigee uncer- 
tainties were computed as a function of the timing of a corrective burn. 
Selected times for corrective burns and resultant perigee uncertainties 
are plotted on Figure 4. 5. 4. 6-3. The problem of performing a corrective 
burn prior to the second perigee is comparable to those for the initial 
deboost phase previously discussed. 

4. 6. 4. 7 Aerobraking Navigation Uncertainty Analysis Sensitivities 

The aerobraking navigation subsystem, analyzed in previous sections 
utilized an observation schedule with specific navigation update rates, 
observation sample intervals, and observation sample initialization times. 
This section analyzes the navigation uncertainty sensitivities to chanqes 
in these parameters, specifically the parameters addressed are: 

o Variation of navigation update rates during landmark tracking. 

o Variation of Kalman filter reinitialization times after perigee. 

Variation of Navigation Update Rates during Landmark Tracking 

The landmark tracker navigation update rate during the landmark tracking 
interval prior to the first periqee was 5 and 10 seconds between samples. 
Samples close to perigee were taken every 5 seconds and were increased 
to 10 seconds between samples at higher altitudes. Other cases considered 
were 10 and 25 seconds between samples and 25 and 100 seconds between 
samples. The results of this analysis are shown in Figure 4. 6. 4. 7-1. 

The figure indicates that the navigation uncertainty at the conclusion of 
landmark tracking is insensitive to navigation update rate during the 
landmark tracking interval prior to perigee. 

The landmark tracker navigation update rate during landmark tracking after 
Kalman filter reinitialization subsequent to perigee was varied in the 
same manner. The results of this analysis are shown Figure 4. 6. 4. 7-2. 

This figure indicates a distinct sensitivity to variations in navigation 
update rate during this landmark tracking interval. The effects of 
this sensitivity is illustrated in Figure 4. 6. 4. 7-3 where the second 
orbit apogee and perigae are greatly affected. The figure indicates 
that a navigation update rate of 25 and 100 seconds between samples 



AV BURN 
AT APOGEE 


FIGURE 


AV BURN 
AFTER 1300 SEC 
LANDMARK 
TRACKING 


AV BURN 
AFTER 500 SEC 
LANDMARK 
TRACKING 


AV BURN 
AFTER 
HORIZON 
SENSOR 
UPDATES 


4. 6. 4. 6-3. RADIAL POSITION UNCERTAINTY AS A FUNCTION OF CORRECTIVE 
BURN TIMING 




RSS POSITION UNCERTAINTY (KFT) 



TIME FROM INITIATION OF LANDMARK TRACKING (SECONDS) 


FIGURE 4.6. 4.7-1. EFFECTS ON NAVIGATION DUE TO VARIATION OF NAVIGATION UPDATE 

RATE DURING LANDMARK TRACKING PRIOR TO PERIGEE 


D5-17142 



RSS POSITION UNCERTAINTY (KFT) 



FIGURE 4.6. 4. 7-2. EFFECTS OF NAVIGATION DUE TO VARIATION OF NAVIGATION UPDATE 

RATE DURING LANDMARK TRACKING AFTER PERIGEE 


D5-17142 













NAVIGATION UPDATE 













D5-17142 




4. 6. 4. 7 (Continued) 

produces large second perigee errors. The 10 and 25 second interval 
navigation update rate also produces significantly higher apogee uncer- 
tainties than for the 5 and 10 second update interval. Therefore, the 
5 and 10 second interval sample rate is recommended during the landmark 
tracking after perigee. 

Variation of Reinitialization Times for the Kalman Filter after Perigee 

The Kalman filter reinitialization time after perigee was varied between 
300 and 600 seconds for both the 2 day (10 pass) and 11 day (60 pass) 
missions. The results of this analysis are presented in Figure 4. 6. 4. 7-4. 
The major observations derived from the figure are: 

o The 2 day reinitialization time is very critical and should 
not occur more than 500 seconds after perigee. Best results 
are achieved if the filter is reinitialized as soon as pos- 
sible after leaving the atmosphere following perigee. 

o The filter reinitialization time for the 11 day (60 pass) 
mission indicates the lack of any sensitivity. Because of 
this, the longer duration missions can be reinitialized at a 
later time without effecting the apogee uncertainty. 

Oblateness Sensitivities 


The earth's oblateness has a significant effect on orbital characteris- 
tics and must be taken into account when targeting for a perigee which 
will achieve the desired mission duration. Although the orbital pertur- 
bations due to oblateness are large, they are considerably more predictable 
than those due to the atmosphere. Therefore, the uncertainties of long 
duration prediction of oblateness may result in trajectory perturbations 
that may require burn correction. These uncertainties are expected to 
be small in magnitude, 

4.6.5 Astrionic System Configuration 

The previous Tug astrionics study (see prior Reference 3. 3. 0.0-1) per- 
formed by IBM used shuttle-era technology and components for the Space 
Tug astrionic system. Since that study, some of the shuttle concepts 
have changed. Therefore, to present an up-to-date system, selected 
Space Tug astrionic components were replaced with comparable shuttle-era 
components. In addition, the previous effort included a spectrum of 
missions and an all-purpose Tug concept. For this study, the astrionic 
system was updated to reflect the requirements imposed by a synchronous 
orbit mission only and the deletion of excess weight penalties due to 
extended periods in space, such as the 180-day quiescent mode. The 
resulting configuration and its associated component weights and charac- 
teristics for the non-aerobraking synchronous orbit mission (60-hour 
length) was used as a baseline for this study. 


4-200 


RSS APOGEE POSITION UNCERTANTY (NM) 


V 





TIME OF KALMAN FILTER REINITIALIZATION (SECONDS) 


FIGURE 4. 6.4. 7-4. 


SECOND APOGEE RSS POSITION UNCERTAINTY VS. TIME OF 
REINITIALIZATION OF KALMAN FILTER 


D5-17142 


D5-17142 


4. 6. 5.1 Astrionic System Configuration Description 

The updated Space Tug astrionic system configuration is shown in Figure 

4.6.5. 1- 1. A centralized computer (CPU/10) concept with a data bus for 
data transmission is used as the baseline. A core main storage (40K 

of 32 bit words) and a magnetic tape mass storage supplement the computer. 
Acquisition, Control and Test (ACT) units provide a standardized interface 
between the data bus and the other subsystem components. 

The navigation sensors are as shown, with inertial reference provided by 
the IMU, automated rendezvous and docking capability provided by the 
laser radar, attitude update capability provided basically by the star 
tracker, and navigation update capability provided by the horizon sensor 
and landmark tracker. 

The onboard sequencing and monitoring function and interface with engine 
controls are also provided by the ACT units. ACT/Select Buffer provides 
the capability of extracting selected data from the data bus and is used 
in conjunction with the communication and external vehicle interfaces. 

Power is provided to the astrionic system by 1 Kw fuel cells with thermal 
control provided by a combination active/passive system using coolant, 
cold plate/radiators/louvers combinations. The astrionic components are 
housed in an astrionic module. 

It should be noted that the philosophy for design of the astrionic system 
is based on using the shuttle components, if possible and as applicable, 
to provide commonality between space vehicle astrionics and to minimize 
development cost and associated hardware cost for other components used 
on the Tug. Selection of other components, such as 1 Kw fuel cells instead 
of 6 - 7 Kw fuel cells used on the shuttle, is based on meeting the astrionic 
system requirements with shuttle-era technology. 

4. 6. 5. 2 Weight and Power Summary 

The weight and power summaries for unit components and for redundant basic 
astrionic system based on the redundancy study (detailed in Section 4.6.6) 
is shown in Figure 4. 6. 5. 2-1. The subsystem weight and power summaries 
and astrionic module weight and power totals are summarized in Fiqure 

4. 6. 5. 2- 2. 

4.6.6 Redundancy Analysis 

A redundancy analysis was performed to determine the redundancy required 
for the nominal 60-hour Space Tug synchronous mission and to determine 
the weight deltas to perform for additional periods of up to 15 days. 

The astrionics configuration was described in the preceding section. This 
section explains the redundancy analysis performed for this study. 


4-203 


V' 


NOTES: 



• REDUNDANCY NOT SHOWN 

• ACT = ACQUISITION CONTROL AND TEST UNIT 

• CPU/IO - CENTRAL PROCESSOR UNIT/INPUT-OUTPUT UNIT 

• IMU = INERTIAL MEASUREMENT UNIT 


FIGURE 4. 6. 5. 1-1 TUG ASTRIONIC SYSTEM CONFIGURATION 


D5-17142 



















D5-17142 



UNIT 

REDUNDANT ( 

ION FIG UR AT 

[ON i 






Total 

Total 

COMPONENT 

Power 

Weight 

Quantity | 

Power 

Weight 


(watts) 

(lbs) 

p 

B 

(watts) 

(lbs) 

DATA MANAGEMENT 






! 

CPU/10 

136 

34 

i 

1 

136 

68 

Main Storage (8K BOM) 

43/56(1) 

28.7 

5 

1 

227 

142 

ACT 

1.5 

4 

40 


60 

160 

ACT/Select Buffer 

75 

13 

3 


150 (2) 

39 

Magnetic Tape 

15 

10 

1 

1 

15 

20 

Data Bus 


7.5 

2 

1 

-- 

22.5 

Total 





588 

452 

NAVIGATION | 







IMU 

120 

47 

1 

1 

120 

94 ! 

Landmark Tracker 

40 

30 

1 


40 

30 

Horizon Sensor 

10 

10 

1 


10 

10 

Star Tracker 

21 

18 

1 


21 

18 

Laser Radar 

30 

28 

1 

1 

30 

56 

Total 





221 

208 

POWER 







Fuel Cell 


79 

1 

1 

__ 

158 

o 2 /h 2 

h|/ 02 Tanks 
DC Regulator 



1 

1 



90 

— 

5 

2 


— 

10 

Power Distributor 

— 

39 

1 


— 

39 

Aux. Pwr. Distributor 

-- 

13 

4 


— 

52 

Junction Box 

— 

2 

8 

! 

— 

16 

Mounting Hardware 

— 

-- 

- 


-- 

20 

Wires & Cables 

— 

— 

- 


-- 

130 

Total 





0 

515 

COMMUNICATIONS 







USB Equipment 

23 

32 

1 


23 

32 

USB Antenna 

— 

2.5 

2 



5 

USB Dir. Antenna 

10 

30 

1 


10 

30 

Command Decoder 

12 

21 

1 

i 

12 

42 

Antenna Selector 

5 

1 

1 


5 

1 

Signal Processor 

13 

15 

1 


13 

15 

TV Camera & Control 

9 

7 

1 


9 

7 

Total 



i 


72 

132 


NOTES: (1) 43 watts standby/56 watts active 

(2) Assumes 2 of 3 active 
P - Primary; B - Backup 


Figure 4. 6. 5. 2-1. Nominal Mission Weight and Power (Sheet 1 of 2) 






















D5-17142 



TOTAL 

TOTAL 


POWER* 

WEIGHT 

ASTR IONIC SUBSYSTEM 

(watts) 

(lbs) 

DATA MANAGEMENT 

588 

452 

NAVIGATION 

228 

208 

ELECTRICAL POWER 

— 

515 

COMMUNICATIONS 

72 

132 

THERMAL CONDITIONING 

150 

303 

INSTRUMENTATION 

TBD 

50 

STRUCTURES 

— 

300 


TOTALS 1031 1960 


♦Nominal active units 


Figure 4. 6. 5. 2-2. Height and Power Summary 


D5-17142 


/ 


\ 


4. 6. 6.1 Redundancy Study Assumptions 

Several assumptions were made for this study. These assumptions along 
with explanations, as applicable, are as follows: 


o Equipment on/off failure rate ratio was assumed to be 3:1, i.e., 
equipment was 3 times as likely to fail when operating as when 
it v/as powered off. 

o The goal was to achieve a mission success probability equal to 
or greater than 99% for all mission times considered. As 
mission time increased, system modifications to improve re- 
liability were made to maintain the 99% success criteria. 

o Only components deemed mission critical were considered in the 
redundancy analysis. The components considered and their 
associated typical failure rates are given in Figure 4. 6. 6. 1-1, 

o The component unit weights considered for this study are 
assumed to include Built-In-Test-:‘quipment (BITE), self- 
test hardware, or software test capability to obtain 95% 
coverage (which is the probability of successful fault detec- 
tion, isolation, and switching) without adding additional 
hardware. 


4. 6. 6. 2 Analytic Programs 


The reliability calculations for this study were performed on an APL/360 
time-sharing terminal system. The use of such a system enables an inter- 
active design process in that the engineer can quickly model the system 
and ascertain its expected performance. The system parameters can be 
easily modified and the improvement (or degradation) in performance as- 
certained. The design process is continued until an acceptable system 
configuration is attained. 


The basic important input parameters are component failure rates, coverage 
and on/off failure rate ratios. Using the above input data, the program 
calculates component reliability, expected component failure numbers, 
mission success probability, the mission losses due to equipment failure 
and to uncoverage and the weights of the configured redundant systems. 

4. 6.6.3 Reliability Enhancement/Coverage 

In applying redundancy to a simplex system, several means of enhancing 
reliability are prominent. Redundant simplex components may be added 
and/or equipment coverage (the capability to detect, isolate and switch 
failed component), may be increased. The former adds weight as simplex 
units are added, while the latter case adds weight in the form of fault 
recognition reconfiguration and recovery hardware to aid in coverage. 

It is significant to note that adding spares alone can go only so far in 
enhancing reliability. When spared components are increased above some 



4-207 


COMPONENT 


D5-17142 


FAILURE RATE 
(Per 10 6 Hours) 


CPU/IO 50 

Main Store (8K BOM) 50 

Magnetic Tape 77 

IMU 200 

Landmark Tracker 25 

Horizon Sensor 8 

Star Tracker 38.5 

Laser Radar* 300 

Fuel Cell 71 

Coolant Pump 19.4 

USBE 1 .3 

USBE Antenna 0.09 

Command Decoder 40 

Antenna Selector 0.18 

Signal Processor 1.3 

*Required for rendezvous and docking at synchronous orbit only. 


Figure 4. 6. 6. 1-1, Failure Rates of Critical Components 


4-208 


DS- 17142 


4. 6. 6. 3 (Continued) 

level, additional spares will not significantly aid reliability unless 
failures in the original spares can be detected, isolated and switched. 

This explains the importance of increasing component coverage as well 
as the need to add spares. 

For this study, a 95% baseline coverage is assumed for all components. 

With this assumption, greater reliability enhancement is obtained by 
initially adding spares. After some redundant units have been added, 
coverage becomes the dominant factor in reliability enhancement, and 
becomes more and more prominent as the mission time increases. The 
redundancy study results reflect the addition of both spares and coverage 
to the Tug astrionic system. 

For the reliability calculations previously described, a coverage weight 
model, based on previous study experience which indicated that the BITE 
approximately doubled the weight of a MARCS computer when circuits were 
added to obtain coverage of 99.5%, was used to obtain the weight deltas 
for increasing the coverage above 95%. It was further assumed that the 
unit weight of the components listed in Section 4.6.5 included 70% for 
simplex hardware weight and 30% for BITE weight to obtain 95% coverage, 

\ 4.6. 6, 4 Nominal Mission Without Aerobraking 

The initial redundancy study effort was to enhance the reliability of the 
simplex system configuration described in the precedinq section. Using 
a mission time of 60 hours, redundant components were added until the 99% 
mission success goal was attained. The results of this analysis, showing 
the redundant system with its associated weights, was shown in previous 
Figures 4.6. 5. 2-1 and -2. 

4.6. 6.5 Aerobraking Redundancy Impacts 

Using the reliability enhanced configuration developed for the basic 60 
hour mission time and also using the assumptions previously detailed, 
delta weights were developed for maintaining the 99% mission success 
probability as the aerobraking maneuver increased the mission time by 
up to 15 days. Weight analyses were made for delta mission times (above 
the 60 hour basic mission) of 48, 90, 120, 180, 240 and 360 hours. The 
resulting weight deltas due to aerobraking are summarized in Figure 
4. 6. 6. 5-1. The results show that the weight deltas for the 4 - 8 day 
aerobraking mission range from 100 to 240 pounds while increases of greater 
than 500 pounds are expected for the 15 day aerobraking mission. 

4. 6. 6. 8 Redundancy Analysis Observations 

The redundancy weights increase as a function of mission time and minimum 
weight deltas would be obtained by minimizing aerobraking mission time. 

Another observation which is significant is the sensitivity of sparing 
and coverage to mission time. Equal reliability enhancement occurs 


4-209 


WEIGHT (LBS) 



DAYS 

(HOURS) 


FIGURE 4. 6.6. 5-1 


AEROBRAKING REDUNDANCY WEIGHT IMPACTS 


D5- 17142 





D5-17142 


4. 6. 6. 6 (Continued) 

whether spares are added or coverage is increased in the 4- to 6-day 
region. Below 4 days, adding spares supplies the greatest enhancement, 
while above 6 days, increasing coverage is the pacing factor. Since 
the implementation of fault detection, isolation and switching is not a 
trivial problem, increasing aerobraking time above 6-8 days is a risk in 
the area of redundancy implementation. Therefore, aerobraking mission 
time below 6-8 days incurs less redundancy implementation risk. 

4.6.7 Power Weight Impacts 

The Space Tug power subsystem is included in the astrionic module for 
this study. As aerobraking mission time increases, the onboard total 
power consumption increases and there is an associated weight delta. 

4.6. 7.1 Space Tug Power Requirements 

To determine the weight deltas for power, the astrionic module power 
requirements, based on this study and the previous studies, were cal- 
culated to be an average of approximately 1 Kw. This assumes that the 
astrionic module is powered continuously, and the observations from this 
study proved that is a realistic assumption. The selected 1 Kw fuel cell 
can handle these loads with inherent fuel cell capability for handling 
larger peak loads. If the fuel cells were also used for both propulsion 
module and astrionic module power loads (500 w and 1 Kw, respectively), 
then a redundant fuel cell operating during peak power periods could 
easily handle the expected power requirements. 

4.6. 7.2 Aerobraking Weight Deltas 

The H 2 and O 2 reactant weights and the tanks (required to store the 
reactants) weights were calculated to obtain the weight delta required 
for power during aerobraking. The H 2 /O 2 consumption is approximately 
0.83 Ibs/Kwh, while the tanks are approximately 0.4 lbs/lb reactant. The 
weight for 1 Kv/h of power would then be 

= 0.83 Ib/Kwh x 1 Kwh + 0.4 lb/ lb reactant x 0.83 lb reactant 

= 0.83 lb + .332 lb - 1.162 Ib/Kwh 

Assuming an approximately 30% contingency for boiloff, peak loading, 
unexpected power loads, and/or increased mission time, the weight per 
Kwh would be approximately 1.5 pounds. 

The weight increase for H 2 /O 2 and tanks is a linear function. Since the 
average power load is 1 Kw/hr and the weight is 1.5 Ibs/Kwh, then the 
weight increase is approximately 1.5 pounds for each additional hour 
used for aerobraking. For example, a 200 hour aerobraking mission would 
require approximately 300 pounds of weight for power. The power weight 
deltas versus aerobraking mission time are shown in Figure 4. 6. 7. 2-1. 

4.6.8 Astrionic System Radiation Impacts 

This section summarizes the results of a study of natural radiation belt 
effects on the electronics of Space Tug systems for the 2, 5 and 10 day 


4-211 



FIGURE 4.6. 7. 2-1 






D5-17142 


4.6.8 (Continued) 

aerodynamic aerobraking profiles, 

A qualitative explanation of the external fluence (particles/ci/) calcu- 
lation is presented first, followed by the internal dose analysis. 

Finally, the ramifications of the internal dose to electronic devices 
are considered. 

4.6.8. 1 External Radiation Environments 

Each aerobraking mission consists of a series of elliptical orbits 
(approximately 30° inclination) with monotonically decreasing eccentri- 
cities starting with maximum apogee at synchronous orbit altitude. The 
natural radiation belts are distributed in an approximately toroidal 
shape around the earth between 70° north and south latitudes and radially 
between about 160 NM to 6,500 NM for the protons. The omnidirectional 
fluxes (particles/cm 2 sec) vary strongly with altitude, orbital inclina- 
tion, particle energy and weakly with time. 

The method by which the cumulative particle fluences were estimated for 
each orbit was to derive an expression for the length of the radial 
vector extending from the center of the earth to the orbit periphery 
and to divide the ellipse into equal areas. 

From the trapped radiation models by Dr. Vette (References 4.6.8, 1-1 
and -2) the flux was obtained at apogee, perigee and •-'adius for each 
orbit. The mean flux between each of these orbital points was determined. 
Given the period of each orbit and employing the Areal velocity law of 
orbital mechanics, (this is, equal areas swept out in equal times), 
the period was divided by 4 and multiplied by each mean flux value to 
obtain the particle fluence encountered in the traversal of each segment 
of the orbital path defined by apogee, perigee and radius. These partial 
fluences were summed over each orbit and again over all the orbits for a 
particular mission length. These particle fluences are shown in Figure 
4.6.8. 1-1 under the headinqs electrons/cm 2 and protons/cm 2 . This appro- 
ximate technique tends to give an over-estimate of the fluence. 

4. 6. 8. 2 Internal Radiation Environments 

Charged particulate radiation damage in electronic devices is manifested 
in two general modes, ionization and atomic displacement. The former is 
characterized by rads (Si), which is the amount of radiation energy 
absorbed relative to. silicon, and the latter by equivalent fission neutron 
fluence (neutrons/cm 2 ). 

The dose in the electronic devices is dependent on the intensity and 
spectral distribution of the radiation penetrating through component 
packaging and vehicle shielding. These in turn are dependent on the 
fluences and spectra incident on the vehicle. A further complication is 
the altitude dependence of the electron and proton energy spectra. Because 
of the large variation in altitude, nominal spectra were used to estimate 



' ft 


■ T-*-. 


* i •? 


... 


• -jfev J^£ vy: n'^-v ^ t ns s^- 


Missions 

IONIZING DOSE AND EQUIVALENT FISSION NEUTRON FLUENCE 

ELECTRONS 

PROTONS 

TOTAL DOSE* 

EXTERNAL DOSE 

INTERNAL DOSES 

EXTERNAL. DOSE INTERIM 

AL DOSES 

INTERNAL 

2 

Electrons/cm 

Rads (Si) 

2 

Neutrons /cm 

2 

Protons/an 

Rads (Si) 

2 

Neutrons/cm 

Rads (Si) 

Case 1 

10 Orbits 
2 Days 

2.1 x 10 11 

103 

7.4 X 10 6 

3.1 x 10 9 

< 310 

10 10 

< 413 

Case 2 

30 Orbits 
5 Days 

3.8 x 10 11 

190 

1.4 x 10 7 

4.6 x 10 9 

< 460 

2 x 10 10 

< 650 

Case 3 

60 Orbits 
10 Days 

8.4 x IQ 11 

418 

3.3 x 10 7 

9.4 x 10 9 

< 940 

3 x 10 10 

< 1350 


o 

*Sum of electron and proton ionizing dose behind 0.63 g/cm aluminum. 


Figure 4.6.8. 1-1. Summary of External and Internal Van Allen Radiation Environments 

For Three Proposed Mission Profiles 


D5-17142 


4. 6. 8. 2 (Continued) 

internal doses. The typical shielding configuration used is shown in 
Figure 4. 6. 8. 2-1. 

Internal Electron Dose 


The internal electron and associated Bremsstrahlung dose/unit incident 
electron fluence was obtained as a function of aluminum shielding thick- 
ness by computer transport calculations (Reference 4.6.8.2-1). This 
function is illustrated in Figure 4. 6. 8. 2-2. It is apparent that for the 
assumed shielding thickness, the electron dose dominates the Bremsstahlunq 
contribution. This function can be used, for example, to compute the 
internal electron dose for case 1 (2 days) behind 0.63 gm/cm 2 by taking 
the product of the ordinate value for that thickness (5 x 10~'0) and 
the incident electron fluence (2.1 x loll) to obtain 103 rads (Si). It 
is also evident that a slight change in shielding thickness will dras- 
tically change the internal dose from Van Allen electrons for shielding 
thicknesses less than 2 gm/cm 2 . 

The equivalent fission neutron fluence for the incident electron spectrum 
was obtained by a conversion of the electron spectrum to an equivalent 
1 Mev electron fluence and the results are shown in Figure 4.6.8. 1-1. 

An appropriate 1.0 Mev electron to fission neutron conversion permits 
integration over shielding thickness to give equivalent fission neutron 
fluence behind 0.63 gm/cm 2 of aluminum. 

Internal Proton Dose 


The natural proton belt spectrum varies so strongly with altitude that 
the choice of a dose/unit incident fluence factor is probably best 
chosen as the maximum value at cutoff energy due to shielding. For 
0.63 gm/cm 2 , the proton cutoff energy is approximately 20 Mev. That is, 
only protons of energy E > 20 Mev will penetrate the electronics. The 
maximum fluence to dose conversion factor at this energy is about 10~? 
rads (Si) - CM 2 /proton at the cutoff. This results in an over-estimate 
of the internal dose by X 2 to X3 at high altitudes and under-estimate 
by the same amount at lower altitudes. This factor was used to convert 
the proton fluences to internal dose in Rads (Si) to obtain the values 
given in Figure 4. 6. 8.1-1. 

Relative to fission neutrons, the protons cause considerably more atomic 
displacement damage than electrons. The values of equivalent neutron 
fluence for the protons are indicated as "worst case" or upper bounds 
because they are based on a solar flare proton spectrum which is somewhat 
lower energy than the Van Allen proton spectrum. 

4. 6. 8. 3 Electronic Device Effects 

At the internal doses indicated in Rads (Si) in Figure 4. 6.8. 1-1 there 
will be no significant effects on electronic devices in the Space Tug 
astrionics. However, it can be inferred from the progression of doses 



4-216 


VEHICLE SKIN 


FI 





) 


INCIDENT RADIATION 


RADIATION DIMENSIONS: 

COLD PLATE - 0.353 

RADIATOR- 0.171 

LOUVERS- 0.103 

0.627 gm 


FIGURE 4 . 6 . 8 . 2- 1 PRESCRIBED SHIELDING CONFIGURATION 


4 


D5- 17142 



BREMSSTRAHLUNG AND ELECTRON DOSE/UNPT FLUENCE (RADS (Si) - cm2/ELECTRONI 


D5-17142 







D5-17142 


4. 6. 8.3 (Continued) 

for the three aerobraking cases that the dose for an aerobraking mission 
time of 15 days (80 orbits) will be well ever 1300 Rads (Si). At doses 
approaching ICr Rads (Si), MOSFET devices begin to suffer a shift in 
threshold voltage. Consequently, there is some question regarding the 
use of these devices should the 15 day mission time be selected. 

The displacement damage effects of the electrons are negligible. However, 
the displacement damage effects of the protons for the 10 day mission 
will begin to cause significant gain degradation in silicon transistors 
with an alpha cutoff frequency of less than 10 MHz, operating at nominal 
current levels and having an end-of-life Beta greater than 100. 

In general, no significant radiation impacts are expected for the aero- 
braking mission times studied. Some electronic components, however, are 
very sensitive to radiation and the observations related to these components 
and for the radiation impacts study are included in this section. 

4. 6. 8. 4 Observations 

If a 15 day mission is selected, it may be necessary to characterize the 
radiation response of M0S devices for a comparison of radiation induced 
threshold voltage shift to circuit tolerances. Should the 2 to 10 day 
mission times be chosen, there is no problem in this regard. 

If silicon transistors with alpha cutoff frequency less than 10 MHz and 
end-of-life Beta greater than 100 are proposed for astrionics design it 
may then be necessary to find substitute transistors for all aerobraking 
missions, depending on the circuit tolerance for forward current transfer 
ratio decrease. The latter is not normally a difficult problem at the 
radiation levels of interest. 

It should be noted that device packaging was not included in the shielding 
configuration. This results in a radiation overestimate for some devices 
with relatively heavy packaging such as power devices. For others, such 
as flatpacks with 10 mil aluminum covers, there is no appreciable over- 
estimate of the dose. Consequently, very slight quantities of localized 
shielding should attenuate the electron and proton fluences to quite 
tolerable levels of MOSFETs and power transistors when device packaging 
is considered. 

Because of the technique used in determining their values, the radiation 
doses established in this study are slightly overestimated. Greater 
accuracy can be attained with computer codes (Reference 4. 6. 8, 4-1) spe- 
cifically designed to calculate fluxes for this type of mission profile. 

Such calculations should be performed prior to making any firm decisions 
regarding the use of additional shielding or generating design groundrules 
for circuit design. 


D5-17142 


4.6.9 New Technology Implications 

Throughout performance of the Jug aerobraking study, functional require- 
ments of the astrionics were compared to conventional implementations 
and evaluations made to determine the need for development of new 
technologies. This section describes the specific items identified which 
would provide a significant contribution to the realization of an aero- 
braking mode Tug. 

4.6.9. 1 Redundancy Implementation 

As discussed in Section 4.6.6 (Redundancy Analysis), the successful 
completion of the Tug aerobraking mission depends on an operational 
astrionics system which in turn depends on "coverage 11 (failure identi- 
fication) to "key-in" redundancy components at the proper time. Current 
technology depends heavily on BITE (Built-In-Test-Equipment), off-line 
dynamic response testing, voting or comparison or reasonableness testing 
to identify failed components. These methods have had limited success 
with electromechanical sensors and hence necessitates the development of 
a new component evaluation technology. 

The basic conventional evaluation limitations can be overcome by using a 
random or pseudo-random noise input and correlation techniques. The 
advantages of using correlation identification techniques are: 

o The system may be checked out "on-line". 

o Test signals can be kept small and will not interfere with 
normal operation. 

o Results can be obtained in the presence of random noise and 
parameter drifts. 

o The technique can be easily applied to existing hardware. 
Correlation is inherently a simple digital process. 

The objective of such a new technology study would be to define the 
suitability and application of "on-line" system evaluation by digital 
methods. 


4. 6. 9, 2 Navigation Sensor Integration 

The navigation analysis performed in this study utilized an optimal 
filter (Kalman) implemented on a general purpose floating point computer 
(IBM 360/75). In a space application the sensor integration routines 
must be programmed in a limited memory machine in fixed point arithmetic. 
A new technology task is recommended to define a sub-optimal filter 
routine in 16 bit fixed point arithmetic to provide integration of the 
landmark, horizon and star sensors and IMU. 


4-.219 


D5-17142 


4. 6. 9, 3 Navigation and Guidance Analyses 

The study effort to date has evaluated navigation uncertainties and 
preliminary evaluated burn corrections with associated navigation up- 
dates. To enhance this effort, a study should be made of the vehicle 
attitude control, retargeting and predictive problems associated with 
orbit disturbances that result from uncertainties in atmospheric drag, 
navigation sensor inaccuracies and vehicle dynamics. The study would 
involve development of a guidance law to (1) predict future orbital 
variations based on past inputs (accelerations, state vectors), (2) com- 
pute new orbital perigees to achieve mission objectives, and (3) control 
vehicle forces during atmospheric braking to achieve the desired end 
conditions. The guidance law would then be tested using available 
simulations to verify its operation under a variety of disturbances. 

4.6.10 Follow-On Study Effort 

The Space Tug Aerobraking study performed to date was an overview of 
selected astrionic system areas and parameters to aid in the determina- 
tion of feasibility of the aerobraking concept to the Space Tug synchronous 
mission. During the course of the study, areas requiring study and/or 
further study have been identified. The following gives brief recommenda- 
tions of follow-on study efforts for aerobraking (and in some cases 
better Tug definition) to aid in evaluation of parameters. 

4.6.10.1 Astrionic System Configuration Analysis 

The analysis performed for the present study was an updating of the 
initial Space Tug astrionic system design using the latest Shuttle 
concepts and components as applicable to streamline the Tug to perform 
only the synchronous mission instead of the broad spectrum originally 
studied. 

Present Shuttle emphasis appears to be in the area of cost without weight 
being a pacing item. For the Space Tug, weight and cost would be pacing 
items. Therefore, additional effort should be expended to integrate 
Shuttle astrionic system concepts and components into the Tug while 
maintaining minimum weight where possible. Although this is' not an 
aerobraking analysis per se, it is important to provide a well-defined 
baseline system configuration as a basis for future aerobraking study 
effort. 

4.6.10.2 Redundancy Analysis 

The redundancy effort to date has defined the typical redundancy weight 
deltas to be expected for aerobraking mission. In the vein of Section 
4.6.10.1, additional effort should be expended to provide a more detailed 
redundancy analysis using the updated astrionic system components, and to 
look individually at each component to determine methods of reliability 
and coverage enhancement. In addition, the risks associated with 
redundancy management for long duration missions should be addressed. 



D5-17142 


4.6.10.3 Navigation Timeline Analysis 

The navigation analysis to date selected a typical navigation update 
timeline which provided satisfactory update accuracy. However, more 
analysis is requied to perform a detailed operations analysis, using 
attitude constraints, sensor acquisition and reacquisition constraints, 
loss of navigation update, burn perturbations, length and frequency of 
updates, etc., to investigate the limitations of autonomous navigation 
during aerobraking. 

4.6.10.4 Advanced Sensor Systems 

This study would look at existing and potential autonomous navigation 
sensors to investigate improvement of autonomous space navigation by 
using new technology sensors and/or concepts. The critical post-perigee 
sighting should be given priority treatment and include investigation 
of the state-of-the-art in interferometer or other all-weather landmark 
tracker technology. A second part of the study would include an analysis 
of navigation sensor hardware to determine means of enhancing reliability 
and to determine operating modes and limitations of the hardware. 

4.6.10.5 Updating Capability versus Control Requirements 

This study effort would be an expansion of the effort included in the 
navigation analysis. The effort would perform detailed trades to de- 
termine the updating capability of onboard navigation components versus 
the required control penalties. This would include attitude control 
deadbands and requirements during sensor observations as well as the 
accuracy versus RCS penalties to perform navigation update burns for 
various times during the aerobraking orbit prior to perigee. 

4.6.10.6 Radiation Analysis 

The radiation analysis for this study was a cursory evaluation to identify 
any significant astrionic system impacts due to repeated passes through 
the Van Allen radiation belt. Additional study effort would include a 
more accurate determination of the elliptical orbit profiles using the 
BOFES (Burrell Orbital Flux and Spectra) code. The impacts of both single 
aerobraking missions ond repeated aerobraking missions for a ground based 
Tug would be evaluated. Sensitivities of various components for various 
shielding and radiation doses would be addressed. 

4.6.10.7 Astrionic System New Technology Component Analysis 

Weight is the pacing item for the Space Tug. The astrionic system weight 
can be reduced if new technology components (in lieu of an off-the-shelf 
Shuttle components) are considered. However, development costs and 
certain development risks would be associated with the new technology 
components. A study effort to evaluate the relative merits of using 
new technology components in the astrionic system of the Tug, giving 
the advantages and disadvantages of weight and cost trades is recommended. 
This would include a survey of potential new technology components (such as 


4-221 


D5-1.7142 


4.6.10.7 (Continued) 

LSI computers) and the relative development progress of each. 

4.7 AEROBRAKING KIT MATERIALS SELECTION 

The materials selection portion of this study was limited to identifying 
suitable materials to be used for the aerobraking kit components without 
a detailed material analysis and selection. The location of the aero- 
braking kit components on the Space Tug .^suited in a wide range of 
potential temperatures on the componcrics . Selection of materials was 
further complicated by the parametric nature of the study. The shorter 
the aerobraking mission time, the more severe the thermal environment 
the aerobraking kits would encounter. 

The materials portion of the study used the aerobraking kit designs 
(Section 4.2) and the thermal data (Section 4.5) as input data for the 
selection of materials. Material selection criteria was established and 
the materials were then selected for each of the aerobraking kit components. 
For the same component, different materials were recommended for use 
dependent on the thermal environment imposed by the mission duration. 

4.7.1 Materials Groundrules and Criteria 

The study was not able to investigate the materials requirements to the 
depth necessary to fully understand the materials/structure/performance 
interfaces. The study groundrules recognized this situation and limited 
the scope of the materials investigation. More detailed studies are re- 
quired in advanced technology and follow-on Space Tug activities. The 
key material study groundrules are as follows: 

a. The Boeing Pre-Phase A Space Tug configuration (prior Reference 
1.1. 0.0-1) scaled upward for increased propellant capability was 
selected as the baseline Space Tug. Modifications were restricted 
to changes to apply the aerobraking kit to conventional Tug con- 
figuration. 

b. The materials technology was restricted to state-of-the-art thermal 
materials. 

c. The TD-nickel -chrome with 2000°F capability was used as an upper 
radiative thermal protection system material limit. Materials 
such as tantalum and columbium were considered advanced state-of- 
the-art, 

d. The use of ablative materials were restricted to single pass 
applications. The recycling capability was considered advanced 
technology. 

e. The maximum allowable temperature the payload may experience was 
limited to 300° F. 


4-222 



D5-17142 


4.7.1 (Continued) 

The materials selection criteria was identified for the aerohraking kit 
components in terms of the type of thermal protection system it would 
provide, i.e., radiative or insulative. Key selection criteria are 
shown in Figure 4. 7. 1.0-1. All of the criteria shown is self-explanatory 
except for oxidation resistance, bonding characteristics, form and purging. 
Oxidation resistance refers to the ability of the material to withstand 
the environments without an additional protective coating which would 
require replacement after each mission. Bonding refers to methods of 
applying insulation to structure (e.g.» welding, brazing, adhesives, 
etc.). Form refers to the ability to obtain the materials in bars, sheets, 
rods, plate, etc. Purging refers to the ability of the insulation type 
materials to be purged of entrapped, heated gases with a cooling gas or 
by drawing a vacuum. Other less critical selection criteria are shown on 
the materials properties sheets (Figure 4,7. 3.0-1 ) . 

4.7.2 Thermal Environments of Aerobraking Kit Elements 

The aerobraking kit elements consist of (1) an aft heat shield, (2) side- 
wall insulation over the propulsion module, astrionics module and payload 
adapter (where required), (3) a flare, and (4) a payload adapter. 

(Astrionics aerobraking kit elements are identified in Section 4.6.) 

The study investigated four different mission durations. These were a 
one, two, five and 11 day missions. Four aerobrakinq configuration conr 
cepts were studied. These concepts included a no flare configuration 
and three different flared configurations. Figure 4. 7. 2. 0-1 tabulates the 
thermal data for the four configurations. 

Although shorter mission durations than the one day (5 pass) were not fully 
investigated, the short duration missions were qiven a preliminary thermal 
investigation. The temperatures that would be encountered on the aft heat 
shield for the basic (no flare) configuration are approximately 3600°F 
to 4000°F for one to two pass missions (References 4. 7. 2. 0-1 and -2). 

For the configurations with flares, the temperatures encountered for one 
to two pass missions would vary from approximately 3300°F to 3700°F. 

For these temperatures, either ablatives or some of the advanced state- 
of-the-art tantalum and columbium alloys (e.g., Fansteel 60) would be 
required for the aft heat shield. 

Short mission durations (less than one day) will cause the sidewall tem- 
peratures to be considerably higher than those shown in Figure 4. 7. 2. 0-1. 
These high sidewall temperatures would require a more complex design with 
both radiative and insulation materials combined to offset the temperatures. 
The temperatures on the flares would also increase and may require the 
relatively heavy TD-nickel -chrome to withstand the thermal environment. 

The payload adapter could be fabricated from aluminum for the flared 
configurations but would require some insulation to protect the sidewalls 
and a cover to protect the aft end of the payload (300°F payload 
temperature limit). For the basic (no flare) configuration, the payload 
adapter is not shielded by a flare and would encounter more severe 


4-223 



CRITERIA 


DENSITY 

THERMAL CAPABILITY 

THERMAL CONDUCTIVITY 

EMISSIV1TY 

TENSILE STRENGTH 

WELDABILITY 

MACHINEABILITY 

OXIDATION RESISTANCE 

BONDING CHARACTERISTICS 

FORM 

PURGING 

COST 

AVAILABILITY 


DOME & FLARE 
(RADIATION) 


SIDEWALL 

(INSULATION) 


LOW WEIGHT | 

LOW WEIGHT 

HIGH TEMP. 

HIGH DECOMP. TEMP. 

— 

LOW 

HIGH 


HIGH AT ELEV. TEMP. 

M 

STRENGTH & EASE 

— 

GOOD PROPERTIES 


NO SPECIAL TREAT. 

— 


EASE 

— 

VARIOUS 


GOOD 

NOT IDENTIFIED IN THIS STUDY 

SHUTTLE ERA (1973 - 74) 


.0-1: MATERIALS SELECTION CRITERIA 


D5-17142 






















































D5-17142 


4.7.2 (Continued) 

temperatures. This payload adapter would require an inconel housing with 
additional insulation on the sidewall and on the payload aft end cover. 

In addition to the four selected configurations, the basic (no flare) 
configuration with no aft heat shield was assessed. The aft heat shield 
must protect the engine systems from the re-entry aerobraking temperatures. 
The RL-10A-3-8 engine used in the Space Tug for this study has many 
limited temperature capability components. This constraint was one reason 
the unshielded aerobraking Tug was not considered practical. For example, 
some elements only have 160°F temperature capability. These elements 
would have to be insulated even when protected by the heat shield. The 
temperature capabilities of these systems are shown in Figure 4. 7. 2. 0-2. 

These temperature limits coupled with the complex aerodynamic flow fields 
of an unshielded nozzle and the aerothermodynamics of the exposed engine 
during aerobraking eliminated the unshielded aerobraking Tug configuration 
from further consideration. 


The payload may only experience a temperature limit of 300°F. The 
temperatures encountered at the payload base for the basic (no flare) 
configuration are as follows: 


Trajectory 


Equilibrium Temperature (°F) 


5-Pass 682 
10-Pass 591 
30-Pass 451 
60-Pass 364 


Therefore, for the basic (no flare) aerobraking Tug configuration, the thermal 
protection system must completely encase the payload. The flared con- 
figuration's payload adapter temperatures were not defined due to the 
complexity of the thermal analyses required to determine the temperatures 
on the payload adapter under the flare and the payload adapter base. 

However, the design of the payload adapter does include a reflective painted 
surface to reduce heat input to the payload. The temperatures at the 
payload adapter base are expected to be below 300°F for the mission durations 
which optimize payloads (30 pass - 5 day missions). For shorter duration 
missions, a protective cover would be required. 

4.7.3 Materials Selection 

The materials which were investigated are shown in Figure 4. 7. 3. 0-1 and 
4. 7. 3.0-2. The effect of temperature on the strength to density properties 
are shown in Figure 4. 7. 3. 0-3. Based on the criteria shown in 4.7.1 and 
the thermal environments shown in 4.7.2, the followinq materials were 
selected for each of the aerobraking kit elements. 


4-226 



D5-17142 


Component 


Limiting Item Material Max , A 1 low,_.Temp_ l 


t ■; 

Thrust Chamber 

Tube Braze 

PWA 85 

1900°R 


Thrust Control 
Va Lve 

Thin Wall 
Housing 

AMS 4130 ' 

760°R 

i 

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trol Valve 

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AMS 4130 

760°R 


Turbopump 

Assembly 

i 

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760°R 

\ ; 

Fuel Injector 
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Angle 

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Teflon 

860°R 

■ >■ 

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620°R 


Pressure' Switch 

Elec. Insulation • ■ 

620°R 


Global 

Assembly 

‘ Conical Mount 
Housing 

AMS 4139 

760°R 


* Reference 4. 7.2. 0-3 


Figure 4. 7. 2. 0-2 RL-10A-3-8 ENGINE MAXIMUM ALLOWABLE TEMPERATURE 

LIMITS FOR NON-OPERATING ENGINE 


4-227 


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D5-17142 
































































































































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D5-77142 








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30 fii 
i3ir>° 

r» 2 t* 
70° - 
1625° 

34. 

12 Hr 
1000° 
F 

16 6 
ki 70° 

■ 


TUNGSTEN 

W 

■ 

KISS* 

1 

HI2° 

OxiHi- 

zee 


032 tt 
70° 

,839*i 

70° 

.279 si 
2192° 

. 311 ftr 
2fi00° 

3 2ft 
70° - 
4532° 

430. 

(.2h dia. 
Wire) 

20. ** 
31100° 
F 

i 59.fi 
70° 
50 + ft 
2102° 

mm 


ZIRCONIUM 

Zr 

399.3 

(.230) 

3:i2<i* 



M».:i u 
70° 

. 06.19 
o to" 


4fi 

3 57*i 
70° - 
2732° 

02 


,13 Tn 
70° 

1 




Figure 4. 7. 3. 0-1 Continued 


t 


D5-17142 







































9 


MKTALS - t-NAhLOYKD 


common or trade name 

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oxidizeH 

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932 

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. fiB h 70 " 


. 04 Ci 7 (P 

3.62 St 70 ° 

120-200 

20 - 30 

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uxtrfizei 

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U 163 ^ 

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rapidly 

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FIGURE 4. 7. 3.0-1 Continued 


0 


tfti annj<rmi "crow 





4-233 


rAllllOX AND GRAPHITE MATERIALS 


COMMON OR TRADE 
NAME 

v: 

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gs 

W 

U b 

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t-, 

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b 

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£ S? 

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2 

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fc 

cs 

c 

o 

Is 


Carbon 

C 

Block 

140.5 

(.0313) 

r.740° 

Oxidi- 

zes 

Al»ve 

500° 



. 1C5 


.99 

3.-24. 
& 70° - 
200° 


- 




Corboo Ptitc H-205 

c 

Plate 

Sheet 

109,2 

(♦063) 

5600° 

i* 


960 




250. 

58. 

av. 

1300° 


3.20 

(Comp. 

HG) 


* 

Great Lakes Carbon 

fH rotvac Graphite 

c 

Plate 

sheet 

124 . n 

(.0744) 

6600° 

If 


13 WG 
2580 AG 

.45 

■ 

.5 WG(a 
1800 
.42 WG 
&2600° 

1.1 - 2.2 

15. WG 
SO. AG 

■ 

■ 



Pyrogenlca 

Carbon Foam nC-4S 

c 

Foam 

Block 

! (.0022) 

6600° 

n 


0.4 

AV. 

.4 


.95 


.33 


B 

■ 


Alrco Speer 

Graphite Foam RG-45 

C ; 

j 

Foam . 
Bloc* 

i 4.99 
I (.0029) 

GM0° 



.4176400°- 

-1600° 

.32A\fe 

70°— 

3000° 

■ 

.95 


.045 


B 



Alrco Speer 

Graphite Foam RG-I5T 

C ; 

Foam 

Block 

9.93 

(0053) 

2600° 

M 


0.4 AV, 

M 


.95 


.44 

■ 

■ 



Alrco Speer 

Graph oil (Pyrolytic G . ) 

! 

c 

Flex 

Sheet 

GO 

(♦035) 

6600° 

It 


15.6WG 
300 AG h 
2000° 


■ 

■ 

r 

15. WG. 
-0.2 AG 6 
70°-2000° 



.2 



Nat 1 ! Carbon 

Carbon Felt VDG 

c 

Fib. 

MatU 

5.2 

(jQoat) 

SG00° 

If 


.25fs 100° 
4. Sir 
3500° 

♦ 1T& 
70° 4(r 
2500 

S 

.99 


10’; Def. 
fi.TPSl 




I 

National Carbon 

Graphite Felt 

c 

Fib. 

Mat'l 

5.3 

(.0031) 

5GOO° 

♦r 


1.2 (u 100° 
2.0 (a 
3500° 

. I7(a70° 

. 4 *» 

2500° 


.99 


10T Def. 

(* .35 


- 



Nat'l. Carbon 


FIGURE 4. 7. 3, 0-1 Continued 


























Johns ManviUe 


FIGURE 4. 7. 3.0-1 Continued 


D5-17142 























































D5-17142 





























































S|02 

Ca~0 

Fth, 

Mall* 

Board 

Eagle Pic her lnd 

Ceramic Block 

SlOj 

Ai 2 t>a 

Fib, 

Mail. 

Board 

Eagle Pic her tnd 

Fiberfrax 970 

Aiao, 

Si o 2 

Fib. 

Felt 

Paper 

Carborundum 

rnmmmmmm 

AUOj 

StO^ 

Fib. 

Block 

Carborundum 

EankakJ Paper 

AljO, 

SiO, 

M o 0 

Fib. 

Paper 

Norton Refractories 

mi h 

AM, 

FcjOj 

Fib. 

Mall. 

Norton Refractories 

r 

Ztrconla Felt ZVF - 100 

Zro, 

«f°a 

v 2°3 

Fib. 

Mall. 

Union Carbide 



FIGURE 4. 7. 3. 0-1 Continued 


D5-17142 










































































4-237 


ms 


J* 




CERAMIC - CERMET MATERIALS 


COMMON* OR TRADE NAME 

z * 

— r- 

Z u 

1 

7. 

o 

b. 

Xr 

U * 5^ 

a a *3 

a w 

Sc: 

E H 

H O 

^ a 

u 

J£ 

U 

O 

-5 b. 
w c 

Si (- 

W 

o 

O ~ 
zz u 

tfi h 

xS 
< jr - 
^ H 

CONDUCTIVITY 
BTU» TlR/FT 2 /IN/°F 

5 

So- 
5s a 

U « 

£ H 
ctj a 

to 

£2 
2 X 
£ O 

« Ui 
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Bi 

a u* 

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cn 

7% 

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u*? 

H o 
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1 | 
s * 

£ £ 

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5 o 1 * 
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r y, 

H ~ 
Tr. . 

O -3 
W 

o 

o 

>: 

w 

3 

o 

a 

2 

£ 

a 

d 

2 ■ 


Alumina 

Aluminum Oxide 
Cast-Prewed-Fuwd 

AL-52.9*. <h - 47. IV 

al 2 c> 3 
(98. H7) 

Dense 

Solid 

247 3 

3I>«0° 

2700° 

3000° 

3100° 

240. to 
70° 

50. to 
2000° 

.18 w 
70° 

. 2H fi 
2000° 

.00 to 
2700° 

.4 to 
2700° 

4*4 to 

70° - 1650° 

140. - 
380* - 



1800-3000 

(75-857 

Dense) 


Altgciflt, Porous 
4 Ti Density 

AT O 
2 3 

Brick 

120 

(.070J 

3680° 

2700° 

2300° 

3000° 

fa. 

2400° 

. 1H to 

70° 
.28 to 
2000° 


.4 to 
2700° 



I 

10. t> 
2400° 

1000 160C 

pst '50% 
Dense) 


Alsmlnlt Foam Brick 
is. 2*4 Density 

AI 2 % 

Brlcx 

37. 

(.021) 

3GG0° 

2700° 

2900° 

3000° 

3. fi 
1700° 

. IS to 
70° 
.28 fc 
2000 


,4 to 
2700° 


11, 7 

1 

i-o 



Alumlmm Silicate 
(Sllllraantie S«t*jn] 

Al 2 0 3 - G3V . SjOj - 3Ti 

A1 2 CIs 

SKtj 

Block 

Brick 

Caat 

202 

(H7) 

2950° 

2550° 

1 

2730° 

12.3 & 
200° 
8.8 to 
2732° 

,25 to 

1100° 


*6to 

1800°- 

2600° 

3.12 to 
77° - 
2730° 

75 

(879 

Dense) 

■ 

21. to 
70° 



Alcmitctn Silicate 
MulUte Svstem ■ 

SA^ - 71. 87. 2 SiOj 28V 

3 Al^Oj 
2S102 

Block 

Brick 

Cast 

173. 5 
(.099) 

3326° 

2830° 

3090° 

42. to 70° 
27. 6 hi 
2000° 

.25to 

1100° 

! 


3.35 to 
70° - 
2500° 

100 - 
150* 

1 

20.5 
to 70 a 

1000 - 
3500 PSI 
(75 - S5V 
Dense) 


Magnesium Alumfnat* 

(Spinet Svttentl 

JUjO, - 71.87. 2 MgO - Wi 

At,Oj 

•JMjO 

Fused 

Block 

220 

(-124) 

3873° 

1 

| 

3452° 

102 fa 
200° 
36, U 
2000 ° 

.29 to 

70 °o 

2000 

| 

i , 

5.0to 
70° - 
2000° 

270. 

8. 5 to 
3000° 

34. 5 
e 7« 
20. Mi 
‘2400°' 

i 


Alum. - Lithium - Silicate 
Pvroeeram. CER-V1T C-100 
Al t 03 -24%- Ll 2 C0 3 -t7.77. SjOjj 
5H7 

AUO n 

Si«2 

Fused 

Shapes 

ISfi 

[.0903) 

2532° 

1400° 

i 

1650° 

S.Gfa 

1500° 

.196 



.0 to 
70° - 
1000° 


■ 

13.3 (> 
650° - 
1000^ 

13.000 




FIGURE 4. 7. 3,0-1 Continued 


D5-17142 





















































































































































































D5-17.42 




























































































241 










SERVICE TEJf>. 800*F TO 

I8S0T 





MATERIAL 

WiNAl 

C0WSITI0N 

l 

OXIDATION 
A£S I STANCE 

•F 

•CITING 

RANGE 

•F 

COST 

S/so ft 

.020 IM 
THICK 

OCRS ITT 
LBS/ IN 3 

THERMAL 
EXPANSION , 
IN/IN X 10’* 

TENSILE 

KSI 

YIELD 

KSI 

ELONGATION 

X 

VELLABIltTY 

MACH INABILITY 

AVAILABILITY 
*0 W 1 STATUS 

Inconel 

m 

IM 50-55 
C r 17-21 
Cb*Ta 4.75-5.5 
•4c 2.14.1 
Co 1.0 «• a. 

T1 .65 • 1.15 
Fa tel 

1900 

2200- 

2450 

4.03 

0.296 

TOT - S.9 
?00*F 7.2 
500 *F 7.9 

1000T a. 2 
1SOOT 9.3 
2COGT 10.0 

imm 

jillil 

Rooa - IS 
120QT-15 
140CT 8 

1600T 88 
18Q0T 170 
2000T 125 

Mary Good 

Wald and 
annealed 

with no 

spontaneous 

hardening 

unless 

cooled 

Slowly 

Poor 

Sheet 

Strip 

Plate 

ter 

State- of -art 

Inconel 

1750 

NUCo 70.0 *1n 
F« 5. 0-9.0 
Cr 14.0-17.0 
T1 2.25-c 75 
Cb*Ta 0.>0- 1.2 


2350* 

2600 

2.46 

296 

70T - 5.7 
200T 7.0 
500* F 7.4 
1000T a.i 
1500T 9.1 

Moos - 175 
1 20CT -128 
1 400"F - 98 
1600T 53 

10OOT 9 

Rooa - 125 
I20CT 104 
1400T 76 

1600T 40 

1800T 5 

Rooa . 20 
120QT 10 
1400T 5 
1600*F 12 
1800T 89 

Satisfactory 

Poor 

Tube 

Sheet 

Rod 

ter. etc. 
State -of -art 

lift* 41 

Cr 19.0 
Mi 53 
Cc 11 
•Id 10 
T1 3.2 
A1 1 .60 
Fa 2.0 


2250- 

2500 

9. 14 

.296 

7GT - 
200* F 6.6 
40OT 6 e 
1000T 7.5 

iaoot «.; 

•ooa - 175 
1200T 164 
1400T 140 
•WOT 88 
180CT 40 

Rooa * 130 
1200T 130 
1400T 121 
16O0T 84 

1WOT 25 

Rooa - 10 
120GT 8 

MOOT 5 
1600T 5 
1800T 20 

Good 

Like other 
High alua. 
4 nigh 
tltenlua 
alloys 

Poor 

Machine In 
aged condition 
with Tungsten 
carbide tools 

Soft 6 gwy 

Sheet 

Plate 

Foil 

ter 

Forging 

Billets 












Like Aus tenet 1c 
S. S. in anneal 
condition 

Sute-of-art 







SERVICE TE*V 1800°F TO 2200T 





TO 

Rtckel 

Th02 2 

Ml tel 

2000 

2650 


0 322 

Mean 

70*100 OT 
8.5 

70* F - 65 
180GT 18 
2000T 14 
2400T 9 

TOT 45 
1800'F 17 
2000T 12 
2400*F 7 

TOT 15 
1BOOT 7 
2OG0T 7 
2400T 5 

Poor 

truing 
recowended 
80S Parent 
Material 
strength 

Poor 

Soft & guaev 

Slabs 

Billets 

Rods 

Plates 

Sheets 

Tuba Hollows 
Foil 

VI re tubinq 
State -of-ert 

TO 

Nickel 
O mania 

Cr 20 
ThO, 2 

■1 *tel 

7400 

2600 

110.00 

.306 

Near* 

70-1 800* F 
6 8 

Am 125 
1200T 65 
15Q0T 35 
2000T 16 

Rooa 80 
1200T 54 
15GOT 30 
2000T 15 

Rooa 18 
1200T 8 
1500T 7 
2000T 3 

Poor 

Brazing 

reccaaended 

50-701 

parent 

aaterlel 

strength 

Poor 

Soft A guray 

Sheet 

Rod 

State-of-ert 


FIGURE 4. 7. 3. 0-2: RADIATIVE MATERIAL PROPERTIES 


>■ 




05-17142 



4-242 


MATERIAL 

NOWMAl 0XI3ATI0N 

co*>os:tio* resistaro. 

•f 

KLTPIG 

R8lG£ 

*F 

RAW MATER- 
IALS COST 
U LB 

aensm 

LBS/ 19 J 

SERVICE TDf. 220CT TO 25QQT 

THERMAL TE9SIIE 

EX°A9S!09 , «SI * SI 

lH/iH X 1C" 

ELOAGATIO* 

NELMBIllTY 

HAChlKEABlLlTY 

A/AILA8ILITT 
FGRM A STATE 


Cb 80 


4710 

95. 6C 

0.347 

. 

Room 85 

Rojm 65 

Room 25 

Acceotable 

Good 


C-I2S 

T* 9-11 






2000-f 39 

203CT 29 

2000*F 40 

Electron Beam 


Advanced 


W 9 -11 






2400*F 26 

24)0T 22 

2400T 3 

or TIG 

Threading and 

state-of-art 



MOTE: Coluflfciua alloys exhibit * ductile brittle 


2700T - 

27 »T - 

2700*F - 


tiering are 




transition at subzero temperatures 


3000T 11 

JOOOT 10 

3000*F 65 


often difficult 









3500*F 5 

3500T 4 

3 SOOT 60 


to accomplish 


Coluefciu* 

r+f 10 

. 

4400 


C. 323 

Mr in 

Room 61 

Sr- 5* 4) 

Rcom 25 

Acceptable 

Good 

Advanced 

C-103 

Ti 18 





32-2*2** 

2000T 27 

2000V 2u 

2000*F 45 



state-o^-art 


Z* O.i 





Aon . 0 60 

2400T 13 

2400T 11 

2400T 70 

TIG or 

dig* duett lit* 



Cb 9*1 






2 7 00*F 9.5 

2700V 9 

270CT 70 

Electro**, Bear 

causes -robletis 









300C*F 5 

30UQV 4 

30C0** 70 


dr- Tlfrg 













and tirrfrj 


Fans tael 

T* 26-29 


4695 


CM) 

m 

Stress 

Stress 

Stress 

'•cod 

p l*r 


85 

W 10-12 






Relieved 

Re 1 :eved 

Relieved 





It C.6-1 . 

.1 





Sheet 

Wet 

Sheet 

TIG or 

Terrene* to 

Ingot 


Cb Sal 






-320*F 170 

-32f*F 1)5 

-320T 17 

Electror 9ein 

tear and oall 

Plate 








RT 92 

RT 78 

RT 18 



Stria 








2000*f 45 

20C0*F 42 

2000T 22 



Foil 








24O0T 22 

2400V 21 

2400T 61 



Welded Tubinq 








260C*F 16 

2600V 14 

2600*F 80 



Fat. Parts 








Aecrystalue* 

Recrvst*li:ed 

Recrvstel 1 ized 



Advanced 








-3?0*F 150 

-MOT 130 

-320*F 16 



state-of-art 








RT 82 

RT 60 

RT 23 











2OO0*F 35 

2003T 22 

2000* F 40 











c 400 " F 21 

2401V 19 

240CT 58 











2600T 18 

2600V 15 

263C*- 3. 










SERVICE TE*>. 

2SOO-3SOOT 







Tantalum 

W 8.5 - 11.0 4500 

5495 

95. 3i 

0 608 

Mean 

RT 80 


RT 65 


RT 25 


Feasible 

'•otd 

Inoct 

Fans tee 1 

Ta Bal 




32-21 3*F 

2000T 

. 

200DT 

. 

20XT 

• 

GTA Melded 


B*? I lets 

60 





3.6 

2400T 

22 

2400V 

17 

2400T 

49 

Electron 

Mlon duct* ’ * t* 

Bars 







2 TOOT 


2700V 

. 

27>3T 

- 

Beam 

1 o*u hardness 

Plate 







3000 T 

14 

3000V 

11 

3000T 

50 

Resistance 

is chief cause 

Sheet 







3500T 

9.6 

3500V 

8.4 

3500T 

SO 

Welding 

of machining 

Foil 














problems 

Rod 














Threading and 

Wire 














tapping are 

Tube 














often difffcult 

Fabrication 















Advanced 















state-of-ar 


FIGURE 4. 7. 3. 0-2: 


RADIATIVE MATERIAL PROPERTIES (Continued) 


D5-17142 


TENSILE STRENGTH/DENSITY (INCHES) 



FIGURE 4. 7. 3. 0-3. STRENGTH TO DENSITY PROPERTIES OF HIGH TEMPERATURE STRUCTURAL MATERIALS 


05-17142 


D5-17142 


4.7.3. 1 Aft Heat Shield Materials Selection 

The study qroundrules limited the state-of-the-art radiative thermal pro- 
tection material to 2000°F with TD-nickel -chrome . Therefore, the materials 
shown for temperatures above 2000°F represent areas where material tech- 
nology would be required. The density of the high temperature materials 
are two to three times that of the lower temperature materials. Therefore, 
a different material was selected for the eight-foot movable cap in some 
instances than was used for the fixed portion of the aft heat shield to 
reduce the overall heat shield weight. The materials used are as listed 
below. 


CAP SECTION OF HEAT SHIELD FIXED SECTION OF HEAT SHIELD 


'N 


No. of 

Basic 

U i 

o 

o 

45° 

60° 

Basic 

30° 

45° 

60° 

Passes 

Tuq 

Flare 

Flare 

Flare 

Tuq 

Flare 

Flare 

Flare 

1-4 

A 

A 

A 

A 

A 

A 

A 

FS60 

5 

FS60 

FS60 

FS60 

FS85 

FS60 

FS60 

FS60 

FS85 

10 

FS60 

FS60 

FS85 

TDNC 

FS60 

FS60 

FS85 

TDNC 

30 

FS85 

TDNC 

TDNC 

R41 

FS85 

TDNC 

TDNC 

R41 

60 

FS85 

TDNC 

R41 

R41 

FS85 

TDNC 

R41 

R41 


LEGEND: 

A = Ablatives TDNC = TD-nickel-chrome 

FS60 = Fansteel 60 R41 = Rene' 41 

FS85 = Fansteel 85 


The latching mechanism for the heat shield will use the same materials as 
were proposed for the cap. Heat flow back to the Tuq propulsion module aft 
skirt is reduced by heat blocks. The engine is located well aft (five to 
seven feet) of the heat shield. The limited temperature capability 
components are protected from the radiative heat transfer from the dome 
by (1) the retracted portion of the two position nozzle, and (2) by reflec- 
tive metallic foil mounted atop the insulated limited temperature components. 

4. 7. 3. 2 Sidewall Insulation 

The sidewall temperatures are shown in the prior Figure 4. 7. 2. 0-1. The 
highest temperatures range from 1585°F at the aft skirt of the no flared 
configuration down to 1120°F at the astrionic module/propulsion module 
interface for the 5 pass mission. The lowest temperatures range from 
609°F at the aft skirt of the 60° flared configuration to 430°F at the 
astrionic module/propulsion module interface for the 60 pass mission. 

For all of the configuration options, John Mansville microquartz insulation 
was selected. The material has a long life temperature capability to 
2000°F (melting point 3000°F) and has a low density (3 pounds per cubic 
feet). The thickness was varied so that the maximum allowable temperatures 
on the sidewall was 400°F. The maximum thickness for the 5 pass basic 


4-244 



D5-17142 


4. 7.3.2 (Continued) 

Tug at the aft skirt (1535°F) was 1.08 inches and at the propulsion/as- 
trionics module joint was 0.78 inches thick. The maximum thickness for 
the 60 pass, 60° flare at the aft skirt (609°F) was 0.82 inches and at 
the propulsion/astrionic module joint was 0.38 inches thick. 

To protect the microquartz from damage due to handlinq, transportation, 
vibration, etc., the microquartz will be covered by a thin outer skin 
of titanium (.002 inch). For temperatures too high for the titanium skin, 
an Inconel 718 foil (.002 inch thick) may be used. 

4. 7. 3. 3 Flare Materials Selection 

The selection of the material for the flare was based on the high strength 
to density properties of Inconel 718 over other materials in the 1000°F 
range. The temperatures on the panels near the outer edqe of the flares 
are as shown below: 


No. of 

30° 

45° 

60° 

Passes 

Flare 

Flare 

Fla 

5 

1260 

1075 

970 

10 

1175 

900 

829 

30 

850 

670 

648 

60 

700 

550 

538 


The use of other materials were investigated but were not subjected to a 
weights analyses. Titanium has a lower density than Inconel 718 and may 
be a reasonable flare material for mission with durations qreater than 
the 5 day (30 pass) mission. For longer duration missions or for large 
flares, it may be possible to use a metalized Kapton polyimide film bag 
material (short duration capabilities up to 800°F) as an inflatable flare. 

In addition to the panels which constitute a major portion of the flare 
weight, there are support struts, piano hinqes, cables, and spring hinqes. 

The 36 support struts will not see significant temperatures as they will 
be shielded by the flare panels. Titanium tubular rods may be used for 
all the flared concepts. If necessary, insulation may be wrapped around 
the rods for the high temperatures encountered with short duration missions 
These remaining items constitute a small portion of the total flare weight. 
These items may be fabricated from any material compatible with the inconel 
panels and titanium struts and capable of withstanding the temperatures 
encountered. 

4. 7. 3. 4 Payload Adapter Material Selection 

The allowable temperature limit of the payload (300°F) necessitates a pay- 
load adapter that serves both as a docking device and as a payload heat 
shield for the basic (no flare) configuration. For the flared configura- 
tions, the payload adapter performs the additional functions of flare 
mountinq fixture and flare actuation mechanism housing. The payload adapte 
will consist of (1) skin and stringers, ring flames, cross beams, guide 



05-U142 


4. 7.3.4 (Continued) 

tubes and guide cones. These will be fabricated from aluminum. The 
temperature will have to be controlled so that only 300°F will be en- 
countered. For the basic (no flare) Tuq, a microquartz insulation with 
a titanium outer cover will be used. For the flared configurations, no 
additional protection is required beyond the aluminum skin as the flare 
shields the payload adapter for the medium and lonq duration missions. 

The very short duration flared Tuo missions may require the same micro- 
quartz insulation/titanium cover as was used with the basic (no flare) 

Tug. 

4.7.4 Options and Recommendations 

The materials selected for the aerobrakinq kit were based on a specific 
aerobraking concept and selected kit designs. Other methods of aero- 
braking (forward flare, active cooling with exhaust qases, larqer flares, 
etc.) will significantly impact materials selection. As weight of payload 
is directly impacted by the aerobraking kit weights, low weight is a key 
factor in selection of materials. 

4.8 WEI 6HT S AN D MASS PROPERTIES 

Weight penalties associated with implementation of an aerobrakinq return 
mode of operation are presented in this section. These weights are qiven 
as a function of the number of perigee passes for each of the aerobraking 
kit elements. Figure 4.8.6. 0-1 is a summary of the aerobrakinq weight 
penalties and resulting payload for each Tug configuration. Mass proper- 
ties used in the controls and aerodynamic analyses are presented in 
Section 4.8.1. 

4.8.1 Mass Properties Summary 

Mass properties for the four Space Tuq configurations are presented in 
Figure 4. 8. 1.0-1. The mass properties, which include the weiqht, centers 
of gravity, and mass moments of inertia, were calculated for the conditions 
existing at the start of aerobraking. 

4.8.2 Structural Weights 

Structural weight penalties resulting from the modification of the basic 
nonaerobraking Space Tug are presented in Fiqures 4. 8. 2. 0-1 throuoh 
4. 8. 2. 0-6. Fiqures 4. 8. 2. 0-2 and 4. 8. 2. 0-3 present the weights for an 
aerodynamic shield (aft heat shield) and a flared aerobrake (flared skirt), 
respectively, as a function of the number of passes. The flared skirt was 
sized for repose angles of 30, 45, and 60 degrees. The aft heat shield 
was sized for the aerodynamic effects resultinq from these flare angles in 
addition to a "no flare" configuration. Figure 4. 8. 2. 0-4 presents the 
weights for a combination payload adaoter/f lared skirt support structure. 

Figure 4. 8. 2. 0-6 presents the total structural weiqht penalties as a 
function of the number of passes. 



D5-17142 


WEIGHT AT START OF AEROBRAKING (POUNDS) 
CENTER OF GRAVITY X (INCHES) 

CENTER OF GRAVITY Y (INCHES) 

CENTER OF GRAVITY l (INCHES) 

ROLL MOMENT OF INERTIA (SL.-FT. 2 ) 

PITCH MOMENT OF INERTIA (SL.-FT. 2 ) 

w mi i uAur %it r r » / #»• *? \ 

i rvrt rnjiiuni uf isilim in \ jL.-r I > 


NO 

FLARE 

30° 

FLARE 

45° 

FLARE 

60° 

FLARE 

14,430 

14,430 

14,430 

14,430 

243.3 

234.0 

234.2 

230.7 

0 

0 

0 

0 

0 

0 

0 

0 

10,388 

12,443 

17,021 

21 ,268 

101 ,686 

94,748 

97,078 

94,604 

*• r% 

IU 1 ,ooo 

A * * A 

,/HO 

57 ,076 

54 ,604 


FIGURE 4. 8. 1.0-1 SPACE TUG MASS PROPERTIES 







5 

PASSES 

10 

PASSES 

30 

PASSES 

60 

PASSES 


MATL. 

t 

(IN.) 

MATL. 

(IN.) 

1 

MATL. 


MATL . 

UN.) 

f NO FL^RE 

A 

.137 

A 

.091 

A 

.047 

A 

.030 

l 

30° c LARi 

A 

.091 

A 

.066 

A 

.041 

A 

.030 

! 45° FLARE 

A 

.058 

A 

.038 

A 

.030 

A 

.030 

^60° flare 

A 

.037 

A 

.032 

A 

.030 

A 

.030 

'no flare 

A 

.030 

A 

.030 

A 

.030 

A 

.030 

3C° FLARE 

A 

.030 

A 

.030 

A 

.030 

A 

.030 

45° FLARE 

A 

.030 

A 

.030 

A 

.030 

A 

.030 

^60° FLARE 

A 

.030 

A 

.030 

A 

.030 

A 

.030 


A RENE 41 

A TD nickel chrome 

A FANSTEEL 85 

A TANTALUM FANSTEEL 60 


• MINIMUM MATERIAL THICKNESS = 0.030 (IN.) 
t t * EQUIVALENT MONOQUE THICKNESS 


FIGURE 4. 8. 2. 0-1 AFT HEAT SHIELD MATERIALS SUMMARY 


D5-17142 












FLARE WEI3WT (LBS) 


D5-17142 


1 

.1 1 II 


— 

/^FLARE ANGLE 

— 


1 400 r 



30° FLARE 


+3o ATMOSPHERIC 
DENS I TY '/ARIA! I ON 


45° FLARE 


NUMBER OF PERIGEE PASSES 

FIGURE 4.8. 2. 0-3 FLARE WEIGHT VS NUMBER OF PASSES 











D5-17M2 



. NO FLARE 

! CONFIGURATION 
i (LBS) 

FLARED 

CONFIGURATIONS 

(LBS) 

SKIN/STRINGERS 

200 

200 

RING-FRAMES 

45 

45 

CROSS BEAMS 

-- 

40 

GUIDE TUBES 

25 

25 

GUIDE CONE 

30 

30 

PAYLOAD LATCH FR % GUIDES 

i 

50 

50 

TOTAL 

350 

390 


FIGURE 4. 8 . 2 . 0-5 WEIGHT SUMMARY - PAYLOAD ADAPTER/ FLARED SKIRT SUPPORT 



TOTAL STRUCTURAL WEIGHT (POUNDS) 


D5-17142 



FIGURE 4. 8. 2. 0-6 TOTAL STRUCTURAL WEIGHT VS NUMBER OF PASSES 


4-253 





D5-17142 


4.8.2. 1 Aft Heat Shield 

The aft heat shield is required for protection of the primary propulsion 
system from aerodynamic effects during the aerobraking return mode. The 
basic configuration of the heat shield consists of a stiffened skin elliptical 
dome with a b/a ratio of 2:1. The base diameter of the shield is 168 
inches. A 96 inch diameter removable hatch is included for operation of 
the primary propulsion system both before and after aerobraking. The aft 
heat shield was sized to react the pressure loadings given in Section 4.1. 

The materials used are those specified in Section 4.7 and were selected 
in accordance with the temperature environments of Section 4.5. The heat 
shield operates as a hot structure with no external thermal protection. 

Figure 4.8. £.0-1 gives a summary of the materials and equivalent monocoque 
thickness (t) used in obtaining the shield weight. Figure 4. 8. 2. 0-2 
presents the aft heat shield weight as a function of the number of passes. 

4. 8. 2. 2 Flared Skirt 

The flared skirt is designed as a stiffened frame covered with a thin facing 
sheet. The skirt is supported by 36 tubular struts and can be stowed within 
a 15 foot diameter mold line. Inconel 718 is used for the frame and facing 
sheets with titanium 6A1-4V being used for the support struts. The skirt 
s will operate as a hot structure with no thermal protection being required. 

\ The skirt was sized to react the pressure loading given in Section 4.1 

and the temperature environments of Section 4.5. Figure 4. 8. 2. 0-3 shows 
the total weight of the flared skirts, including deploy/ retract mechanisms, 
as a function of the number of passes. 

4. 8. 2. 3 Payload Adapter/Flared Skirt Support 

The payload adapter is required for protecting the payload durinq aero- 
braking and, with flared configurations, for providing an attach point 
for the flared skirt support struts. The adapter is an aluminum stiffened 
skin structure 168 inches in diameter and 120 inches lonq. Three ring 
frames are employed; a lower bolt ring for attachment to the astrionics 
module, an intermediate ring for support of the flare struts, and an aft 
ring for support of the payload guide. Figures 4. 8. 2.0-4 and -5 show the 
weight of the adapter, less external insulation, for the basic no-flare 
configuration and for the flared configurations. The higher adapter 
weight for the flared conf i gyration is due to the addition of a cross beam 
for support of the skirt deploy/retract mechanism. 

4.8.3 Thermal Protection System Weight 

The thermal protection system (TPS) weight as a function of the number of 
passes is presented in Figure 4. 8. 3. 0-1. The TPS weights for the no-flare 
configuration are for the Tug sidewall, astrionics module, payload side, 
and payload base region. The 30°, 45°, and 50 c flare configuration TPS 
weights are for the Tug sidewall and astrionics module only. 


4-254 


THERMAL PROTECTION SYSTEM WEIGHT (POUNDS) 


D5-17142 






D5-17142 


4.8.4 Astrionics System Weight 

The weight of the astrionics system as a function of the number of passes 

is given in Figure 4. 8. 4. 0-1. The astrionics system weight increase 
with the number of passes reflects the systems sensitivity to total 
operating time. An increase in total power supplied by the system, and 
in increased redundancy requirements, to maintain the same level of 
reliability, are reflected in added weight as the number of passes increase. 

4.8.5 Control System Weight 

Reaction control system weight for the different Tug configurations as a 
function of the number of passes is given in Figure 4. 8. 5. 0-1. The total 
control system weight represents the hardware and consumable propellant 
required to maintain attitude control. 

4.8.6 Total Tug Weight 

The total weight of the Space Tug at start of aerobraking as a function 
of the number of passes is presented in Figure 4. 8. 6. 0-1. This total 
was determined by adding the respective aerobraking weight penalties for 
each of the aerobraking kit elements (including RCS fuel) to the baseline 
Tug weight of 9,718 pounds, obtained by uprating the 39,800 pound capacity 
synchronous orbit primary propulsion module of prior Reference 1.1. 0.0-1 
to a 45,000 pound capacity propulsion module. The payload capability is 
the difference between this total and the 14,430 pound maximum allowable 
Tug weight at start of aerobraking based on available delta velocity 
capability. 



FIGURE 4. 8. 4. 0-1 ASTRIONJCS SYSTEM WEIGHT VERSUS NUMBER OF PASSES 

4-257 









REACTION CONTROL SYSTEM WEIGHT (POUNDS) 



FIGURE 4. 





TOTAL TUG WEIGHT AT START AEROBRAKING (LBS. X 10“ 3 ) 


D5-17142 










D5-17142 


5.0 SENSITIVITY ANALYSIS 

The sensitivity analysis was performed to examine various parameters 
peculiar to aerobraking and to provide an overall impact assessment of 
the technical studies. In performing this analysis, data from Section 
4 was utilized to determine the effects of the aerobraking kit inert 
weight penalties, delta velocity effects, reaction control system versus 
main engine operations, circularization altitude and Shuttle rendezvous 
effects. Operating modes were examined to define the payload capa- 
bilities of the four aerobraking Tug configurations (Section 5.1). 

The effect of navigation uncertainties and atmosoheric dispersions 
were defined to determine their impact on aerobraking Tug temperatures, 
materials, navigation accuracy, mission duration, operational modes 
and payload. The need for guidance schemes, better understanding of 
atmospheric variations and additional study of navigation and atmospheric 
effects was identified (Section 5.2). 

This analysis also compared the aerobraked Tuq with the conventional 
trajectory single reusable stage Tug used in the geosynchronous role. 
Equivalent payload Tug sizes were computed and the performance parameter 
sensitivities were compared (Section 5.3). 

The sensitivity analysis (1) indicated methods of increasing the payload 
capability of the aerobrakinq Tug, (2) defined the impact of atmospheric 
dispersions and navigation error on thermal protection requirements, 
mission operating mode and payload capability, and (3) evaluated the 
conventional and aerobraked Tuq performance parameters and payloads. 

From the above analysis, conclusions and recommendations were developed 
to provide direction to technology and Space Tuq follow-on studies 
(Section 5.4). 

5.1 PAYLOAD/AEROBRAKING PERFORMANCE OPTIMIZATION 

The round trip payload geosynchronous mission was the study's baseline 
mission. This mission requires that equal weight payloads be carried 
to and from geosynchronous orbit by the same Tug. Alternate missions 
(subject to cursory examination in the sensitivity analysis only) 
considered payload placement and retrieval. The conventional trajectory 
single reusable stage Tug used a starting point for this study (Figure 
5. 1.0. 0-1), has no payload capability for placement, retrieval or 
round trip missions. 

The aerobraking operational mode was examined in detail for the geo- 
synchronous round trip mission, since this mission imposed the most 
stringent performance requirements on the Tug systems. In this 
mission profile, the Tug is deployed by the Shuttle in 100 NM, 28° 
inclination orbit. The Tug transfers the payload via Hohmann transfer 
to equatorial synchronous orbit with the 28° plane change being made 
at apogee. In synchronous orbit the payload is exchanged for an equal 
weight return payload. At the proper time the Tug and payload are 
placed on the aerobrakinq return ellipse by applying the deorbit and 
28° plane change delta velocity. During the aerobrakinq return the 

.PRECEDING PAGE BLANK NOT FILMED 
5-1 


WEIGHT ESTIMATE (LBS) 



ASTRIOHiCS 

MODULE 




33 ‘5* 


\ 


lh 2 

TANK 


14*0” DIA 





• PROPULSION MODULE 

• PROPELLANT 

• INERTS 

• ENGINE 639 

• PROP/MECH 801 

• THERMO/MICRO 573 

• STRUCTURE 2912 

• CONT. & RESID. 943 

• ASTRIONICS MODULE 

• ELECTRICAL 841 

• AVIONICS 625 

• STRUCTURE 400 

• THERMAL 660 

• PAYLOAD ADAPTER 

TOTAL SPACE TUG WEIGHT 
MASS FRACTION 


RL-10A-3-6 ENGINE 
(NOZZLE RETRACTED) 


45, 000 
5,868 


50, 868 


2,526 


200 

53, 594 
0,840 


FIGURE S. l.O. 0-1: CONVENTIONAL (N0N-AER0BRAKING) SPACE TUG CONFIGURATION 


D5-17142 


1 


D5-17142 


5.1 (Continued) 

apogee of the ellipse is reduced to 270 NM by drag dissipation of the 
orbital kinetic energy on one or more passes throuqh the upper atmos- 
phere. The orbit is circularized at 270 NM for phasing with the Shuttle 
at 100 NM and at the proper time a final Hohmann transfer is made to 
100 NM for rendezvous with the Shuttle. 

Using this mode, the four aerobrakina configurations all have significant 
positive payloads, in the baseline round trip mission. The payload 
capabilities for the two alternate missions (payload placement and 
retrieval) are more than double that of the round trip mission. The 
following subsections discuss the general parametric results (Section 

5.1.1 - not configuration oriented) and the specific payload capability 
assessment (Section 5.1.2 - configuration oriented). 

5.1.1 General Parametric Results (not configuration oriented) 

This subsection discusses the parametric analysis results pertaining 
to the aerobraked geosynchronous mission. This analysis was used to 
provide early visibility of the effects of weight penalties, mission 
delta velocities, and specific impulse on payload capabilities. Al- 
' ternate mission modes and the sensitivity of payloads to these mission 

modes were investigated to determine if the 100 NM departure orbit to 
geosynchronous orbit and back to a 270 NM circularization orbit 
followed by a 100 NM rendezvous with Shuttle sequence described above 
is near optimum. 

5. 1.1.1 Inert Weights and Delta Velocity Effects 

The initial aerobrakinq delta velocity budqet utilized in the sensitivity 
analysis was 14,100 ft/sec outbound and 8,000 ft/sec return (total 
mission budget of 22,100 ft/sec). With this velocity budget and the 
baseline Tug coni iguration shown in Figure 5. 1.0. 0-1, the aerobraking 
Tug round trip payload capability was computed. Figure 5. 1.1. 1-1 para- 
metrically illustrates the payload capability as a function of aerobraking 
inert weight penalties. Assuming no aerobraking kit inert weight penalty, 
the maximum round trip payload would be approximately 4450 pounds. 

Because the round trip mission is, to a first order approximation, a 
direct one-to-one substitution of inert weight for payload, the payload 
is zero when the aerobraking penalty is 4450 pounds. The vertical bars 
on Figure 5. 1.1. 1-1 illustrate the percentages of the 287 synchronous 
missions that can be captured (accomplished). For example, with an 
aerobraking kit inert weight penalty of 1450 pounds and the associated 
3000 pound payload capability, approximately 95% of all geosynchronous 
missions can be captured. The remaining 5% of the geosynchronous pay- 
loads in the mission model are heavy (7000-10,000 pounds) and are beyond 
the round trip payload capability of the aerobraked Tug, (Note: This 

payload may be accomplished in a placement or retrieval mode). There- 
fore, a 3000 pound round trip capability for the aerobraked Tuq appears 
to be a desirable design goal. 




5-3 



AWiTO TUG FOR AEROBRAKiNG MODIFICATIONS (LBS) 



WEIGHT OF ROUND TRIP PAYLOAD (LBS) 


FIGURE 5.1. 1. 1-1, EFFECT OF AEROBRAKING MODIFICATION WEIGHTS ON ROUND TRIP PAYLOAD 


D5-17142 


5. 1.1.1 (Continued) 

Figure 5. 1.1. 1-2 extends the parametric investigation beyond that shown 
in Figure 5. 1.1. 1-1 above. The effects of both aerobraking inert weight 
penalties and return delta velocities are depicted. The return delta 
velocity of 8000 ft/sec with zero inert weight penalty results In a 
4450 pound payload. However, if the return delta velocity were 7500 
ft/sec (zero inert weight penalty), the payload capability is increased 
to 5000 pounds. 

Linear interpolation can be used vertically (along equal return delta 
velocity lines) in Figure 5. 1.1. 1-2. The lines across the figure labeled 
with the inert weight penalties are not straight lines because of the 
natural logarthims involved. The slope of these lines varies from 
approximately 1.07 to 1.23 lbs/ft per second for the 8500-3000 and 
7000-7500 ft/sec bands, respectively and indicates the payload sensi- 
tivity of the aerobraked Tuq to changes in delta velocity. 

As the delta velocity is dependent on the operational Shuttle/Tuq mode 
used, Figure 5. 1.1. 1-2 provides insight to the effect on payload capa- 
bility of the selected operating mode's delta velocity. For example, 
Figure 5. 1.1. 1-2 can be used to estimate round trip payload capabilities 
for various total mission delta velocities. To do this, add the 14,100 
ft/sac outbound and a return velocity (e.q., 8000 ft/sec) and compare 
with the computed total mission delta velocity. The difference between 
these two values is used as a delta to move right or left from the 
same return velocity selected on the chart. Usinq the inert weight 
difference (penalty) associated with a particular configuration, the 
resulting payload can be estimated. An example of this is shown below: 


Configuration A (Computed) Figure 5, 1.1, 1-2 Values 


Outbound 4V (ft/sec) 14,500 
Return 4V (ft/sec) 8,100 
Total Mission 4V (ft/sec) 22,600 
Total Inert Weight (Lbs) 9,594 


14,100 

8,000 

8,594 


Difference in total mission 4V = 22,600 - 22,100 = +500 ft/sec 
Difference in total inert weight = 9594 - 8594 = 1000 lbs ( 4 Wt) 

Read up vertically from 8500 ft/sec (8000 + 500) to the 4Wj = 1000 line 
The estimated round trip payload is 2900 pounds 








D5-17142 


5. 1.1.1 (Continued) 

Figure 5. 1.1. 1-3 and -4 show the same type of data as Figure 5. 1.1. 1-2 
for the two alternate missions, payload placement and retrieval, respectively. 
The interpretation and use of these two figures is similar to that of 
Figure 5. 1.1. 1-2. The maximum payload placement capability is approxi- 
mately 9600 pounds (8000 ft/sec return delta velocity and zero inert 
weight penalty). With the same 8uGQ ft/sec return delta velocity and 
2000 pounds inert weight penalty, approximately 5300 pounds of payload 
can be placed. Payload capabilities for the retrieval mission using the 
same velocity budget and aerobraking kit weight penalties are 8300 and 
4600 pounds of payload, respectively. The retrieval mission, shown in 
Figure 5. 1.1. 1-4, is more sensitive to delta velocity than the placement 
mission but for return delta velocities of 7000-7200 ft/sec, the place- 
ment and retrieval payload capabilities are essentially equivalent. 

5. 1.1. 2 RCS Isp Effects 

The Tug's GO 2 /GH 2 RCS system has an effective specific impulse (Isp) of 400 
seconds compared to 460 seconds for the main RL-10A-3-8 engine (prior 
Reference 1.1. 0.0-1). Because of this lower Isp value, there will be a 
payload reduction associated with extensive use of the RCS system during 
the aerobraking phase of the mission. In addition to the normal RCS 
functions of limit cycle operation and stabilization as discussed in 
Section 4.4 (Control Analysis), the RCS might substitute for the main 
engine in the final low earth circularization, phasing and rendezvous 
maneuvers . 

Figure 5. 1.1. 2-1 shows the effect on entry weight by using the RCS to pro- 
vide all of the propulsive delta velocity requirements after the completion 
of the aerobraking phase. The change in Tug gross weight at the beginning 
of the aerobraking phase is shown on the ordinate. The abscissa represents 
a range of possible total mission propulsive delta velocity requirements. 

The trajectory analysis (Section 4.3) utilized a baseline delta velocity 
budget of 22,400 ft/sec. For this delta velocity value, the delta entry 
weight (using the main engine) is zero. With this same total budget 
(22,400 ft/sec) and using the RCS to provide all propulsive requirements 
after aerobraking, the Tug’s gross entry weight is decreased approximately 
75 pounds. The entry weight difference between the main engine and RCS 
curves is less for smaller velocity budgets because of the decreased 
velocity requirement imposed on the lower Isp RCS. Similarly, the entry 
weight difference is greater for the larger velocity budgets because of 
the increased demands on the RCS. The additional propellant required by 
the lower Isp RCS is included in the entry weights shown. Therefore, the 
entry weight deltas do not represent completely the payload penalties 
associated with using the RCS as a substitute for the main engine. These 
payload penalties are further discussed in the following paragraphs. 

Figure 5. 1,1. 2-2 extends the parametric analysis of Figure 5. 1.1. 2-1 
above. The effects of using the RCS for all propulsive maneuvers after 
the completion of aerobraking are plotted as a function of initial Tug 
and payload gross weight at 100 NM. The weight differences between the 


5-7 




c n 







DELTA ENTRY WEIGHT - LBS 


• ROUND TRIP PAYLOAD 
FROM 100 NN 

• FIXED PROPELLANT 
UT 9 45000 LBS 


w; ; 


MAIN EN6INE USED 
FOR ALL AV AT 
END OF AEROBRAKING 


RCS USED FOR ALL 
AV AT END OF 
AEROBRAKING 


I uiHBniHSSstrasaias 

BUS 


i&ssp 


JUKI 


22000 ! 


. 23000 


FIGURE 5. 1.1. 2-1 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS 
ORBIT - ATMOSPHERIC ENTRY WEIGHT 

J 

■ s 

r - - \ — ' J 
i : i 

! 

j 

! 
























D5-17142 






D5-17142 


5. 1.1. 2 (Continued) 

main engine and RCS curves are identical to those of Figure 5. 1.1. 2-1. 

Also shown in Figure 5. 1.1. 2-2 is the effect of offloading propellant to 
maintain a fixed initial Tug and payload gross weight (57*740 pounds). 

With this fixed initial weight, the aerobraked Tug's maximum delta 
velocity capability is 22,400 ft/sec (main engine). Using the RCS for 
all propulsive maneuvers after the completion of aerobraking reduces 
the maximum capability to 22,170 ft/sec. Neither propellant loading 
technique (full nor offloading) depicted on the figure permits an initial 
Tug and payload gross weight greater than 59,500 pounds for the expected 
aerobraking delta velocity budget. 

If the RCS were used to provide the propulsive delta velocity after aero- 
braking, there will be a reduction in round trip payload because of the 
lower specific impulse. This payload reduction (total mission delta 
velocity budget = 22,400 ft/sec) is approximately 300 pounds for a fixed 
initial propellant weight of 45,000 pounds. The maximum delta velocity 
capability is exceeded for the fixed initial weight approach. However, 
if this were not true, and to provide a comparative value, the payload 
reduction would be approximately 230 pounds. 

The two approaches shown in Figure 5. 1.1. 2-2 (fixed initial propellant 
weight or fixed initial qross weiqht) were compared to determine the 
preferable operational mode. Fixing propellant weight is the most loqical 
approach to take since using all available propellant for a fixed stage 
size is more efficient than' off loadinq propellant. But, fixing propel- 
lant causes a change in entry weight which in turn causes a chanqe in 
inert weight (due to heating and loads). These changes cannot be 
evaluated directly from the results of the technical studies (Section 4) 
since the trajectory, thermal, control and weights analyses were per- 
formed for a fixed entry weiqht. However, as will be shown in Section 

5.1.2 below, the actual changes in entry weights are small. Therefore, 
any inert weight penalties would be very small and have an insignificant 
impact on the payload capabilities of the aerobraked Tug configurations. 

Figure 5. 1.1. 2-3 shows the round trip payload penalties associated with 
using the RCS for various percentages of the return trip delta velocities. 
The notes on Figure 5. 1.1. 2-3 refer to Figures 5. 1.1. 1-1 and -2. For 
example, if the RCS accomplished 15% of the 8000 ft/sec return delta 
velocity (1200 ft/sec), the round trip payload penalty would be approxi- 
mately 200 pounds. The payload penalties shown in this figure were used 
iri the configuration oriented payload computations discussed in subsequent 
subsections. 

5. 1.1.3 Circularization Altitude and EOS Rendezvous Effects 

One of the objectives of the sensitivity analysis was to investigate 
alternate mission modes that could enhance the payload potential of the 
aerobraking technique. As seen in prior Figure 5. 1.1. 1-2, the mission 
delta velocity budget is a critical payload parameter. Figure 5. 1.1. 3-1 
shows the reductions in total mission delta velocity that can be obtained 


5-12 



PERCENT OF RETURN <4V ACCOMPLISHED BY RCS 



D5-17142 







FIGURE 5. 1.1. 3-7 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS 

ORBIT - MISSION TOTAL DELTA VELOCITY SENSITIVITY 
TO FINAL ORBIT ALTITUDE 


ZHL i-sa 














D5-17142 


5. 1.1. 3 (Continued) 

by using alternate Tug/Shuttle mission profiles for rendezvous. The top 
dotted curve on the figure illustrates the nominal Tug circularization 
at 270 NM followed by a propulsive transfer to 100 NM for Shuttle ren- 
dezvous (see prior Figure 1.2. 0.0-1). The mission delta velocity budget 
for this mode could vary from 20,150 to 22,400 ft/sec dependent upon 
the rendezvous, docking, and reserve requirements (see notes on Figure 
5. 1.1. 3-1). If a circularization altitude of 200 NM, followed by a transfer 
to 100 NM were selected, the delta velocity requirement is reduced 
approximately 360 ft/sec. The mission modes shown by the solid lines 
eliminate the transfers to 100 NM. The Tug circularizes at the altitude 
shown and awaits the Shuttle in that orbit. Significant decreases in 
the mission delta velocity budget are obtainable by using these direct 
rendezvous modes . 

The trajectories for the alternate mission modes shown on Figure 5. 1.1. 3-1 
were generated by varying the first pass vacuum perigee so that the final 
pass apogee decayed to the desired altitude. The maximum variation in 
the first pass perigee was approximately 230 feet to change the final pass 
apogee from 270 NM to 100 NM. 

Figure 5. 1.1. 3-2 shows the results of the analysis of an alternate aero- 
braking mission mode. The mode depicted has the following sequence: (1) the 
EOS ejects the Tug and payload in 100 NM/28.5 0 circular orbit; (2) the Tug 
delivers the payload into geosynchronous orbit, picks up another payload 
and deorbits; (3) the aerobraking phase is completed with a circularization 
burn into a variable altitude shown on the ordinate of Figure 5. 1.1. 3-2; 
and (4) the Tug transfers to 100 NM/28.5 0 to meet the Shuttle for return 
to earth. 

Figure 5.1. 1.3-2 shows the effects of (1) circularization altitude selection, 
(2) jettisoning part of the aerobraking modification kit, and (3) use of 
the RCS to perform the circularization and transfer burns. For tirs 
example, an aerobraking kit inert weiqht penalty of 1000 pounds was 
assumed (800 pounds assigned to propulsion module and 200 pounds assigned 
to the astrionics module). Referring to the top hatched area, with this 
inert weight penalty and an 8000 ft/sec return delta velocity, the basic 
payload would be 3450 pounds using the main engine to circularize at 270 
NM and then to transfer the Tug from 270 to 100 NM. If the propulsion 
module aerobraking kit (800 pounds) were jettisoned immediately after the 
last atmospheric passage, the payload capability would increase about 100 
pounds. Circularizing at higher altitudes (above 270 NM) with the main 
engine decreases the payload because of the higher propulsive require- 
ments to circularize at the higher altitude and then to transfer to 100 
NM. The impact on payload capability by jettisoning the propulsion module 
aerobraking kit at the higher circularization altitude is also small 
(^125 pounds) for a circularization altitude of 400 NM. 

Circularizing at altitudes below 270 NM increases the payload capability 
because of the lower circularization propulsive requirements. The payload 
capability effect of retaining or jettisoning the aerobraking kit 


5-15 




05-17142 








D5-17142 


5, 1.1.3 (Continued) 

modifications becomes even less siqnificant than the effect at higher 
altitudes. Circularizing at 200 NM rather than 270 NM will increase 
payload capability approximately 500 pounds. 

The bottom hatched area presents similar data but utilizes the RCS as 
a substitute for the main engine to perform all return trip low earth 
burns. The same trends noted for the main engine are shown. The payload 
penalty using the RCS as a substitute for the nominal 270-100 NM burns 
is approximately 250 pounds. 

Figure 5. 1.1. 3-3 has the same mission mode sequence as Figure 5. 1.1. 3-2 
above but assumes a more severe inert weight penalty for aerobraking 
kit modifications. The propulsion module aerobraking kit was assumed to 
weigh 2400 pounds, This case was included to investigate the effect of 
jettisoning a heavier aerobraking kit (e.g., containing a large flare). 

Using the main engine and with the nominal post aerobrakinq operations, 
i.e., 270 ->100 NM transfer, approximately 225 pounds of payload increase 
can be obtained by jettisoning 2400 pounds of modifications. This payload 
capability increase varies from negligible (circularizing at 100 NM) to 
300 pounds (circularizing at 400 NM). 

Figure 5. 1.1. 3-4 examines the effect on round trip payload by using the 
Shuttle to accomplish more of the low earth orbit (circularization, phasing, 
rendezvous and docking) propulsive requirements. The mission mode (shown 
by the solid lines) includes the nominal 100 NM/28.5 0 departure, geo- 
synchronous orbit, deorbit, and aerobraking phase. The Tug then 
circularizes at any altitude shown on the ordinate. If this circularization 
altitude is between 100 NM and 270 NM the Shuttle recovers the Tuq in 
this circular orbit, thereby eliminating the requirement for any Tuq 
transfer burn. If the Tuq circularized above 270 NM, the Tug then 
transfers to 270 NM and is recovered by the Shuttle, thereby reducinq the 
Tug's propulsive transfer requirement. The dotted lines on Fiqure 
5. 1.1. 3-4 represent the mode (transfer to 100 NM for EOS recovery) shown 
in prior Figures 5. 1.1. 3-2 and -3 and are shown for comparative analysis. 

Comparing the Tug payload capabilities of the two operational modes at a 
circularization altitude of 270 NM, an increase of approximately 700 
pounds is gained by EOS/Tug rendezvous at 270 NM over the payload capa- 
bility of the Tug transferring down to 100 NM. With a Shuttle/Tug rendezvous 
altitude of 200 NM, approximately 1000 additional pounds of payload can be 
achieved. This 1000 pound increase is 500 pounds more than is achieved 
by the Tug circularizing at 200 NM ? n d cransferring to 100 NM. These 
operational options are indicative of the many modes available and 
demonstrate the major impact that the circularization and rendezvous 
operation modes have on aerobraking payload capability. The payload 
performance associated with the use of the RCS as a substitute for the 
main engine in the return low earth burns are also shown. 


5-17 


ALTITUDE OF CIRCULARIZATION (N. M. ) 








ALTITUDE OF CIRCULARIZATION (N. M. ) 


(a) 100 <h <270 N.M., 

CIRCULARIZE & AWAIT EOS 

(b) h > 270, CIRCULARIZE & 
TRANSFER TO 270 N.M. FOR EOS 

CIRCULARIZE & TRANSFER 



^WEIGHT OF PAYLOAD (LBS) 

FIGURE 5. 1.1. 3-4. EFFECT OF TUG RECOVERY ALTITUDE AND RECOVERY METHOD 

ON DELTA PAYLOAD 


D5-17142 



05-17142 


5.1 .1.4 General Parametric Conclusions 

From the data presented, three conclusions were drawn. 

o Payload capability would be significantly increased by de- 
creasing the nominal circularization altitude from 270 NM 
to approximately 200 NM. Very low circularization altitudes 
(e.g., 100 NM) have certain disadvantages includinq: (1) Or- 

bital decay if the Tug must wait some significant period of 
time for EOS recovery; and (2) the lack of an altitude 
tolerance band in which to circularize after repeated 
passages through an unpredictable atmosphere. 

o Jettisoning the propulsion module aerobraking kit to gain 
payload should not be considered as the aerobraked Tug's 
baseline operational mode. The payload increase does not 
appear sufficiently significant ( a; 250 pounds payload at 200 
NM for 2400 pounds inerts) to justify the normal expenditure 
of the kit for this reason alone. 

o The combination Shuttle/Tug capability should be assessed to 
determine not only optimum recovery but also optimum 
departure orbits. 

5.1.2 Specific Payload Capability Assessment (Configuration Oriented) 

The four aerobraking configurations selected for analysis during this study 
are discussed in Section 4.2 and are shown in Figure 5. 1.2. 0-1. The fully 
fueled weight statements (standard atmosphere and no navigation errors) 
for these four configurations have been extracted from Section 4.8 and 
are shown in Figure 5. 1.2. 0-2. The inert weights of the basic propulsion 
module (5868 pounds) and astrionics module (1960 pounds) are summed in 
the column labeled Tug Inert Weight. 

The maximum gross Tug weight (less payload) shown in Figure 5. 1.2. 0-2 is 
56,685 pounds for the 5 pass, 45° flare configuration. Using prior Figure 
5. 1.1. 1-2, with an effective delta inert weight of 3091 pounds, a maximum 
round trip payload of 1000-2000 pounds could be expected. The total gross 
payload for the Shuttle would be approximately 58,500 pounds. Similarly, 
the minimum gross Tug weight of 54,643 pounds for the 30 pass basic (no 
flare) configuration would have an estimated maximum 3000-4000 pounds of 
payload for a total gross Shuttle payioad of 58,500 pounds. Both of these 
gross weights are less than the Shuttle's 100 NM/28.5 0 payload capability 
of 65,000 pounds shown in Figure 5. 1.2. 0-3 (Reference 5. 1.2. 0-1). 

From the general parametric studies of Section 5.1.1 above* a Tug cir- 
cularization and recovery altitude of approximately 200 NM appeared 
attractive. Figure 5. 1.2. 0-3 indicates that a representative Shuttle 
payload capability to a 200 NM/28.5 0 orbit would be 60,000 pounds. There- 
fore, both operational modes were carried forward for further analysis. 
Within this subsection, the only exception to this is the payload placement 
mission where the Tug with payload has a gross weight greater than the 


5-20 


BASIC 

3CT FLARE 

45 FLARE 

(NO FLARE) 

r- PAYLOAD ADAPT. 



; PAYLOAD 

g i 

c t -p- — 

! t — — 

V ASTRIONICS 

r7'"% 




PAYLOAD | / 

FLARED / 
SKIRT A 

ASTRiONICS | 3 ° B 


PAYLOAD l 


ASTRIONICS 45 ° 


PAYLOAD 

ASTRIONICS 


S ?5 




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AFT HEAT SHIELD- 


FIGURE 5. 1.2. 0-1: SELECTED SPACE TUG AEROBRAKING CONFIGURATION CONCEPTS (1ST PHASE) 



5-22 


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* STANDARD ATMOSPHERE AND NO NAVIGATION ERRORS 
** INCLUDES PAYLOAD INSULATION AND PAYLOAD CAP ACTUATION DEVICE 


FIGURE 5. 1.2. 0-2 FULLY FUELED WEIGHT STATEMENT 







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FIGURE 5.1. 2.0-3: H-33 CONFIGURATION EXTERNAL H2 TANK ORBITER 

ALTERNATE MISSION CAPABILITY 


D5-17142 



D5-17142 




5.1.2 (Continued) 

Shuttle can deliver to 200 NM. For identification, the mode of departinq 
from 100 NM, circularizing at 270 NM and transferring back to 100 NM, is 
denoted by "270-100 NM recovery". The 200 NM mode is denoted by "Depart 
and Recover at 200 NM" . 

The following subsections discuss the mission delta velocity budgets and 
the payload capabilities of the four aerobraked configurations under the 
standard atmosphere and zero navigation error conditions. The sensitivity 
of the payload capabilities to atmospheric perturbations and navigation 
errors is contained in Section 5.2. 

5. 1.2.1 Delta Velocity Budgets 

The total mission delta velocity budqet for an aerobraked Tug configuration 
is the sum of (1) the normal main engine and RCS propulsive requirements for 
the non-aerobraking phases of the geosynchronous mission, and (2) the added 
RCS requirements for stability and control during the aerobraking phase. This 
subsection discusses both of these requirements in order to establish a total 
mission delta velocity budqet for the payload computations to follow. The 
first two figures show the non-aerobraked propulsive delta velocity requirements 
for the four configurations and establish the operational mode dependent fixed 
base requirements. The last figure in this subsection shows the added aero- 
braking phase RCS delta velocity requirements and the total mission budqet. 

Figure 5. 1.2. 1-1 shows the sensitivity of the delta velocity budget to 
(1) the descent olane change angle* (2) the number of passes to lower the 
apogee to 270 NM, and (3) the aerobraked Tug configuration. Increasing the 
plane change requirement from the nominal 28° to 30° results in a delta 
velocity increase of 150 ft/sec. Similarly, a decrease in the plane change 
requirements of 2° to 26° results in a delta velocity decrease of 150 ft/sec. 

The sensitivity of delta velocity to the number of passes is relatively in- 
significant and is the result of varying the first pass perigee (see Para- 
graph 5. 1 .1 .3 above) . 

All four of the aerobraked configurations have nearly identical delta ve- 
locity budgets (fixed plane change angle). The delta velocity budqet 
difference between the basic (no flare) and the 60° flare configurations 
is approximately 13 ft/sec for all mission durations. The maximum 
difference in delta velocities (fixed plane change angle) is approximately 
38 ft/sec (5 pass basic no flare to 60 pass 60° flare). These small dif- 
ferences can be absorbed within a rendezvous and reserve delta velocity budget 
of 800 ft/sec and are not considered payload significant. 

Fiqure 5. 1.2. 1-2 illustrates the sensitivity of the 30° flared configuration's 
burnout weight (inerts plus payload) to the mission mode selected. The delta 
velocities shown on the abscissa include 1200 ft/sec for rendezvous and 
reserves. The delta velocity (using the main enqine for post aerobraking 
propulsion) of 22,400 ft/sec (see prior Fiqure 5. 1.1. 2-2) has a zero delta 
burnout weight. This delta velocity, with the additional velocity allowan- 
ces made, approximates the 270-100 NM mode (labeled "D"). Other mission 


5-24 


* a ItiD. LtLUs.a 


KCUFTCU \ csscn CO. 









ROUND TRIP AV 
AV INCLUDES 

800 EPS FOR RENDEZVOUS 
and RESERVE 
FINAL PASS APOGEE ■ 

ALT * 270 NM 

FINAL CIRCULAR ORBIT 

ALT = 100 NM 

BASED ON TRAJECTORIES 

IN THE HOHINAL ATMOSPHERE 






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+ i i t p 

PASSES TO LOWER APOGEE TO 270 MM ‘ 




SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS 
ORBIT - MISSION TOTAL DELTA VELOCITY SENSITIVITY 
JO PLANE CHANGE ANGLE AND CONFIGURATION 






HfHfiHH 

mS 
















WEIGHT 


f-1200 FT/SEC FOR 
!_ REND. & DOCKING 


v ROUND TRIP PAYLOAD 
FROM 100 NM 
• PLANE CHANGE * 28° 


FIXED PROPELLANT (45,000 LBS) 
(VARIABLE INITIAL WT & ENTRY WT) 


FIXED INITIAL WT (57,740 LBS) 

(OFFLOAD PROPELLANT @ ENTRY WT ’ 14,430 LBS) 


MAIN ENGINE 
USED FOR ALL 
AV AT END OF 
AEROBRAKING 
(I SP = '460 SEC) 
RCS USED FOR 
ALL AV AT END 
OF AEROBRAKING 
(Ic„ = 400 SEC) 




AV RANGE FOR 30° FLARE 
5 TO 60 PASSES 


FINAL PASS 

FINAL CIRC 

APOGEE ALT 

ORGIT ALT 

TOO NM 

100 NH 

200 

200 

200 

100 

270 

100 


baseline of s - 
sec 4.3 — — 


! 22500 


MISSION TOTAL AV - FT/SEC 


FIGURE 5.1 .2.1-2 SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS , 
ORBIT - PAYLOAD SENSITIVITY TO DELTA VELOCITY 


05-17142 






















D5-17142 


5. 1.2.1 (Continued) 

modes, all departing 100 NM, are similarly labeled. Also shown on Figure 
5. 1.2. 1-2 are the effects of using the RCS to provide the post aerobraking 
propulsive requirements. 

The data for the 30° flare configuration shown on Figure 5. 1.2. 1-2 confirms 
the general parametric conclusions of Section 5.1.1 above. For example, 
the 30° flare configuration's delta velocity budget can be reduced approxi- 
mately 900 ft/sec by returning direct to 100 NM rather than usina the 
270-100 NM mode. This results in a round trip payload increase of 810 pounds 
(fixed initial or entry weight) or about 1000 pounds (fixed 45,000 pounds 
initial propel lant weight) . Without consideration of the RCS control re- 
quirements, all of the four aerobraked configurations will have approximately 
the same payload increase because the delta velocity budqets are nearly 
identical . 


Section 4.4 contains data on the RCS propellant consumed by the various 
configurations durinq the aerobrakinq phase of the mission. This RCS 
propellant can be converted to an equivalent delta velocity by the re- 
lationship shown below: 


( 1 ) 


Weight of Propellant (W ) = 

r 


Total Impulse (I) 
Specific Impulse (Isp) 


(2) I = (mass) x (delta velocity) 


These equivalent delta velocities were added to the nominal mission re- 
quirements. Also included was 400 feet per second for midcourse corrections 
due to navigational errors during the aerobrakinq phase. The resulting 
overall mission delta velocity budgets are shown in Fiqure 5. 1.2. 1-3. The 
top portion of this figure reflects the equivalent delta velocity required 
by the RCS during the aerobraking phase and are configuration dependent. 

The fixed base totals shown at the bottom of Figure 5. 1.2. 1-3 are opera- 
tional mode dependent. Only one fixed base value was utilized for all 
configurations because of the insensitivity of configuration to delta 
velocity as shown in Figure 5. 1.2. 1-1. The total mission delta velocity 
for a specific configuration and number of passes and for an operational 
mode is the sum of the two values. For example, the basic (no flare) 30 
pass configuration operating in the 270-100 NM transfer mode has a total 
mission delta velocity of 22,500 (21,950 + 550) ft/sec. The 60° flare 30 
pass configuration in the same operational mode has a total mission delta 
velocity of 22,020 ft/sec. This decrease in total mission delta velocity 
for the flared configurations tends to compensate for their qreater inert 
weights shown in the prior Fiaure 5, 1.2. 0-2. 

5. 1.2. 2 Basic (No Flare) Configuration 

Figure 5, 1.2. 2-1 shows the qross weiqht distribution of the basic (no flare) 
Tug configuration and its round trip payload at the start of the aerobrakina 
phase (i.e., at the time of first atmospheric entry) using the 270-100 NM 
recovery mode. The Tuq inerts are those shown previously in Figure 
5. 1.2. 0-2. The aerobrakinq penalties are the sums of the inert weiqht 


5-27 



REQUIRED ADDITIONAL MISSION DELTA VELOCITY 
(FEET PER SECOND) 


05-17142 



0 10 20 3 0 40 50 60 

NUMBER OF PASSES 


FIXED BASE TOTAL MISSION DELTA VELOCITY 



270-100 N.M. 
TRANSFER MODE 

200 N.M. DEPARTURE 
& RECOVERY MODE 

TO SYNC. ORBIT 

14,100 

13,810 

DOCKING @ ORBIT 

400 

400 

DEORBIT 

5,993 

5,993 

CIRCULARIZATION(S) 

1,057 

307 

& TRANSFER 


RESERVE FOR 

400 

400 

Maneuvers 


FIXED BASE TOTAL i 

21,950 FT/SEC 

20,910 FT/SEC 


FIGURE 5,1 .2.1-3, TOTAL MISSION DELTA VELOCITY EQUIVALENT VS. 

NUMBER OF PASSES 

5-28 



TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF POUNDS) 


D5-17142 



FIGURE 5.1. 2. 2-1: AEROBRAKING WEIGHT VERSUS NUMBER OF PASSES 

(BASIC - 270-100 N.M. RECOVERY) 




5-29 


f 


D5-17142 


5. 1.2. 2 (Continued) 

penalties (Figure 5. 1.2. 0-2) and the penalty of using the RCS to provide 
part of the equivalent mission delta velocity (Figure 5. 1.1. 2-3). Since 
the main engine was utilized to circularize at 270 NM and to transfer from 
270 to 100 NM, this RCS Isp penalty was relatively small (a/ 200 pounds of 
payload). The propellant remaining for the main engine and the RCS is that 
usable propellant required to perform the remainder of the mission during 
and after the aerobraking phase. 

As the number of passes in the mission increases, a small increase in the 
remaining propellant can be noted. This is caused by the increased equiva- 
lent total mission delta velocity shown in Figure 5. 1.2. 1-3. The bucket in 
the aerobraking penalty curve (ai30 passes) represents that shown in 
tabular form in Figure 5. 1.2. 0-2. Because of the almost insignificant slope 
of the gross weight line, this aerobraking penalty bucket also represents 
the point at which the maximized round trip payload for the basic (no flare) 
Tug will occur. 

Figure 5. 1.2. 2-2 shows similar data to that in Figure 5. 1.2. 2-1 but uses 
the operational mode of EOS delivery and recovery at 200 NM. As shown in 
Figure 5. 1.2. 1-3, the mission delta velocity for this operational mode is 
decreased approximately 1000 ft/sec from that of the 270-100 NM mode. The 
same trends are noted with this operational mode as with 270-100 NM mode. 

The round trip payload maximizes for missions having approximately 30 
passes. The basic (no flare) configuration has approximately 1200 pounds 
more round trip payload capability in the 200 NM mode (^45# more) than in 
the 270-100 NM mode (3950 pounds versus 2750 pounds). 

5. 1.2. 3 30° Flare Configuration 

The gross weight distribution at the start of aerobraking (270-100 NM 
mode) is shown in Figure 5. 1.2. 3-1. The rationale of the data shown is 
similar to that discussed for the basic (no flare) configuration in Figure 
5. 1.2. 2-1 above. The 30° flare (as well as the 45° and 60° flares to 
follow) has a smaller total mission equivalent delta velocity than the basic 
(no flare) configuration. The RCS Isp penalty is also smaller and has 
lesser impact on the total aerobrakina penalty. The bucket in the aero- 
braking penalty curve is not sharp or well defined. Therefore, the round 
trip payload capability is nearly constant for missions havinq 20-40 passes 
(3.7 to 7.4 days). 

Figure 5. 1.2. 3-2 shows similar data for the 200 NM mode. The same trends 
as discussed in Figure 5. 1.2. 3-1 above are observed. In particular, the same 
relatively insensitivity of payload capability for missions of 20-40 passes 
is seen. The round trip payload capability for the 200 NM mode is 
approximately 1275 pounds greater than for the 270-100 NM mode. This 
increase in payload capability is slightly greater than for the basic (no 
flare) configuration because of the impact of the decreased delta velocity 
offsets the increased inert weights. 


5-30 


TUG WEIGHT AT START OF AEROBRAK'.NG (THOUSANDS OF POUNDS) 


( 


D5- 17142 



NUMBER OF PASSES 

FIGURE 5. 1.2. 2-2: AEROBRAKING WEIGHT VERSUS NUMBER OF PASSES 

(BASIC - DEPART AND RECOVER AT 200 N.M.) 




5-31 





TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF POUNDS}. 


I 


D5-1 7142 





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0 10 20 30 40 50 60 


NUMBER OF PASSES 

FIGURE 5. 1.2, 3-1: AEROBRAKING WEIGHT VERSUS NUMBER OF PASSES 

(30° FLARE - 270-100 N.M. RECOVERY) 


5-32 






TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF POUNDS) 


D5-17142 


GROSS WEIGHT p START OF AEROBRAKING 




ROUND TRIP PAYLOAD 


AEROBRAKING PENALTIES 

• HEAT SHIELD 

• SIDEWALL INSULATION 

• PAYLOAD & FLARE ADAPTER 
•FLARE 

* RCS Up 

• AM MODS I I | 


PROPELLANT REMAINING FOR 
.MAIN ENGINE & RCS _ 


TUG INERTS = 7,828 LBS. 

• PM -5,868 

~ • AM = 1,960 

I ! 


NUMBER OF PASSES 


FIGURE 5.1 .2*3-2: AEROBRAKING WEIGHT VERSUS NUMBER OF PASSES 

(30° FLARE - DEPART AND RECOVER AT 200 N.M. ) 




D5-17142 


5. 1.2. 4 45° Flare Configuration 

The 270-100 NM mode 45° flare configuration qross weight distribution is 
shown in Figure 5. 1.2. 4-1. This configuration has a relatively heavy flare 
and heat shield for a 5 pass mission (see Figure 5. 1.2. 0-2). The 10 pass 
mission aerobraking inert weights significantly decreased and this fact is 
reflected in the sharp drop in the aerobraking penalty curve between 5 and 
10 passes. Similar to the 30° f'iare configuration, the aerobraking penalty 
bucket is not sharp. Therefore, the round trip payload is relatively insen- 
sitive to trip times over 2 days (10 passes). 

Figure 5. 1.2. 4-2 shows the 45° flare configuration in the 200 NM mode. The 
characteristics of the 45° flare curves are similar to the 30° flare 
configuration and the 200 NM mode provides the same increase in payload 
capability as did the 30° flare. 

5. 1.2. 5 60° Flare Configuration 

The 60° flare configuration in the 270-100 NM mode is shown in Fiqure 
S. 1.2. 5-1. The major flare weight difference between the five and ten 
pass missions (Figure 5. 1.2. 0-2) has the greatest impact on the decrease 
in aerobraking penalty weights shown in that region. Similar to the other 
flares, the 60° flare configuration is relatively insensitive to mission 
duration (after about 10 passes or 2 days). 

The 60° flare data for the 200 NM mode is shown in Fiaure 5. 1.2. 5-2. The 
same relative insensitivity of payload to mission time (missions of at 
least 10 passes) is demonstrated. The 200 NM mode provides about 1275 
pounds of additional payload capability (approximately 50%) over that of 
the 270-100 NM mode (3875 pounds versus 2600 pounds). 

5. 1.2. 6 Configuration Payload Comparison 

Figures 5. 1.2, 6-1 and -2 compare the round trip payload capabilities of 
the four configurations in the 270-100 NM and 200 NM modes respectively. 

The 30° flare configuration has the greatest payload capability in both 
modes. Its maximum round trip payload capability in the 270-100 NM mode 
is 2950 pounds and 4225 pounds in 200 NM mode. (Note; The 45° and 60° 
flare payloads are comparable to the 30° flare payloads if flare 
lengths/weights are equivalent.) 

The shape of the curves is similar in both figures but the reduction in 
mission delta velocity by using the 200 NM mode has affected the flare 
configuration more than the basic (no flarej configuration. This is shown 
by the difference in relative spacing among the curves on the two figures. 
The small payload difference between the 45 e and 60° flares (25-60 passes) 
is due to the small difference in RCS requirements (equivalent delta 
velocity) because the total inert weights of these two configurations are 
almost identical. 



TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF POUNDS) 


05-17142 



FIGURE 5.1. 2.4-1: AEROBRAKING WEIGHT VERSUS NUMBER OF PASSES 

(45° FLARE - 270-100 N.M. RECOVERY) 


5-35 


TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF ROUNDS} 




• MAIN ENGINE PERFORMS 
CIRCULARIZATION 

i t i t 


GROSS WEIGHT @ START OF AEROBRAKING 

i ill ill 


ROUND TRIP PAYLO 


AEROBRAKING PENALTIES _ 

• HEAT SHIELD 

• SIDEWALL INSULATION 

• PAYLOAD & FLARE ADAPTER 

• FLARE 

• RCS l sp 

• AM MODS 


PROPELL 
MAIN ENG 


TUG INERTS « 7,828 

• PM = 5,868 
— . ® AM = 1,960 


HB 

■j ScJ 

mm\ 


FIGURE 5. 1.2. 4-2: AEROBRAI 

(45° F 






TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF POUNDS) 


I I I I I I 

GROSS WEIGHT § START OF AEROBRAKING 




ROUND TRIP PAYLO, 


IB! IB 


AEROBRAKING PENALTIES ~ 

• HEAT SHIELD 

• AM MODS 

• SIDE WALL INSULATION 

• PAYLOAD & FLARE ADAPTER 

• FLARE 

• RCS L„ 


PROPELLANT REMAI 
MAIN ENGINE AND RC 


TUG INERTS = 7,828 

• PM = 5,868 I 
_ * AM = 1,960 J 






TUG WEIGHT AT START OF AEROBRAKING (THOUSANDS OF POUNDS) 


GROSS WEIGHT @ START OF AEROBRAKING 


I i I 
.ROUND TRIP PAYLO 


AEROBRAKING PENALTIES _ 

• HEAT SHIELD 

• AM 

• SIDEWALL INSULATION 

• PAYLOAD & FLARE ADAPTER 
•FLARE 

• RCSI sp 

• AM MODS | | | 


■■■■■■ 

mm 

B88S8B 

■■■■■■ 


■SI 


PROPELLANT REMAU 
■ MAIN ENGINE AND RC 


TUG INERTS = 7,828 LBS. 

• PM * 5,868 I 

• AM = 1,960— ] 


FIGURE 5. 1.2. 5-2: AEROBRAKIN 

(60° FLARE 


ROUND TRIP PAYLOAD (THOUSANDS OF POUNDS) 


O' ^ 



0 10 20 30 40 50 60 

NUMBER OF PASSES 


FIGURE 5. 1.2. 6-1: ROUND TRIP PAYLOAD VERSUS NUMBER OF PASSES 

(270 - 100 N.M. RECOVERY) 


D5-17142 


cn 

i 

4 * 

O 



FIGURE 5. 1.2. 6-2: ROUND TRIP PAYLOAD VERSUS NUMBER OF PASSES 

(DEPARTURE AND RECOVERY AT 200 N.M. ) 


D5-17142 




D5-17142 


\ 


5, 1.2.6 (Continued) 


For missions of greater than 20 passes, the 45° and 60° flares have slightly 
less payload sensitivity to mission time than do the other two configurations. 
However, for very short missions (5-10 passes), these two flares have the 
greatest sensitivities. The basic (no flare) configuration has the least 
payload sensitivity to mission time over the entire range studied (5-60 
passes). The 30° flare configuration represents an intermediate case over 
the entire range. 


Figures 5. 1.2. 6-3 and -4 are the same basic plots as 5. 1.2. 6-1 and -2 
above. Overplotted are isotemperature lines that correspond to the maximum 
heat shield nose steady state equilibrium temperatures developed in Section 
4.5. These specific temperatures were selected because they represent the 
various selected material temperature ranges of Section 4.7. These materials 
and their temperature ranges are repeated below. 


1200-1800°F 

1800-2200°F 

2200-2500°F 

2500-3500°F 


Rene 41 

TO-nickel -chrome 
Fansteel 85 
Fansteel 60 


The 2000°F line on these figures represents the state-of-the-art of radiative 
materials. Above this temperature, the new high temperature alloys of 
tantalum/col umbium/titanium such as the Fansteels 85 and 60 are required. 
These require advances in the state-of-the-art. 

Restricting the TD-nickel-chrome to 2000°F reduces the payload capability 
of the 30° flare about 100 pounds or 2-1/2%. The payload capability at 43 
passes (2000°F temperature limit) is 2850 pounds. The other flares are 
not affected by this temperature limit since both maximize payload at 
temperatures less than 2000°F. Therefore, for this nominal case (standard, 
atmosphere and no navigation errors), TO-nickel -chrome is a suitable material 
for the heat shield of the flared configurations for the maximized payload 
mission durations. Because the flared configurations sidewalls and flares 
have maximum temperatures less than the heat shield nose, TD-nickel-chrome 
and/or lesser temperature materials can be utilized. 

The basic (no flare) configuration maximizes payload at approximately 2500°F 
steady state and requires more than 60 passes to reduce the temperature to 
2200°F. This particular configuration was subjected to a transient analysis 
(Section 4.5). This analysis showed that the 2200°F limit was reached at 
approximately 35 passes and the 2000°F limit at 55 passes. These results 
were not shown on Figure 5. 1.2. 6-3 and -4 iri order to provide a common 
basis for comparison of all configurations. The effect of this transient 
analysis on the basic (no flare) configuration's heat shield material is 
to reduce the requirement from Fansteel 60 to Fansteel 85 for maximum 
payload mission duration (30 passes). 


The alternate payload placement mission capabilities for the four con- 
figurations are shown in Figure 5. 1.2. 6-5. This mission could not be 
flown from 200 NM (as with the payload retrieval and round trip missions) 


5-41 


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FIGURE 5. 1.2. 6-3: ROUND TRIP PAYLOAD VERSUS NUMBER OF PASSES 

(270 - 100 N.M. RECOVERY) 


D5-17142 








PLACEMENT PAYLOAD (THOUSANDS OF POUNDS) 


D5-17142 


IBB I 


BASIC (NO FLARE) 


• NO DOCKING ALLOWANCE 

IN GEOSYNCHRONOUS ORBIT 
FOR PAYLOAD PLACEMENT 

• DEPARTS FROM 100 N.M. 

• MAIN ENGINE PERFORMS 
CIRCULARIZATION AND 
TRANSFER 


30 FLARE 


45 u FLARE 


GROSS WEIGHT OF TUG & 
PAYLOAD ~ 62,300 LBS. 

_ UPON DEPARTURE 100 N.M. 


NUMBER OF PASSES 


FIGURE 5.1 .2.6-5; PLACEMENT PAYLOAD VERSUS NUMBER OF PASSES 

(270 - 100 N.M. RECOVERY) 









5. 1.2. 6 (Continued) 


because the gross weight of the Tug and payload exceeded the EOS capability. 
Approximately 8000 pound placement payloads were easily achieved with the 
nominal 270-100 NM mode; however, the 10,000 pound payload placement missions 
could probably not be accomplished even with optimization of departure and 
recovery altitudes. The payload placement mission does not require a docking 
delta velocity allowance at geosynchronous orbit. Therefore, the outbound 
delta velocity budget was established at the nominal 14,100 ft/sec. In 
addition, this particular mission impacts the basic (no flare) configura- 
tion's aerobraking inert weights shown in prior Figure 5. 1.2. 0-2. There 
is no payload during the aerobrakinq phase so that only a fixed insulation 
cap was placed over the exposed end of astrionic module. The conventional 
trajectory Tug's payload adapter was substituted for the aerobraking 
payload/flare adapter. The resulting payload placement mission basic 


(no flare) configuration 
changes in the ballistic 

weight statement 
coefficient: 

is shown 

below and reflect 


5 Pass 

10 Pass 

30 Pass 

60 Pass 

w i 

9,872 

9,528 

9,289 

9,545 

w p 

45,000 

45,000 

45,000 

45,000 

Gross 

54,872 

54,520 

54,089 

54,545 


Wt 


The flared configurations require all components of the aerobraking kit 
and their weights are similar to those shown previously in Figure 5. 1.2. 0-2 
but reflect the W/CpA changes. This flexibility in selection of aero- 
braking kit components is directly reflected in Figure 5. 1.2. 6-5 with the 
basic no flare configuration showing a relatively large payload advantage. 

The sensitivities to mission time and maximum equilibrium temperature con- 
straints are similar to the round trip payload case. 

The other alternate mission capability, that of payload retrieval, is 
shown in Figure 5. 1.2. 6-6. This mission could be flown using the 200 NM 
mode since only the fully fueled Tug required low earth orbit insertion. 

As in the placement case, further payload increases by departure and recovery 
altitude optimization could be achieved. The payload retrieval capabilities 
are slightly greater than the placement capabilities. This is because: 

(1) A different operational mode was used; and (2) aerobraking reduces the 
delta velocity requirements of the heavier Tug and retrieved payload during 
the return trip. Carrying no payload to orbit (large propulsive requirement) 
with aerobraking return of a retrieved payload (small propulsive require- 
ment) results in the basic (no flare), 45° flare and 60° flare configurations 
to be approximately equivalent for missions greater than 5.5 days (30 
passes). 


RETRIEVED PAYLOAD (THOUSANDS OF POUNDS) 


D5-17142 








■ 


■ 

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• DOCKING ALLOWANCE IN 
GEOSYNCHRONOUS ORBIT 
_ FOR PAYLOAD RETRIEVAL. 

- BAS 

1 


B 

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■ 



• DE 
AT 



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0 TO 20 30 40 50 60 


NUMBER OF PASSES 

FIGURE 5.1 .2.6-6: RETRIEVED PAYLOAD VERSUS NUMBER OF PASSES 

(DEPART AND RECOVER AT 200 N.M.) 


5-46 



D5-17142 


5. 1.2. 6 (Continued) 


Figure 5. 1.2. 6-7 is a summary capture map illustrating the capabilities of 
aerobraking to accomplish the 287 geosynchronous missions contained in the 
mission model utilized. The shaded area in all three blocks shows the 
cumulative percent of the mission model within the payload weight incre- 
ments. For example, 64% of all payloads weigh less than 3000 pounds and 95% 
weigh 3000 pounds or less. There is a large discontinuity beyond payload 
weights of 3000 pounds with the next level estimated to be approximately 
7000-10,000 pounds. 

The round trip capability (200 NM mode) easily captures the 3000 pound 
payload class and below. The <v 1000 pound excess payload capability margin 
is an indication that if the inert weights and/or delta velocities used 
in this analysis were optimistic by a significant amount, the 3000 pound 
round trip payload missions could still be accomplished. The placement 
(270-100 NM mode) and retrieval (200 NM mode) payload capabilities are 
approximately 8000 and 9000 pounds respectively. Neither of these capa- 
bilities capture the proposed 10,000 pound payloads. However, reduction 
of these payload weights to those matching the aerobraked Tug's capability 
would still provide for placing and retrieving relatively large payloads 
in geosynchronous orbit. The figure illustrates only the basic Tug and 
the 30° flare configuration. The 45° and 60° flare configuration will 
also capture 95% of the missions. 

5.2 ATMOSPHERIC PERTURBATION AND NAVIGATION ERROR SENSITIVITIES 

Sections 4.3, 4.5, and 4.8 (trajectory, thermal and weights analyses, 
respectively) contain data on the effects of atmospheric perturbations 
from the 1962 Standard Atmosphere used in this study. Section 4.6 
(astrionics analysis) discusses the effects of state vector uncertainties, 
navigational errors, and midcourse correction burns. This subsection dis- 
cusses the combination of the results of these previous analyses and the 
resulting payload sensitivities to these parameters. The sensitivities 
shown are for the high atmospheric density range and for 1 sigma (<r) or 
3 sigma (<r) errors (or uncertainties) from the astrionics analysis. The 
study scope did not permit a more extensive analysis in this area. 

5.2.1 Effects on Nominal Perigee Altitudes 

Figure 5. 2. 1.0-1 shows the 1 a radial perigee position uncertainties for 
various midcourse correction burn options. As discussed in Section 4.6, 
a corrective burn after the Horizon Sensor updates (at altitudes far from 
perigee altitude) actually increases the perigee error. With no burn 
correction applied, the 3 o perigee uncertainty Is approximately 5.1 NM 
(31,000 feet). Correcting to a lower error value, such as 0.5 NM, can be 
relatively inexpensive in terms of RBS propellant consumption (see prior 
Figure 4. 6. 4. 6-2). Therefore, these two large error cases were eliminated 
from the analysis. 


5-47 


PAYLOAD WEIGHT (K LBS) PAYLOAD WEIGHT (K LBS) 



CUMULATIVE PERCENT OF PAYLOADS 


PAYLOAD PLACEMENT 



CUMULATIVE PERCENT OF PAYLOADS 


FIGURE 5.1. 2.6-7 


PAYLOAD WEIGHT (K LBS) 


BASIS: 

• EOS CAPABILITY ^ H-33 ORBITER 

• 287 GEOSYNCHRONOUS MISSIONS 

• 30 PASS MISSIONS 

CONCLUSIONS: 

• 95% OF ALL GEOSYNCHRONOUS 
MISSIONS CAPTURED 



OF AEROBRAKED TUG 


05-17142 



RADIAL POSITION UNCERTAINTY 
AT PERIGEE (N.M.) 



NO BURN 
CORRECTION 


AV BURN 

AFTER 

HORIZON 

SENSOR 

UPOATES 


AV BURN 
, AFTER 
500 SEC 
LANDMARK 
TRACKING 


AV BURN 
AFTER 
1300 SEC 
LANDMARK 
TRACKING 


FIGURE 5. 2. 1.0-1: RADIAL POSITION UNCERTAINTY AT PERIGEE 






D5-17142 




D5-17142 


5.2.1 (Continued) 

Figure 5. 2. 1.0-2 shows the atmospheric perturbation data utilized for the 
high and low cases. This information was furnished by NASA/MSFC (prior 
Reference 4.3.4.0-1). The original data from HSFC included the expected 
range under both summer and winter conditions. Because the study limitations 
did not permit an investigation of each season independently, the two ranges 
were compared and the extreme high and low ranges selected. The selected 
limits are shown on Figure 5. 2. 1.0-2 as Winter Low and Summer High. This 
conservative approach then encompasses all cases with the exception of 
geomagnetic storms (not considered). The percentage of expected pertur- 
bation is greater at higher altitudes than at the aerobraking periqee 
altitude region where maximum equilibrium temperatures occur. In the perigee 
region, the highest densities expected vary from 140% to 165% of the 1962 
Standard Atmosphere with the lowest densities varying from 50% to 65% of 
the nominal. The range of atmospheric perturbations (or anomalies) depicted 
in Figure 5. 2. 1.0-2 are unpredictable. Not shown in this figure are the 
mean perturbations (dependent on solar activity) which are predictable. 

Within the target perigee range, these mean values are near nominal. 

Because of these two facts (predictability and less variability), the mean 
values could be programmed into the mission sequence as nominal and as 
such, were not analyzed in this study. 

Figure 5. 2. 1.0-3 shows the maximum equilibrium temperatures for the beat 
shield nose as functions of atmosphere density range and navigational 
errors. The solid lines indicate the temperatures experienced in the high 
density atmosphere and the dotted lines represent the Standard atmosphere. 

No transient (heat sink) effects are included so that the data shown will 
be conservative. 

As discussed in the thermal analysis, Section 4.5, the impact of the hiqh 
density atmosphere will result in an approximate 65 and 80°F rise in 
maximum equilibrium temperature for the 60° flare and basic (no flare) 
configurations, respectively, with no navigation errors. 

The impact of only the high density atmosphere is shown along the ordinate 
of Figure 5. 2. 1.0-3. The 30 pass basic (no flare) configuration (atmos- 
phere plus navigation error) will have a maximum (“worse case") temperature 
rise of approximately 360° F. This maximum increase is measured from the 
Standard Atmosphere with zero navigation error to the high density atmos- 
phere with 3ff? (500 seconds of landmark tracking) navigation error. The 
60° flare configuration's maximum rise is approximately 225°F under the 
same conditions. 

1 * 

The 60° flare's maximum equilibrium temperature does not exceed 2000°F 
and therefore TD-nickel -chrome remains a suitable material for the "worst 
case" 30 pass mission. The basic (no flare) has a 30 pass "worst case" 
equilibrium temperature of 2850°F (steady state). Allowing for a possible 
transient analysis decrease of 250°F, this worst case would still neces- 
sitate advanced alloys such as Fansteel 60 ( 2500-3500° F). 


5-50 


D5-17142 




FIGURE 5.2.1. 0-2. RANGE OF ATMOSPHERIC PERTURBATIONS 

5-51 




PERIGEE ALTITUDE ERROR BELOW TARGET PERIGEE (THOUSANDS OF FEET) 


FIGURE 5.2.1 .0-3. MAXIMUM EQUILIBRIUM TEMPERATURE VS. INITIAL PERIGEE ERROR 

(FIRST PASS OF 30-PASS MISSION) 


D5-17142 


D5-17142 


5.2.1 (Continued) 

The 30° and 45° flares are shown for the Standard Atmosphere case only. 

The high density atmosphere thermal analysis was not performed on these 
two configurations, but it is expected that the atmospheric density dis- 
persion impact will be similar to that shown for the other two configurations 
and Figure 5. 2. 1.0-3 indicates the probable results. The 45° flare's 
maximum equilibrium temperature would be under 2200°F (TD-nickel -chrome) . 

The 30° flare's maximum temperature would be approximately 2450°F and 
require advanced state-of-the art materials (Fansteel 60). 

5.2.2 Payload Sensitivities 

The trajectory analysis (Section 4.3) used a first pass target perigee 
based on the nominal Standard 1962 Atmosphere. This assumed, as discussed 
in Section 5.2,1 above, that the atmospheric perturbations were not pre- 
dictable. The impact of the perturbed first pass was then utilized to 
adjust apogee decay on subsequent passes so that the number of mission 
passes (mission time) was held exactly at or approximately at the 
desired constant. In the high density atmosphere, the Tug experienced 
higher equilibrium temperatures and higher pressure loads. As discussed 
in Section 4.8 and shown in Figure 5. 1.2. 0-2, the temperatures and loads 
were major factors in the determination of the heat shield and flare 
weights. Figure 5.2.2.0-T shows the basic (no flare) and 60° flare 
configuration aerobraking weights and delta velocities associated with 
the high density atmosphere. The delta velocities shown are based on "an 
impulse at entry" technique developed for this study. This technique was 
investigated to overcome the effects of the large navigational uncertainties 
encountered at either exit or apogee. The technique has several advantages 
including: 

o Correction burns are made based on the greatest knowledge of 
current position and velocity errors. 

o Atmospheric perturbation correction bums can be combined with the 
navigational error correction burns thereby simplifying the pro- 
cedures and probably reducing the total propulsive Impulse 
required. 

o Relative insensitivity of the required impulse to mission duration 
allowing for better mission planning. 

o Final apogee control sufficiently accurate to permit alternate 
circularization altitudes (e.g., 200 n.m.). 

Figure 5. 2. 2. 0-2 shows the effect of the high density atmosphere on the 
round trip capability of the basic (no flare) and 60° flare configurations. 
Both configurations easily maintain their 3000 pound plus capability but 
some have reduction in capability due to this environment. The reduction 
is mostly due to the added inert weights (short duration missions) and the 
added delta velocities (longer duration missions). The maximum payload 
capabilities (30 passes) are only decreased approximately 10-122 with 


5-53 


5-54 


<T 




0 
<J1 

1 





ro 


* DEPART AND RECOVER AT 200 N.M* 

** INCLUDES PAYLOAD INSULATION AND PAYLOAD CAP ACTUATION DEVICE 


FIGURE 5.2. 2,0-1 FULLY FUELED WEIGHT STATEMENTS FOR HIGH DENSITY ATMOSPHERE 






D5-17142 


5.2.2 (Continued) 

somewhat greater decreases at the mission duration extremes shown. For the 
short duration missions (~10 passes), both configurations are impacted 
approximately the same. For mission durations of 20 passes or more, the 
60° flare configuration is impacted by this atmospheric condition more than 
the basic (no flare) configuration. 

The high density atmosphere's major mission imoact is the decrease in the 
range of mission durations (and configurations) that will allow for 3000 
pound round trip payload missions. Assuming materials such as Fansteel 
60 were available and with a Standard Atmosphere, trip times of less than 
one day (5 passes) were feasible for the 3000 pound payload using either 
configuration. Long duration missions of 11 days (60 passes), again assumina 
a Standard Atmosphere, were easily accomplished by either configuration 
with state-of-the-art materials. With the high density atmosphere, 
the 5 pass 3000 pound payload mission may be marginal. The mission dura- 
tion range is therefore between 10 and 60 pass mission region. 

Also shown on Figure 5. 2. 2. 0-2 are the steady stats temperatures associated 
with the material limits discussed in Section 5.1 .2.6. The high density 
atmosphere impact is an increase of approximately 4 or 5 passes to maintain 
equal maximum equilibrium temperatures. The basic (no flare) configura- 
tion's apparent requirement for a 2500°F plus material (e.g., Fansteel 60) 
is offset by the relative payload insensitivity to missions having between 
30 and 40 passes and by considering transient (heat sink) effects. 

Figure 5. 2. 2. 0-3 shows the round trip payload sensitivities (10 pass mission) 
to atmospheric anomalies and navigational errors. Shown in the first two 
hatched columns of each configuration are the payloads and temperatures pre- 
viously discussed (Fiqure 5.2.2 .0-2 ) . The third hatched column of each 
configuration shows the round trip payloads achievable under the combined 
effects of the high density atmosphere and the 3 sigma nagivational errors 
(1300 seconds of landmark tracking). Because the prior payload analyses has 
reserved 400 ft/second delta velocity for navigation error corrections, 
including the navigational errors has a minimum impact on the round trip, 
payload. Even for the relatively short duration 10 pass mission, the 3000 
pound payload capability is retained under these "worst" conditions depicted. 
Further bounding of the atmospheric anomalies will result in increased payloads. 

The temperatures shown in Figure 5. 2. 2. 0-3 are the heat shield nose stagnation 
temperatures. The expected increases are only 110° for the 60° flare and 
160° for the basic no-flare. These increases are nominal when the wide 
environmental variations encountered are considered. 

5.2.3 Perturbation Summary and Conclusions 

The study scope limited the analysis of the atmosphere density variations 
to the specific cases of a constant high density, standard, or low density 
atmosphere. Pass-to-pass unpredictable variations might be encountered 
in addition to the predictable variations that were considered nominal 


5-56 


ROUND TRIP PAYLOAD 
(THOUSANDS OF POUNDS) 


BASIC (NO FLARE) 
CONFIGURATION 


2990° F 


3085°F 


3130 # F 


60° FLARE 
CONFIGURATION 


2080° F 

iW 


BASIS: 

• MAIN ENG : ME CIRCULARIZATION 

• DEPART & RECOVER AT 200 N.M. 

• MAXIMUM STEADY-STATE 
HEAT SHIELD EQUILIBRIUM 
TEMPERATURES (*F) 

• 10 PASS MISSION 


2160°F 


2190° F 


STANDARD 1962 ATMOSPHERE 
AND NO NAVIGATIONAL ERRORS 


HIGH DENSITY ATMOSPHERE 
AND NO NAVIGATIONAL ERRORS 


n 


HIGH DENSITY ATMOSPHERE 
AND 3 a NAVIGATIONAL ERRORS 
(RSS CORRECTION /IV) 


FIGURE 5. 2.2. 0-3: 


ATMOSPHERIC ANAM01Y AND NAVIGATION ERROR EFFECT ON 
ROUND TRIP PAYLOAD (10 PASS MISSION) 




D5-17142 


5.2.3 (Continued) 

and pre-programmed into the flight trajectory. Prior to this type of 
detailed analysis, these pass-to-pass unpredictable variations should be 
bounded and rates of density changes defined (e.g., how much pass-to-pass 
variation could be expected and how much total variation could be expected 
from the beginning to the end of a 2, 5, or 10 day mission?). 

Only a cursory examination was made of the combination of atmospheric 
variations, navigational uncertainties or errors, and midcourse correct- 
ions. No attempt was made to optimize a guidance scheme to accommodate 
these factors. Possible schemes to minimize these effects include: (1) Tar- 

geting the first pass perigee higher than nominal to insure only the nominal 
conditions are encountered as the worst case, (2) overdesign of the vehicle 
to withstand worst case conditions at nominal target perigees, (3) correcting 
with lift as temperatures or loads vary from nominal, and (4) combinations 
of these. All of these concepts have advantages and disadvantages. The 
approach taken in this section was to consider the aerobraking tug to be 
overdesigned and require additional inert weights and impact the mission 
delta velocity budget. 

The trends and conclusions of the atmospheric perturbation and navigational 
error analysis are as follows: 

o The large flare (45° and 60°) configurations can utilize state- 
of-the-art heat shield and flare materials for the mission 
durations that maximize payload ™ 30 pass missions. 

o The basic (no flare) and 30° flare configurations require advanced 
state-of-the-art materials having maximum temperature limits of 
approximately 2500° (e.g., Fansteel 85) for their maximum 
payload missions (30 passes). 

o The maximum equilibrium temperature of a n-pass high density 
atmosphere is approximately equal to the more severe tempera- 
tures encountered with a four or five less passes mission in a 
Standard Atmosphere. 

o The high density atmosphere reduces the flexibility in mission 
duration selection but does not significantly impact the payload 
capability for the maximum payload mission duration (av 30 
passes) . 

o Prior to atmospheric entry, midcourse corrections to compen- 
sate for navigational errors are desirable. Further these 
corrections can be combined with the atmospheric anamoly 
corrections to achieve operational simplicity. 

o Pass-to-pass mission atmospheric variations require better 
definition. 


D5-17142 


5.2.3 (Continued) 

o An in-depth sensitivity analysis of the combination of naviga- 

tion and atmospheric effects is required in a follow-on activity 
to define guidance schemes and to optimize the design and 
operational modes in a perturbed environment. 

5.3 AEROBRAKING/CONVENTIONAL TRAJECTORY 1 TUG COMPARISONS 

This section discusses the required conventional trajectory Tuq sizes to 
deliver the equivalent aerobraked mode payloads and the sensitivities of 
the conventional and aerobraked Tugs to performance parameters. The com- 
parisons, made between the aerobraking and conventional trajectory Tugs, 
used the same groundrules, i.e., identical (1) specific impulses, (2) de- 
parture and recovery altitudes, (3) docking delta velocity budgets, and 
(4) basic astrionic module weights. Both the aerobraking and the conven- 
tional Tugs were considered to be single stage reusable configurations. 

The propulsion module weights for the conventional trajectory Tuqs were 
derived from prior Reference 1.1. 0.0-1. 

5.3.1 Conventional Trajectory Tug Size Comparison 

Figure 5. 3. 1.0-1 shows the conventional trajectory Tug weights as a func- 
tion of geosynchronous payload weight. The zero payload capabilities are 
not identical because the departure and recovery altitudes were selected 
identical to those of the aerobraking mode (i.e., 270-100 NM placement 
and 200 NM retrieval and round trip). The 45,000 pound propellant conven- 
tional trajectory Tug (53,028 pounds gross) is slightly undersized for the 
zero payload placement case. For the retrieval and round trip payload 
missions, required conventional trajectory Tug weights rapidly increase 
with increased payload. Payload placement is a less demanding mission for 
the conventional trajectory and is reflected in the smaller Tuq weight 
increase per payload increase. 

Shown in Figure 5. 3. 1,0-1 are the maximum payload capabilities of the 30° 
flare aerobraked Tuq (Section 5.1.2) and the gross weight of this Tug. 

The aerobraked 30° flare Tug's (gross 55,066 pounds) payload placement 
capability of 7250 pounds is matched by a conventional trajectory Tug 
having a gross weight of 68,000 pounds (less payload). For the retrieval 
mission, the 55,066 pound aerobraked 30° flare Tug is equivalent to a 

96.000 pound (less payload) conventional trajectory Tug in payload 
capability. The 4425 pound payload round trip mission has a gross Tug 
weight difference of 26,000 pounds (55,066 pounds aerobraked 30° flare Tuq, 

81.000 pounds conventional). All three of the equivalent maximum payload 
capability conventional trajectory Tugs will require two Shuttle launches 
per mission in either the ground based or space based modes. 

Used in the ground based mode (i.e.. Tug begins and ends each mission on 
earth within the Shuttle's cargo bay), the gross weight of the conventional 
trajectory Tug is beyond one Shuttle's capability to deliver it to low 
earth orbit. For the larger conventional trajectory Tug stages, the 60' 
length limitation of the Shuttle's cargo bay and the Shuttle's payload 


5-59 


WEIGHT OF PAYLOAD - 1000 LBS WEIGHT OF PAYLOAD - 1000 LBS 



TUG WEIGHT (LESS PAYLOAD) - 1000 LBS TUG WEIGHT (LESS PAYLOAD) - 1000 LBS 


ROUND TRIP 



TUG WEIGHT (LESS PAYLOAD) - 1000 LBS 


BASIS: 


• PAYLOAD PLACEMEMT MISSIONS 
DEPART AND RETURN TO 100 N.M. 

o PAYLOAD RETRIEVAL AND ROLND 
TRIP PAYLOAD MISSIONS 
DEPART AND RETURN TO 200 N.M. 

• TANDEM STAGE VEHICLE IS TWO EQUAL 
STAGES BOTH RECOVERED 

• STAGE WITH DROP TANK HAS EQUAL 
PROPELLANT IN STAGE AND DROP TANK 

• DROP TANK IS EXPENDED IN SYNCHRONOUS 
ORBIT 

• ASTRIONICS MODULE WEIGHT - 1960# 

• PAYLOAD ADAPTER WEIGHT - 200# 

• STAGE ADAPTER AND STAGE 
SEPARATION MECHANISM - 324# 


FIGURE 5.3.1 .0-1 : PAYLOAD CAPABILITY COMPARISON - AER0BRAKED VERSUS CONVENTIONAL TUG 


D5-17142 


D5-17142 


5.3.1 (Continued) 

capability are dual constraints. Therefore, Tug in-orbit assembly and/or 
fueling is required prior to departing the low earth parking orbit for 
geosynchronous orbit. 

Used in the space based mode (i.e., Tug remains on orbit and Shuttle delivers 
propellant and payload to Tug), the required propellant weights for the 
equivalent maximum payloads are beyond the single Shuttle capability. 

Using this mode, Tug in-orbit refueling is required prior to each geo- 
synchronous mission. 

5.3.2 Sensitivities to Performance Parameters 

The conventional trajectory single reusable stage Tuq was shown in Reference 
1.1. 0.0-1 to be sensitive to performance parameters. The performance 
sensitivity analysis of that previous study was accomoiished by allowing 
stage weight to grow (shrink) so that the payload remaned constant. 

Because of the scope of this aerobraking study, investigation of the aero- 
braking effects on larger (smaller) stages was not conducted. Extrapolating 
the aerobraking impact data (e.g., aerodynamic, thermal, and weights) to 
configure other aerobraked Tugs is not justified until follow-on in-depth 
analyses are made. Therefore, round trip payload was selected as the 
dependent variable for the performance sensitivity analysis and the pro- 
pellant loading in the aerobraked and conventional trajectory Tugs was 
held constant. The 30 pass basic (no flare) configuration (45,000 pounds 
propellant) was utilized as the representative of the aerobraked configura- 
tions. The comparable conventional trajectory single reusable stage has 
a propellant loading of 71,500 pounds (aj 80,000 total Tug weight). 

5.3.2. 1 Mass Fraction (X') 

Figure 5. 3. 2. 1-1 shows the round trip payload sensitivities to mass 
fraction. The conventional trajectory Tug (including astrionic module 
and payload adapter) has a significantly higher mass fraction (0.892) 
than does the aerobraked Tug (0.824) for the baseline payload. Note: The 

large size of the conventional Tug (80,150 pounds) accounts for the high 
mass fraction. The aerobraked Tug is smaller (54,600 pounds) and therefore 
has a part of its lower mass fraction attributable to size effects. 

Both Tugs maintain relatively constant exchange ratios i d /$X') over 
the range of mass fractions shown. The aerobraked Tug's payload capa- 
bility is less sensitive to changes in mass fraction (lower exchange ratio). 
Dropping the mass fraction by 0.015 (-0.015 on the figure) decreases the 
aerobraked payload approximately 1000 pounds and the conventional tra- 
jectory payload decreases 1600 pounds. If the mass fractions were increased 
by 0.015 (+0.015 on the figure), the aerobraked payload increases 1000 
pounds and the conventional trajectory payload increases 1300 pounds. 
Therefore, mass fraction changes will impact the conventional trajectory 
Tug more severely than the aerobraked Tug. 



VARIATION IN MASS FRACTION (X ) 


v 


C/ 





FIGURE 5. 3. 2. 1-1: COMPARISON OF ROUND TRIP PAYLOAD SENSITIVITIES TO MASS FRACTION 


D5-17142 




I 


D5-17142 


5. 3. 2. 2 Specific Impulse 

Figure 5. 3, 2. 2-1 compares the sensitivities to specific impulse. The data 
shown assumes that the variations in Isp could be achieved by equal engine 
weights and nozzle dimensions. (The nozzle's dimensions are important 
parameters in the aerobraked aft heat shield design and weight). For 
higher specific impulse than nominal values (461-470), both Tugs are about 
equally sensitive. Increasing the Isp from 460 to 450 seconds results in 
a small aerobraking advantage. However, the overall comparison is that 
nominal Isp variations affect both Tugs approximately the same. 

5. 3. 2. 3 Delta Velocity 

Figure 5. 3. 2. 3-1 compares the effects of mission delta velocity changes. 

The aerobraked Tuq is somewhat more sensitive to delta velocity than the 
conventional trajectory Tuq because of its lower mass fraction. Adding 
1200 ft/sec to the mission's total delta velocity budqet reduces both 
stage capabilities to less than 3000 pounds. The conventional trajectory 
stage essentially maintains its exchange ratio (dWpid/ddV) through the 
variations shown. The aerobraked stage's sensitivity to delta velocity 
increases with decreasing velocity requirements. 

The round trip payload mission has been flown usinq the 200 NM departure 
and recovery mode and with one docking allowance. Therefore, it is anti- 
cipated that decreases in the baseline delta velocity budgets would be 
minor (800 ft/sec or less). As discussed in Section 5.2., the delta 
velocity budget for aerobraking Tug can increase to account for atmospheric 
variations and midcourse corrections. The conventional trajectory Tug 
could also have increased budgets due to long term space storage, docking 
with an orbital propellant depot, and in-orbit assembly operations. 

Probable delta velocity increases for both Tugs should be less than 1200 
ft/sec. Within these upper and lower bounds, the aerobraked Tug is more 
sensitive to changes in delta velocity than the conventional trajectory 
Tug. 

5.4 SENSITIVITY ANALYSIS CONCLUSIONS AND RECOMMENDATIONS 

The sensitivity analysis has indicated certain definite trends, some of 
which require more comprehensive study in follow-on efforts. Most of 
these trends and follow-on efforts have been discussed in the previous 
sections. Following the tabulations listed below, a summary discussion 
of selected conclusions and recommendations is made. The sensitivity 
analysis conclusions are as follows: 

o The operational mode used by the aerobraked Tuq should optimize 
the Shuttle capability to deliver and recover in low earth 
orbit. 

o An optimum aerobraked configuration should have a stabiliza- 
tion/drag device. (Assuming the 2000°F thermal constraint and 
a Shuttle recovery of the Tug within the seven day on-orbit 
Shuttle capability.) 




5-63 




VARIATION IN TOTAL MISSION DELTA 
VELOCITY (FEET PER SECOND) 


C0£VEN™U^^A^CT0RY_4V ^MzOJT./SEC 
AEROBRAKED A V = 21.46CTft.7sEC. -T — — 


• BASIC (NO FLARE) AEROBRAKED STAGE 

WT. = 54,643 LBS. 

• CONVENTIONAL SINGLE REUSABLE STAGE 

WT. = 80,193 LBS. 

• l $p = 460 SEC. 

• DEPARTS AND RECOVERS AT 200 N.M. 


BASELINE (NO FLARE) AEROBRAKED 
TUG PAYLOAD CAPABILITY 
I 

I 


ROUND TRIP PAYLOAD (POUNDS) 


FIGURE 5. 3. 2. 3-1: COMPARISON OF ROUND TRIP PAYLOAD SENSITIVITIES TO MISSION DELTA VELOCITY 


D5-17142 



D5-17142 


t 


5.4 (Continued) 

o Round trip geosynchronous payloads of 3000-4000 pounds are 
achievable by a 45,000 pound propellant aerobraked Tuq. 

o Placement and retrieval of geosynchronous payloads of 7000-9000 
pounds are within the capability of the 45,000 pound propellant 
aerobraked Tug. 

o The single Shuttle/Tug launch per geosynchronous mission is 
possible for 95% of the missions using Tuq aerobrakinq. 

o Advanced state-of-the-art materials are required for configura- 
tions of less drag and/or shorter mission times. 

o Midcourse corrections can be made at an optimized time prior 
to reaching perigee to reduce the navigation and atmosphere 
uncertainty impact. 

o The equivalent round trip payload conventional trajectory Tuq 
has a gross weight of approximately 26,000 pounds more than the 
aerobraked Tuq (~45% heavier than aerobrakinq Tuq). 

o The aerobraked Tug is less sensitive to changes in mass fraction 
than the conventional trajectory Tug. Sensitivities to specific 
impulse and delta velocity are less pronounced. 

The following recommendations for advanced technology programs and follow- 
on aerobraking studies are made as a result of the sensitivity analysis: 

o Better definition of the short time-span atmospheric variations 
and their rate of change. 

o In-depth analysis of the combinations of atmospheric variations, 
navigational errors, and guidance/midcourse correction schemes. 

o Continued development of high temperature materials, par- 
ticularly in the 2000-2500°F class. 

o Analysis of alternate drag device/stabilization configurations 
with a short high angle flare (near neutral stability) as one 
candidate. 

o Optimization of the Shuttle/Tug delivery and recovery orbits. 

o Impacts of possible Tug interim RCS propellant such as mono- 

propellants and bi propel 1 ants. 

The Shuttle capabilities used as a reference in this analysis are subject 
to change as the Shuttle studies continue. As noted in the sensitivity 
analysis, changing the delivery and recovery mode to more nearly fully 
utilize the capabilities of both vehicles resulted in approximately 1000 


5-66 


D5-17142 


5.4 (Continued) 

pounds of additional round trip payload. This optimization should be 
continued in a follow-on Tug aerobraking study using the results of the 
current Shuttle extended Phase B studies. 

The 30° flare had the maximum round trip payload because of its lighter 
flare weight and low equivalent delta velocity. Smaller physical size 
was the major factor in the weight and its stability resulted in lower 
delta velocities. It also had lower temperatures than the basic (no flare), 
because its increased drag permitted higher perigee altitudes. However, 
this configuration is probably not optimum for the aerobraked Tug. For 
example, a short 60° flare, with the same near-neutral stability as the 
30° flare, could be superior. 

The payload capabilities of all configurations studied maximized at appro- 
ximately 30 passes (5.5 days). Using the Standard Atmosphere, a shorter 
mission (10 passes/2 days) or a lonqer mission (60 passes/ll days) did 
not greatly impact the payload capability. The hiqh density atmosphere 
results indicated that a nominal mission should have 20-40 passes. The 
maximum equilibrium temperatures increased rapidly as the mission time 
shortened from 5.5 days and decreased slowly as mission time increased. 

The 30° flare configuration (30 passes) might experience steady state 
temperatures of approximately 2425° in the high density atmosphere - navi- 
gational error environment. Any configuration of similar drag coefficients 
could be expected to experience similar temperatures. To use the maximum 
payload potential of this configuration (or stabili ty/draq) class, new 
materials to withstand the environment are required. With the transient 
effects considered, the temperature drops to approximately the upper limit 
of TD-nickel-chrome. Shorter mission times (e.g., 20 passes) would 
cause temperatures to exceed this TD-nickel-chrome limit but still be 
below 2500° F. 

This study had the basic assumption that GO 2 /GH 2 RCS development would be 
an integral part of the EOS program. The Tug inert weights, RCS perform- 
ance, and payload capabilities were computed' on this assumption. If it 
appears that this advanced system will not be available at the- time the 
aerobraked Tug is to be placed into operation, an alternate system must 
be considered and the impacts on the aerobraking mode analyzed. 


5-67 


I 


D5-17142 

LIST OF REFERENCES 




1.1. O.O-l 


3.3.0. 0-1 


4.1. 1.1-1 


4.1. 1.1-2 


4.1. 1.1-3 


4. 1.1. 1-4 


4. 1.1. 1-5 


4.1. 1.1-6 


4. 1.1. 1-7 


4.1. 1.1-8 


4.1. 1.1- 9 

4.1.1.1- 10 


4.1.1.1-11 


Technical Study for the Use of the Saturn V, INT-21 
and Other Saturn V Derivatives to Determine an Optimum 
Fourth Stage (Space Tug), The Boeing Company, NAS8-5608, 
Schedule II, February 26, 1971 

Astrionic System Optimization and Modular Astrionics for 
NASA Missions After 1974 - Preliminary Definition of 
Astrionic System for Space Tup Mission Vehicle Payload 
(MVP), Progress Report dated 16 June 1970 to 15 August 
1970, IBM Number 69-K44-0006H, MSFC-DRL-008, Line Item 
Number 268 

Keyes, J. W., "Aerodynamic Characteristics of Lenticular 
and Elliptic Shaped Configurations at a Mach Number of 
Six," NASA TN-D-2606, dated February 1965 

Kinslow, Max, and Potter, J. L., "The Drag of Spheres 
in Rarefied Hypervelocity Flow", AEDC-TDR-62-205, 

December, 1962 

Kussoy, M. I., et al , "Sphere Drag in Near Free Molecule 
Hypersonic Flow", AIAA Journal, Vol. 8, No. 11, 

November 1970 

Mechanics of Rarefield Gases, Volume 5, Section 16 of 
Handbook of Supersonic Aerodynamics, NAVWEPS Report 
1488 (Vol. 5, Section 16) dated 1961 

Schaaf, S. A. and Chambre, P. L., Flow of Rarefield Gases, 
Princeton University Press, 1961 

Cox, R. N. and Crabtree, L. F., Elements of Hypersonic 
Aerodynamics, Academic Press, 1965 

Sentman, L. H., "Free Molecule Flow Theory and its 
Application to the Determination of Aerodynamic Forces," 
LMSC-448514, October 1961 

Schlichting, H., Boundary Layer Theory, McGraw Hill, 1968 

Nielson, J. N., Missile Aerodynamics, McGraw Hill, 1960 

Patterson, G. N., Molecular Flow of Gases, J. Wiley & 

Sons, 1956 

Hoerner, S. F. , Fluid Dynamic Drag, Published by the 
Author, 1961 


PRECEDING PAGE BLANK NOT FILMED 




i 


c 


D5-17142 


4. 1.1. 2- 1 

4. 1.1. 2- 2 

4. 1.1. 2- 3 

4. 1.1. 2- 4 

4. 1.1. 2- 5 

4.1. 1.2- 6 

4. 1.1. 2- 7 

4. 1.1. 3- 1 

4.1. 1.3- 2 

4. 1.1. 3- 3 

4.1. 1.3- 4 

4. 1.2. 1- 1 

4. 1.2. 1- 2 


LIST OF REFERENCES [Continued) 

Seiff, A. and Whiting, E. E., "Calculation of Flow Fields 
from Bow-Wave Profiles for the Downstream Region of 
Blunt-Nosed Circular Cylinders in Axial Hypersonic Flight," 
NASA-TN-D-1147 

Seiff, A., "Secondary Flow Fields Embedded in Hypersonic 
Shock Layers," NASA-TN-D-1304 

Seiff, A., and Whiting, E. E., "A Correlation Study of 
the Bow Wave Profiles of Blunt Bodies," NASA-TN-D-1148 

Truitt, R, W., Hypersonic Aerodynamic, Ronald Press Co., 
1959 

Sentman, L. H., "Tables of Free Molecule Flow Functions," 
Appendix C to LMSC TR-448514, LMSC 448514-1 

Blick, E. F., Aerodynamic Coefficients in the Slip and 
Transition Regime, AIAA Journal Vol. 1, No. 11, 

November 1963 

Matting, F. W., "Approximate Bridging Relations in the 
Transitional Regime between Continuum and Free Molecule 
Flows," Journal Spacecraft and Rockets, Vol. 8, No. 1, 
January 1971 

Wagner, R. D. , and Watson, R., "Reynolds Number Effects 
on the Induced Pressures on Cylindrical Bodies with 
Different Nose Shapes and Nose Drag Coefficients in 
Helium at a Mach Number of 24," NASA TR-R-182 

Henderson, A., "Investigation of Flow over Simple Bodies 
at Mach Numbers of the Order of 20," NASA TN-D-449 

Van Hise, V., "Analytic Study of Induced Pressure on Long 
Bodies of Revolution with Varying Nose Bluntness at 
Hypersonic Speeds," NASA TR-R-78 

Henderson, A., et al, "Fluid Dynamic Studies to M = 41 
in Helium," AIAA Journal, Vol. 4 No.12, December 1966 

DeRose, C. E. , "Ballistic Range Tests of a Drag Ring 
Configuration at Mach Numbers Around 2," NASA TN-D-4291, 
December 1967 

Ericsson, L. E., "Unsteady Aerodynamics of an Ablating 
Flared Body of Revolution Including the Effects of Entropy 
Gradient," AIAA Journal, Vol. 6 No. 12, December 1968 


D5-17142 


4. 3. 1.1- 1 

4. 3. 1.1- 2 


4.3. 1.1-3 


4. 3. 4. 0-1 


4. 5. 1.0-1 


4. 5. 1.0-2 


4.6. 4. 4-1 


4. 6. 4. 5-1 


4.6.8. 1-1 


4. 6. 8. 1-2 


4.6.8. 2-1 


4. 6. 8. 4-1 


4. 7. 2.0-1 


LIST OF REFERENCES (Continued) 

"Space Shuttle System Natural Environment Design 
Requirements Document," NASA Document dated 
May 17, 1971 

Terrestrial Environment (Climatic) Criteria Guidelines 
for Use in Space Vehicle Development, 1971 Revision, 
dated May 10, 1971, NASA TMX 64589 

Space Environment Criteria Guidelines for Use in Space 
Vehicle Development, 1969 Revision, Second Edition, 
dated August 26, 1971, NASA TMX 53957 

Private Communication with R. Smith/0. Smith, Aerospace 
Environment Division, Aerj-Astrodynamics Laboratory, 

MSFC, January 27, 1971 

Chapman, D. R., D. M. Kuehn and H. K. Larson, 

"Investigation of Separated Flows in Supersonic and Subsonic 
Streams with Emphasis on the Effect of Transition," NASA 
Report 1356, 1958 

Larson, H. K., "Heat Transfer in Separated Flows," 

Journal of the Aero/Space Sciences, November 1959 

"Astrionic System Study for Saturn S-II Expendable 
Second Stage," Phase B Study Report, dated 30 June 1971, 

IBM Number 71W-00225, MSFC-DRL-008A, Line Item Number 268 

Toda, Schlee and Obsharsky, "The Region of Kalman Filter 
Convergence for Several Autonomous Navigation Modes," 

AIAA Paper Number 67-623, IBM Corporation 

Vette, Lucero, and Wright, "Models of the Trapped Radiation 
Environment, Volume II Inner and Outer Zone Electrons," 

NASA SP-3024 

Vette, "Models of the Trapped Radiation Environment," 

Volume I Inner Zone Protons and Electrons," NASA SP-3024 

"ELBA - Electron and Bremsstrahlung Dose Rate Code," 

RSIC Computer Code Collection - CCG119, by Space Sciences 
Laboratory, Marshall Space Flight Center, Huntsville, Alabama 

Burrell, M. 0. and Wright, J. J., "Orbital Calculations 
and Trapped Radiation Mapping," NASA TMX-53406 

Kostoff, R. N., "Aerobraking of the Space Tug from Syn- 
chronous Orbit into Low Circular Earth Orbit: Guidance 

and Heating Constraints on First Atmospheric Pass," 

Bell comm Memorandum for File, July 15, 1971 


3 


D5-17142 


4. 7. 2.0- 2 

4.7. 2. 0- 3 

5. 1.2. 0- 1 


LIST OF REFERENCES (Continued) 

Kostoff, R. N.» "Space Shuttle and Space Tua Peak 
Temperatures," Case 237, Bell comm Memorandum for 
File, August 26, 1971 

Pratt and Whitney Aircraft "Ascent arid Re-Entry 
Heating RL-10/Space Tug," dated fey 28, 1971 

Alternate Space Shuttle Concepts Study, Final Report, 
Part II, Volume I, Grumman-Boeing, NAS9-11160, 

July 6, 1971 


D5-17142 


APPENDIX A 

TWO PASS AEROBRAKING SPACE TUG ANALYSIS 


A-1.0 INTRODUCTION AND SUMMARY 

The original contracted activity for the Space Tug Aerobraking Study 
was accomplished from May 27, 1971 through November 27, 1971. This 
study activity was directed to determine the feasibility of aerobraking 
and to define the maximum payload capability aerobraking Tug concept as 
a function of mission duration. The successful completion of this 
activity identified the necessity to investigate short duration mission 
payload data and to impact the effect of flare length on aerobraking 
Tug payload capability. 

This add-on contractual activity was initiated from January 12, 1972 
through April 12, 1972. The three-month activity included two months 
for technical activity with one additional month for incorporation of 
results into the final documentation. This activity investigated the 
following concepts: 

1. A basic (no flare) configuration - 2 pass mission. 

2. A "short" 60° flare configuration - 2 pass mission. 

3. A "short" 60° flare configuration - 30 pass mission. 

4. A "large" 80° nose flare configuration - 2 pass mission. 

The basic (no flare) Tug data (item 1. above) completed the spectrum of 
mission durations from the two' pass (approximately nine hours aero- 
braking return time) to a 60 pass (approximately 11 days aerobraking 
return time). The "short" 60° flare configuration data (items 2. and 
3. above) identified the performance and weights of low weight, high 
performance aerobraking configuration over the two pass to the 30 pass 
mission duration. The large nose flare configuration (item 4. above) 
was designed to achieve a low ballistic coefficient configuration which 
would result in significantly lower pressure loads and lower thermal 
environments than those experienced by the short flared configurations 
when performing the two pass mission. 

The background for the add-on effort was based on the desirability of 
very short duration missions which will minimize the Shuttle's on-orbit 
stay time and which would minimize the monitoring and/or tracking 
operations of ground stations. The initial study effort was designed 
to maximize payload capability of the aerobraking Tug concept without 
considering mission on-orbit duration constraints. 

Other constraints which increased mission duration included the ground- 
rule of 2000° F as state-of-the-art limit for radiative materials. As 
the temperatures for one to four pass missions were believed to exceed 


A-l 


D5-17142 


i 


A-1.0 (Continued) 

any near-term improvement in radiative material capabilities, short 
duration missions were not investigated. The use of ablative materials 
were not considered due to the potential problems associated with 
multiple heating/cooling cycling during aerobraking, ablative out- 
gassing, hot spots, cold soak and surface recession. The add-on study 
activity considered the potential problems associated with ablatives 
to be controllable with the proper material selection criteria and 
design features. 

The objectives of the add-on study were to; 

a. Determine the performance and round trip payload capability of 
a Space Tug utilizing a two pass aerobraking technique from 
synchronous orbit for the basic (no flare) Tug, the short 60° 
flare Tug and the large flare Tug. 

b. Determine the inert weight penalties associated with these con- 
figurations. 

c. Compare the Space Tug relative weights for conventional Tug 
with the new aerobraked Tug data points. 

d. Define the performance of a shortened 60° flare configuration 
to compare with the current 30° flare (30 pass mission). 

Figure A-l. 0.0-1 illustrates the concepts studied in the add-on effort. 
They include a basic (no flare) Tug with an ablative heat shield, a 
"short" 60° flare concept and a large nose flare concept. The basic 
(no flare) Tug for a two pass mission has the same configuration as the 
basic (no flare) Tug for the 30 pass mission. The heat shield, 
however, must use an ablative material rather than a radiative material 
as the temperatures encountered exceed the radiative materials' state- 
of-the-art capability. The sidewall insulation system requires the 
titanium outer foil be replaced by a L-605 material which has a higher 
thermal capability. 

The "short" 60° flare Tug concept for the two pass mission also re- 
quires an ablative heat shield. The sidewall protection system is 
L-605 over microquartz insulation. The short Inconel 718 flare will 
not totally shield the payload and will allow the payload to experience 
temperatures above the desired 300°F. Therefore, the payload adapter 
must be insulated. 

The large nose flare can be Inconel 718 or Rene' 41 as the flare will 
experience temperatures of 1300 to 1400°F. The sidewalls and payload 
are shielded by the flare and do not require insulation (see Appendix 
F). 

The results of the follow-on activity are reflected in the payload 
capabilities of the various configurations are as shown in Figure 
A-l. 0.0-2. The 2 pass basic (no flare) Tug will roundtrip five 


A- 2 


BASIC 

(MO FLAKE) 


SHORT 6<r 
FLARE 


LARGE NOSE 
FLARE 


PAYLQjkP 

AVIONICS 


» AY LOAD ADAPTER 


*' 9 ' AVIONICS | 60° 


AVIONICS 


PAYLOAD ADAPTER 




'"*r5ZT'' 


~St:~ 


xry 




ft w 


ASLATIVE AFT HEAT SHIELD 
(2- PASS MISSIONS ONLY) 


FIGURE A-l. 0.0-1: SELECTED SPACE TUG AEROBRAKING CONFIGURATION CONCEPTS 

(ADD-ON ACTIVITY) 




Short 60° Flare 
(Neutrally Stable) - 
(30 Pass) 

Light Weight Large Flare 
(50‘ Diameter) 

(2 Pass) 


1800 

(See Appendix F) 




Figure A-l. 0.0-2 Geosynchronous Round Trip Payloads 














D5-17142 


A-1.0 (Continued) 

hundred pounds in the geosynchronous mission (assuming the Tug returns 
to a 270 n.m. orbit and then rendezvous with the Shuttle at 100 n.m. 
orbit). If the basic (no flare) Tug departs from and returns to a 
200 n.m. orbit, this payload capability can be increased to approxi- 
mately 1700 pounds. For the 2 pass, 9 hour basic Tug mission, the pay- 
load will be reduced approximately 2300 pounds from that obtained in 
the 30 pass mission due to the increased thermal protection required on 
the heat shield and the Tug sidewalls. 

The short 60° flare has approximately the same small two-pass payload 
capability as does the basic (no flare). The use of a short flare did 
not reduce the thermal and pressure loads significantly. Therefore, 
the aerobraking kit components for the short 60° flare 2 pass concept 
were comparable in weight to that of the 2 pass basic Tug described 
above. As with the basic (no flare) Tug, increasing the rendezvous 
altitude to 200 n.m. orbit will increase the payload capability by an 
additional 1200 pounds to 1700 pounds. 

The large nose flare for the two pass mission has a negative payload in 
either mission mode (270-100 n.m. or 200 n.m.). Approximately a 2000 
pound inert weight reduction would be required to achieve a zero payload 
in the 270-100 n.m. mode or a 800 pound reduction in the 200 n.m. mode. 
The majority of the aerobraking kit weight for this concept is the large 
flare (approximately 99 percent of the weight). This design places the 
structure in compression. Alternative designs indicate that designing 
a tension flare configuration might significantly reduce weight and 
provide a positive payload capability as shown in Appendix F. 

Also shown in Figure A-1.0. 0-2 is the payload capability of the "short" 
60° flare flown in a 30 pass mission mode. This configuration was de- 
signed to have the same drag coefficient as the 30° flare configuration. 
In the initial study effort, the 30° flare configuration had greater 
payload capabilities than the 45° and 60° flare configurations. This 
was believed to be due to the selection of the flare length rather than 
the flare angle selected. The 60° flare, when designed to the same 
drag coefficient as the 30° flare, will be 4.9 feet long compared to 
the 11.25 feet long for the 30° flare. This will decrease the 60° 
flare weight significantly. However, its flare weight saving is offset 
by the requirements for (1) increased RCS fuel (to provide static 
stability) and (2) payload adapter insulation (much higher temperatures 
behind the flare). 

The 60° flare will be neutrally statically stable at 8.68 feet slant 
height. An assessment of payload capability indicated that the weight 
of the longer (8.68 feet) flare is less than the weight of RCS fuel 
and insulation required by the shorter flare (4.90 feet). The longer 
neutrally stable 60° flare's payload capability is shown for the 30 
pass mission (Figure A-1.0. 0-2). This neutrally stable 60° flare con- 
figuration will provide a slight improvement in the payload capability 
(approximately 180 pounds) over the 30° flared aerobraking Tug. 


A- 5 


D5-17142 


A-1.0 (Continued) 

The results of the short mission duration analysis show that it is 
possible to return payloads within a day. However, the payload' capa- 
bility decreases significantly as the mission duration decreases. 

Some technology advances are required to improve the ablative materials. 
Further, the design of large aerodynamic decelerators need to be studied 
in greater depth to reduce the high weight penalties of the designs 
examined to date. 

A-2.0 TRADE STUDIES 

This section of the appendix presents the results of the add-on study 
activity including aerodynamics, configurations, trajectories, control, 
astricnics, thermal, materials and weight analyses. 

A-2.1 AERODYNAMIC ANALYSIS 

Aerodynamic data was developed in accordance to the criteria shown in 
Section 4.1 (basic report). Data was developed to determine the drag 
coefficients for trajectory analysis, the static stability data for 
control analysis and the air loads for structural and weights analysis. 
This data was developed for the two pass missions with a basic (no 
flare) Tug, a Tug with a short 60° flare and a Tug with a large flare 
(low ballistic coefficient). Additional data was developed for the 
short 60° flare configuration flown in a 30 pass mission. This data 
was developed to provide comparable data to a similar 30 pass mission 
flown with a 30° flare configuration. 

Figure A- 2. 1.0-1 is a representation of the payload capability trend 
versus ballistic coefficient (W/Cqa) illustrating the data points de- 
veloped in the initial activity and identifies the recommended addi- 
tional data points. The preliminary assessment indicated that the 
maximum payload capability may be obtained with a combination of a 
short steep angle flare and a moderate degree of static instability. 

One reason for this assessment was to determine the best short 60° 
flare configuration for analysis during the add-on effort (i.e., two 
choices were available; a neutrally stable 60° flare configuration or 
a 60° flare configuration with the same W/CpA as the 30° flare con- 
figuration). 

Flow separation effects should be considered in detail for small, 
steep flare configurations. Flow separation should in general, (1) 
increase with increasing flare semi-vertex angle, (2) increase with 
decreasing perigee altitude (lower time to decay), and (3) increase, 
percentage wise with decreasing flare size. Follow-on in-depth study 
activity including wind tunnel testing must be performed to determine 
the optimum aerodynamic decelerator design. 


INCREASING NET PAYLOAD 


UNSTABLE 

CONFIGURATIONS 


■■■BCraL'IMH 

■■■■■■■■EES 




NEUTRAL 

STABILITY 

BOUNDARY 


••U4-I 

(NO FLARE) 


O - EXISTING DATA POINTS 
O “ RECOMMENDED ADDITIONAL DATA POINTS 
dy - FLARE ANGLE 


NO FLOW SEPARATION EFFECTS CONSIDERED 


EXPECT 
NEGATIVE 
PAYLOAD 
FOR W/C d A = 2 

CONFIGURATION 


(EQUIVALENT) W/C d A 


FIGURE A-2. 1.0-1: PRELIMINARY FLARE CONFIGURATION/NET PAYLOAD SENSITIVITY ESTIMATE 


D5-17142 




D5-17142 


A-2. 1.1 Basic (No Flare) Two Pass Analysis 

The possible ablation effects were not considered in the aerodynamics 
analysis of this configuration. The drag coefficients, normal force 
coefficients, and center of pressure data for the basic Tug are shown 
in Figures 4. 1.1. 1-2 and 4. 1.1. 2-1 (basic report). 

The local pressure coefficient ( c pL0CAL^ d '> s ' tr '* but ^ on f° r basic 
(no flare) configuration over the range of 2 to 60 passes is shown in 
Figure A-2. 1.1-1. The local pressure (P L ) for the two pass mission is 
more than double that for the five pass mission because of the 2 pass 
K n * factor. For example, the five pass maximum nose local pressure is 

less than 0.5 psi (Figure 4. 1.1. 3-2 of basic report) while the two pass 
value is slightly greater than 1.0 psi. 

A-2. 1.2 Short 60° Flare Analysis 


Three short 60° flares of varying slant heights were defined as shown 
in Figure A-2. 1.2-1. Short 60° flare #1, slant height = 0.26 caliber 
(3.64 feet) was sized to provide the same equivalent W/CJ\ as the 30° 
flare (the equivalent W/C Q A for a given configuration is u that constant 
W/CpA which would be required to result in the same initial target 
perigee altitude for the required time to decay). The drag character- 
istics of this configuration, presented in Figure A-2. 1,2-2 appeared 
too low to result in the same equivalent W/C Q A as the 30° flare config- 
uration. Hence, the short 60° flare #2 was defined. 


The short 60° flare #2 configuration (slant height = 0.35 caliber - 
4.90 feet) resulted in approximately the same 30 pass initial perigee 
as the short 30° flare configuration. The drag characteristics of 
this configuration are presented in Figure A-2. 1.2-3. This #2 config 
uration was selected as the baseline short 60° flare and was the 
configuration subjected to design, control, thermal, weights and pay- 
load analyses. This selection was based on the preliminary payload 
estimate (prior Figure A-2. 1.0-1). 


The static stability characteristics of the selected short 60° flare 
(#2) are shown in Figures A-2. 1.2-4 (C Not ) and -5 (CP/D). This con- 
figuration is statically unstable at altitudes below approximately 
400,000 feet. The local pressure coefficients over the heat shield 
and cylinder are the same as those of the 30° flare (Figure 4, 1.2. 5-1 
of basic report). The local pressure coefficients over the flare are 
shown in Figure A-2. 1.2-6. This later figure also shows the K n * 
factors to be used to convert the local coefficients to local 
pressures. 

The short 60° flare #3 configuration (slant height = 0.62 caliber - 
8.68 feet) was sized to the neutral stability characteristics of the 30° 
flare. The drag coefficient data for this flare is shown in 
Figure A-2. 1 .2-7. 


LOCAL PRESSURE COEFFICIENT (C 



D5-17142 


















y/d r * calibers perpendicular to centerline axis 


NOTES: 

© Z-l ELLIPSOID NOSE COMMON TO ALL CONFIGURATIONS 
© MAIN PROPULSION GIMBAL POINT 
© ASTRIONICS MODULE 
PAYLOAD ADAPTER 
PAYLOAD 

TUG DIAM. = D r a 14 FT. 



X/D R « CALIBERS AFT OF NOSE 


© 


CONFIG. S y 

SHORT #1 60° 

SHORT *2 60° 

SHORT # 3 60° 


SLANT LTH, - FT. 


3.64 

4.90 

8.68 


FIGURE A-2.1.2-1: SHORT 60° FLARED TUG CONFIGURATION(S) GEOMETRY 


D5-17142 


Ban 


"REF = 154 FT 
SLANT LTH = .26 CAL 


BRBB^ 


w. 


■■■■■■■■■—a WMwmmm ■■ 

■■■■■■■■■■■By ,«■■■■■■■■■■■■■■■■—■ 

BSiMMgMroMMMMaMmi 


ALT— KM 


FIGURE A-2. 1.2-2: DRAG COEFFICIENTS FOR SHORT 60° FLARE #1 











CP/D ~ CAL AFT OF NOSE 


D5-17142 


.35 CAL. 


% V= 60° 



X/D^CAL. AFT OF HOSE 


SLANT LTH. = .35 CAL. 

:*ref * ,m fj2 jztiziz: 

D REF = WT _ | ! ! ■ !H 4 

1$V = FLARE SEMIVERTEX ANGLE. 


35 KFPS 
30 KFPS 
25 KFPS 
20 KFPS 


mmammm 



■■■"I" "^iMibbibbbbbwbub 

atam 











ALT~ KM 


FIGURE A-2. 1.2-7: DRAG COEFFICIENTS FOR SHORT 60° FLARE #3 




D5-17142 


A-2. 1.3 Large Flare (W/C^A = 2) Analysis 

A spectrum of candidate configurations were sized to produce a ballistic 
coefficient of approximately 2 psf. The drag characteristics of a 
candidate W/CpA = 2 psf "ring tail" flare configuration were estimated 
as a function of altitude in order to determine the importance of rar- 
efied flow effects in increasing the effective configuration drag 
characteristics. The Cp = f (h ) for the candidate configuration is 
presented in Figure A-2, 1.3-1. Based on a preliminary two pass trajec- 
tory analysis, a possible reduction in flare planform area of ~ 10 per- 
cent was defined due to rarefied flow effects. 

Other configurations were investigated and sized based on the prelimi- 
nary data for the ring tail flare. Candidate configurations as shown 
in Figure A-2. 1.3-2 above were subjected to a preliminary design 
analysis to determine the packaging, deployment and retraction feasi- 
bility and the astrionic sensors' field-of-view characteristics. The 
forward facing blunted cone with a flare angle ( 6v) of 80° was se- 
lected. Figure A-2. 1.3-3 shows the drag coefficient data for this 
selected nose flare configuration. Velocity effects on the coefficients 
of drag were not considered because of the primary objective of deter- 
mining realistic weights for this very low ballistic coefficient con- 
figuration. The local pressure loads are approximately 5.3 psf over 
the entire area. The nose flare configuration should provide satis- 
factory static stability over the flight regimes of interest. 

A-2. 2 CONFIGURATION OPTIONS 

This section describes the aerobraking kit design modifications required 
for the additional missions investigated in the add-on activity. The 
add-on effort covered a two month technical activity and, therefore, was 
time restricted to define the feasibility and estimated performance of 
a few vehicle concepts for a few missions. The design activity was 
limited to defining reasonable configuration concepts, functional cap- 
abilities and weights. No attempt was made to optimize the weight or 
performance parameters. 

The groundrules used for the design portion of the add-on effort are 
identical to those used in the initial portion of the study and are 
listed in Section 4.2, Aerobraking Configuration Concepts (basic report). 


The 

aerobraking kit elements 

consist of: 

a. 

Aft Heat Shield 

(Section A-2. 2.1) 

b. 

Sidewall Insulation 

(Section A-2. 2. 2) 

c. 

Reaction Control System 

(Section A-2. 4) 


A-17 


ALT ~ KM 


FIGURE A-2.1.3-1: DRAG COEFFICIENTS FOR "RING TAIL" FLARE 







D5-17142 


i 


APPROX. CONFIG. GEOMETRY: SHAPE & SIZE 




2:1 ELLIPSOID DIAM. 
HEMISPHERE DIAM. ^95’ 



FORWARD 

FACING 

BLUNTED 

CONES 


®L 

6 V 

DIAM, 

1 

80° 

72' 

2 

70° 

73’ 

3 

60° 

78’ 






0P 

A p RANGE 

FORWARD 

90° 

1080-1200 FT' 

FACING 

80° 

1120- <240 FT 

PANEL 

70° 

1150-1280 FT 


60° 

1270-1410 FT 


D, *=D 2 

TRAILING 

“DONUT” 10* 192’ 

20' 96’ 

30* 64* 



TRAILING 
SPHERE 
(OR HEMI- 
SPHERE) 


DIAM.~90’ 


FIGURE A-2. 1.3-2: CANDIDATE CONFIGURATIONS FOR W/C 0 A = 2 PSF 

A-19 






D5-17142 


A-2.2 (Continued) 


d. Astrionics System 

(Section A-2.6) 

e . FI are 

(Section A-2.2. 3) 

f. Payload Adapters 

(Section A-2.2. 4) 

A-2.2.1 Aft Heat Shield 



The aft heat shield is required to protect the engine nozzle and engine 
systems from aerodynamic heating and pressure loads during atmospheric 
braking. For the longer duration missions, radiative heat shields may 
be used. However, for short duration missions, an ablative heat shield 
is required due to the high temperatures encountered. For the three 
configurations investigated as a part of this add-on activity, the 2 
pass mission basic {no flare) Tug and short 60° flare Tug will require 
an ablative heat shield. The large nose flare (2 pass mission) and the 
short 60° flare (30 pass mission) will use radiative heat shield con- 
cepts identical to that shown in Section 4.2.2, Aerobraking Configura- 
tion Concepts (basic report). 

The ablative heat shield was designed to be a cold structure, i.e., the 
backside of the shield was designed to be limited to 300°F. Further, a 
review of the literature on ablative materials indicated that the fixed 
heat shield with a nine foot movable cap over the engine nozzle used 
with the radiative heat shield concept cannot be used for ablative heat 
shields due to the critical gas flow effects at the movable cap joint. 
Sealing at this joint appeared to present a difficult design problem. 
Therefore, the approach selected used a one piece heat shield which had 
the joint at the aft skirt extension interface. The joint now would be 
located in the less critical gas flow region and, therefore, the tem- 
peratures encountered would be lower. 

The ablative heat shield concept is shown in Figure A-2.2. 1-1. It con- 
sists of a silicon/epoxy ablative {Martin-Marietta ablative material 
3560 II A) over an titanium supporting structure. The ablative thickness 
varies from 3.3 inches at the center (stagnation point) to 1.0 inches 
at the interface with the sidewall. The titanium structure consists of /; 
a ring frame at the sidewall joint, a 2:1 elliptical titanium skin .060 
thick and 6 supporting titanium angle braces. The actuation system is 
supported by a set of "I” beam braces located 7.25 feet apart. 

The heat shield is opened or closed by a signal to the electrical drive 
motor. Two driver gears mounted 7.25 feet apart on the drive shaft 
rotate the two hinge links approximately 82°. During this operation, 
the dome moves up and away from the ring frame at point C (see figure). 
The portion of the dome farthest from the hinge slides down the ring 
frame and disengages the 2 locking clips. After the hinge has rotated 
the 82°, the complete assembly then rotates about point A (see figure) 
until the dome can lock into the Tug sidewall at point B (see figure). 

A solenoid driven pin locks the dome to the sidewall during conventional 
Tug engine operations. The closure method reverses this process. 


A-21 




MfUttHMTMOM 


FIGURE A-2.2. 


■MAT MSLft 

tnucnacMD m*w« 

CMtt ilCTtQM 



-1 : ABLATIVE HEAT SHIELD CONCEPT 




D5-17142 


A-2.2.1 (Continued) 

The location of the heat shield along the side of the Tug will impact 
the Tug c.g. location and must be compensated for during conventional 
Tug operations with the gimballing of the main engine nozzle. This 
will result in some performance loss. 

A-2.2.2 Sidewall Insulation 

The sidewall insulation for the longer duration missions (i.e., short 
60° flare 30 pass mission) consisted of a titanium outer skin over a 
microquartz insulation. The thickness of microquartz was designed to 
maintain 400°F at the micrometeoroid shield located under the micro- 
quartz. The temperature of the Tug sidewall decreases going from the 
heat shield/cylindrical wall joint toward the payload. The microquartz 
insulation thickness is tapered downward accordingly. 

For the large nose flare, the flare shields the sidewall and no sidewall 
thermal protection is required. 

For the basic (no flare) Tug and the short 60° flare configuration 
(when used in a 2 pass mission mode), the sidewall temperatures exceed 
the titanium outer foil capabilities and L-605, a Haynes cobalt alloy 
replaces the titanium. The temperature on the payload exceeds 300° F 
and must be thermally protected. The sidewall TPS is extended to cover 
these areas. A payload aft closure must be provided also. A cover of 
aluminum (facing the payload) bonded to microquartz which in turn is 
covered by L-605 outer foil may be used. 

A-2.2.3 Flare 

Two of the three configurations investigated in this add-on study em- 
ployed flares to increase the drag and to provide some improvement in 
static stability. 

The short 60° flare used for the 2 pass and 30 pass missions were de- 
signed to use much of the flare concept proposed in Section 4.2. 2.3 
(basic report). The thermal environments range from approximately 
850°F (30 pass mission) to approximately 1500°F (2 pass mission). The 
previously selected material. Inconel 718, may be used for both flare 
concepts as the inconel has temperature capabilities which exceed the 
temperatures encountered. The support struts may be titanium as the 
temperature directly beneath the flare will not exceed titanium's 
capabi 1 i ty . 

The short 60° flare concept, as shown in Figure A-2.2.3-1, will consist 
of 24 large panels (.020 inches thick) and 24 small panels. The 4.9 
foot long flare is elevated to a 60° angle by the support strut. The 
12 support struts are elevated in turn by threaded rods and followers. 

A reversible drive motor, a drive chain and 12 drive sprockets provide 
the actuation for the support struts. The flare is hinged at the pay- 
load adapter/astrionics module joint. Ring frames in the payload 


A-23 







D5-17142 


A-2.2.3 (Continued) 

adapter and located at forward payload adapter face and at the support 
threaded rod act as "stops" to counteract flare loads. The flare when 
retracted will overlap the large panels. At the ends of the panels 
where the panel circumference is significantly larger than the Tug 
circumference, the smaller panels fold under the large panels to 
facilitate packaging of the Tug with flare into a 15 foot diameter 
Shuttle cargo bay constraint. 

The large nose flare concept is a unique design to provide a ballistic 
coefficient of approximate 2 pounds per square foot. For this concept, 
the flare was placed forward of the propulsion module so that it could 
be combined with the heat shield to form a continuous forward aero- 
dynamic flare as shown in Figure A-2. 2.3-2. The temperatures en- 
countered with the large flare are approximately 1350 to 1400°F and 
will permit the use of a radiative heat shield of Rene' 41. The large 
nose flare, however, does present a difficult packaging problem. When 
extended to its 80° angle, the flare is approximately 72 feet in 
diameter. This large flare must be folded (with considerable over- 
lapping of panels) to allow it to fit within the Shuttle bay. The 
concept, as shown in Figure A-2. 2.3-3, has 24 flare panels, 29 feet 
long. The panels are hinged to the Tug at the start of the cylindrical 
section of the propulsion module. The panels are tapered to provide 
minimum overlap at the forward end and sufficient surface to meet the 
large areas at the aft end of the flare. The flare panels are extended 
first and then the support system is positioned behind, but not 
connected to the flare panels. The nose section of the large nose flare 
corresponds to the radiative heat shields design. (See Section 4.2.2, 
Figure 4. 2. 2. 1-1, Aft Heat Shield, basic report). 

The flare panels are folded in 3 packets of 8 panels each (24 total). 

As shown in Figure A-2. 2. 2-3 on flare deployment sequence, three panels, 
120° apart are deployed first. All of the panels have a lip along 
longitudinal edges which permit the panels to innerconnect. The lips 
also will facilitate the first set of panels to release the second set 
of panels, etc. 

After all of the panels have been released, the support system is 
activated. The support system consists of 12 support rods equally spaced. 
Each rod extends down the length of the Tug sidewall when retracted. 

When extended, the aft portion of each of the 12 rods is pulled forward 
by a motor driven cable system. The aft rod is 30 feet long, 2" o.d. 
aluminum tubing. The forward leg is a 1-1/4" o.d. aluminum tubing rod. 
The upper hinge point between the forward and aft rod has a 1/16" 
diameter steel cable which is connected to each of the other support 
legs at their hinge point. The rods plus the cables thus provide support 
to the extended flare. Because of the loads, each of the 12 aft support 
legs will require a tripod support at their mid-point approximately 
3 inches out from the center of the support rod. Steel cable will run 
over the tripod from the front to the back of the support rod. 


A- 25 


A-26 



D5-17H2 




A-27 



FLARE EXTENDED 
AFT VIEW 



FIGURE A-2.2.3-3: LARGE 





| 



III 

r 

V 


FLARE DEPLOYMENT SEQUENCE 

JOSE FLARE DEPLOYMENT/RETRACTION 




D5-17142 


A-2.2.3 (Continued) 

The retraction method is as follows. First, the support rods are re- 
tracted. The retracted rods can be spaced to fit between the astrionics 
sensor system parts so as not to hamper guidance visibility require- 
ments. After the support system has been retracted, the flare panels 
can be retracted. Each of the panels will have a steel cable connected 
to the back side of the panels. These cables are retracted sequentially 
which in turn retract the panels. The above design for a large nose 
flare for the Tug is only of several possible alternatives investigated 
to reduce the ballistic coefficient. Others may offer lower weight and 
should be investigated in future studies. One of these alternatives is 
the use of a large "doughnut" attached aft of the Tug and tied to the 
Tug by cabling. This mode permits a very large flare which can reduce 
the temperature on the flare to where light weight plastic and/or 
silicon rubber materials may be used. Two of the key problems asso- 
ciated with this concept are (1) packaging the flare within the Shuttle 
dimensional constraints and (2) providing rigidity to the "doughnut" 
flare during maneuvers. 

Another concept which appears feasible is to employ a forward brake 
with an integrated dome concept. A very thin mesh material which can 
be extended into an umbrella like configuration provides a low weight 
drag structure. The material is held in place by the use of a torus 
shape inflatable ring located at the end of the umbrella mesh material. 
The end of the mesh material is wrapped over the ring and its interior 
is insulated to protect the ring. The ring is pressurized by helium 
or hydrogen to a pressure of approximately 9.2 psi. The ring is held 
in position by a series of cables extending from the ring to the Tug 
sidewall similar to the spokes on a bicycle wheel. This concept, while 
offering light weight, requires a complex retraction system. However, 
the reduced weight potential presents a valid reason for further in- 
vestigation of this concept and a preliminary analyses of the data is 
shown in Appendix F, 

A-2.2.4 Payload Adapter 

The payload adapter configuration for the 2 pass missions is identical 
to that proposed in Section 4. 2. 2. 4 (basic report). The thermal pro- 
tection systems for the payload adapter were described in Paragraph 
A-2.2.2 above. 

A-2.3 TRAJECTORY ANALYSIS 

The trajectory data generated in the add-on activity consisted of the 
f ol 1 owi ng : 

o Two and 30 pass trajectories for the short 60° flare 
(#2) configuration with the same W/CqA as the 30° 
flare. 


D5-17142 


A-2.3 {Continued) 

o Two and 30 pass trajectories for the short 60° 
flare (#3) configuration sized for neutral 
static stability. 

o Two pass trajectories for the large "ring tail" 
and "nose" flares (W/C^A «=* 2) . 

The assumptions, groundrules and methodology used were the same as re- 
ported in Section 4.3 (basic report). The effects of atmospheric and 
gravitational perturbations on the configurations or number of passes 
listed above were not considered in the add-on activity. 

The two pass basic (no flare) trajectory data was generated during the 
initial study activity. Figure 4. 3. 3. 0-1 (basic report) shows that the 
basic (no flare) initial target perigee altitude for the two pass mission 
is approximately 233,000 feet. The aerobraking return time is approxi- 
mately 0.39 days (the same as other two pass configurations). 

The drag coefficients used throughout the entire study (including the 
add-on activity) are not constant but depend on altitude and velocity. 

The equivalent W/CqA was defined to aid in the comparison of the con- 
figurations. Equivalent W/C^A of a configuration is that constant 

W/C n A, not a function of altitude or velocity, which would produce the 
same initial perigee for a given number, of passes as the actual W/CgA. 

A-2,3.1 Short 60° Flare Analysis 

Figure A-2.3. 1-1 shows the maximum dynamic pressures for the two short 
60° flare configurations (#2 and #3). The two pass data is plotted as 
small circles (2nd pass values). The 30 pass data is plotted as solid 
lines and indicates the normal trend of maximum dynamic pressure 
occurring on the last pass. The two pass values are approximately ten 
times greater than the maximums experienced during the 30 pass mission. 

The initial perigee altitude for these two configurations and the two 
mission durations are shown in Figure A-2.3. 1-2 below. 


CONFIGURATION/ 
NO. OF PASSES 

INITIAL PERIGEE ALTITUDE 
(FEET) 

Short 60° #2 

“ " J ' ' ■ 

2 

241 ,000 

30 

298,000 

Short 60° §3 


2 

251 .000 

30 

308,000 


Figure A-2.3.1-2 SHORT 60° FLARE INITIAL PERIGEE ALTITUDES 


A-29 


MAXIMUM DYNAMIC PRESSURE - PSF 




v 



FIGURE A— 2.3. 1-1 : SHORT 60° FLARE MAXIMUM DYNAMIC PRESSURE 


D5-17142 


05-17142 


A 


r 


A-2.3.2 Large Flare Analysis (W/CqA*^) 

As discussed in Section A-2.1 above, the ‘'ring tall" flare was used as 
the initial candidate large flare to determine the flow regime effects* 
This flare had an initial perigee altitude of approximately 302*000 
feet (two pass mission). The maximum dynamic pressure encountered on 
the second pass was 2 psf. The "ring tail" flare had an actual equiva- 
lent W/CpA of 1.78 psf which indicated that the size of such a flare 

could be reduced to obtain the desired value of 2 psf. 

The two pass nose flare configuration had an Initial perigee altitude 
of 300,000 feet and a 2nd pass maximum dynamic pressure of 2.5 psf. 

The actual equivalent W/CgA value was 2.02 psf. 

A-2.3.3 Equivalent Ballistic Coefficient Comparison 

Figure A-2.3.3-1 shows a summation of the equivalent ballistic co- 
efficients and the initial perigee altitudes for the configurations used 
in the overall study. Data from the main study is cross plotted with 
the new add-on activity data for comparison. The equivalent W/CqA's 

for any particular configuration are functions of the number of passes 
in the mission. In general, the equivalent W/CpA value for a configura- 
tion increases with decreased trip times. For example, the 45° flare 
has an equivalent W/C^A of 12.5 for 60 passes and 22 for two passes. 

The short 60° flare (#2) and the 30° flare have nearly equivalent drag 
characteristics in terms of initial perigee altitude and equivalent 
W/CpA for the 2 and 30 pass missions. The selection of this short 60° 

flare conforms to the desired flare W/C Q A comparison analysis activity 
planned for this study. 

A-2.4 CONTROL ANALYSIS 

A control analysis was conducted on the short 60° flare (#2) for the 2 
and 30 pass missions, the basic Tug 2 pass mission, and on the large 
nose flare for the 2 pass mission. The assumptions, groundrules and 
methodology reported in Section 4.4 (basic report) were utilized in 
the add-on activity. 

The short 60° flare is statically unstable. The Reaction Control 
System (RCS) was utilized to provide the required stability similar to 
that reported for the basic (no flare) configuration in Section 4.4 
(basic report). 

The basic (no flare) configuration's 2 pass maximum aeromoment was 
greater than the current 200 pound RCS thrusters could overcome. 
Therefore, the pitch/yaw RCS thrusters were resized! to 250 pounds each 
(for the two pass basic configuration only). This increase in 
thrust level was sufficient to offset the aeromoment. 



A-31 












D5-17142 


A-2. 4 (Continued) 

The large nose flare configuration is statically stable. The RCS, for 
this configuration, was only required to provide the limit cycle and 
directional control impulses. 

Figure A-2. 4.0-1 shows the RCS propellant consumptions for the three 
configurati ons. 

A-2. 5 THERMAL ANALYSIS 

The thermal analysis was conducted on the following configurations: 

o Basic (No Flare) - 2 passes 

o Short 60° Flare (#2) - 2 and 30 passes 

o Large Nose Flare - 2 passes 

The assumptions, groundrules and methodology used in the add-on activity 
were extensions of those reported in Section 4.5 (bas'c report). The 
CHAP computer program also has the capability of accommodating ablative 
materials, such as the ESA-3560 I I A selected for tha 2 pass mission 
basic (no flare) and short 60° flare aft heat shields. 

A-2. 5.1 Heating Rates 

Figure A-2. 5. 1-1 shows the heating rate distribution for the 2 pass 
basic (no flare) configuration. Figures A-2. 5. 1-2 and -3 show similar 
data for the 2 and 30 pass short 60° flare. Figure A-2, 5. 1-4 illustrates 
the very low heating rates associated with the 2 pass low W/CJ\ nose 
flare. 

A-2. 5. 2 Maximum Equilibrium Temperatures 

Figure A-2. 5. 2-1 shows the maximum equilibrium temperatures for the 
three configurations. The 2 pass basic (no flare) configuration's nose 
temperature (A) exceeds the limit for all re-radiative materials 
discussed in Section 4.7 (basic report). This high temperature, in 
conjunction with the high heating rates and total heat input shown in 
prior Figure A-2. 5. 1-1, established the criteria for the selection of 
the relatively dense ablative material for the aft heat shield. The 
sidewall and payload temperatures of the 2 pass basic (no flare) con- 
figuration are also higher, requiring more insulation and a heavier 
insulation outer foil (L605 Haynes rather than Titanium). 

The short 60° flare has lower temperatures on the nose and along the 
sidewalls than the basic (no flare) configuration. The 2 pass short 
60° flare has an ablative heat shield to withstand the 3290 degree 
temperature. It also uses the L605 Haynes alloy as the sidewall in- 
sulation outer foil (2 passes). This short 4.9 foot flare does not 
provide the environmental protection to the payload area that the other 
(larger) flares (basic report) provided. Therefore, insulation in this 


D5-17142 



STATIC 

STABILITY 

RCS PROPELLANT CONSUMED (LBS.) 

CONFIGURATION 

2 PASS 

30 PASS 

Short 60° Flare 

Unstable 

256 

271 

Basic 

Unstable 

516 

620 

(No Flare) 
Nose Flare 

Stable 

(250 lb. Thrusters) 
6 

i 

| 

(200 lb. 
Thrusters) 


Figure A-2.4.0-1 RCS Propellant Consumption 


A- 34 




HEATING RATE - BTU/F 



FIGURE A-2.5.1-1 KEATING RATE DISTRIBUTION - BASIC CONFIGURATION 












HEATING RATE - BTU/FT 2 SEC 


7 


V 




FIGURE A-2.5.1-3 FIEATING PATE DISTRIBUTION - 60° SHORT FLARE (30-PASS) 


D5-77142 


HEATING RATF. - BTI l/FT 2 SEC 



FIGURE A-2.5.1-4 HEATING RATE DISTRIBUTION - LARGE FLARE (2 -PASS) 


D5-17142 




CONFIGURATION 


TRAJECTORY 




A 

BASIC (NO FLARE) 

2-PASS 

3680 

SHORT 60° FLARE 

2-PASS 

3290 

SHORT GO 0 FLARE 

30-PASS 

2120 

LARGE NOSE FLARE 

2-PASS 

1410 


10 FT 


MAXIMUM EQUILIBRIUM TEMPERATURE (°F) 


B 

C 

D 

E 

F 

G 

H 

3520 

1789 

1275 



_ 

1170 

779 

3140 

1570 

1240 

1490 

1527 

1138 

758 

2070 

980 

748 

856 

889 

687 

458 

1403 

1380 

1337 

— 

— 


— 


MAXIMUM EQUILIBRIUM TEMPERATURES 


D5-17142 























D5-17142 


A-2.5.2 (Continued) 

region is provided similar to that described for the basic (no flare) 
in Section 4.2 (basic report). 

A-2.5.3 Thermal Protection System 

The aft heat shield ablative material thicknesses are shown in Figure 
A-2.5.3-1. The material thickness was based on 11,500 BTU's input on 
the first pass and 13,500 BTU's input on the second pass. To this, 
sufficient ablative material was used to obtain a 300°F temperature on 
the titanium support structure. The insulation was tapered to match 
the lower heat input on the dome as the contour matches the sidewall. 

Figure A-2.5.3-2 shows the sidewall insulation thicknesses required for 
the 2 pass basic (no flare) and short 60° flare configurations. As 
discussed in A-2.5.2 above, the sidewall insulation extends the entire 
length of the Tug sidewall including the payload to ensure that the 
maximum payload temperature does not exceed 300°F. The micrometeoroid 
shield of the Tug stage is protected to 400°F. The insulation material 
utilized is the same microquartz as reported in Section 4.2 (basic 
report). 

The weight statement for the thermal protection system is shown in 
Figure A-2.5.3-3. The 2 pass basic (no flare) and short 60° flare 
configurations required extensive thermal protection and this is 
reflected in their relatively heavy TPS weights of 3240 and 3125 
pounds, respectively. 

The large nose flare effectively protects the remainder of the Tug and 
payload and, therefore, no thermal protection system penalty is in- 
curred for this configuration. 

A-2.6 ASTRI0NICS ANALYSIS 

No additional astrionics analysis was accomplished during the add-on 
activity. The astrionic configuration and weights are shown in 
Section 4.6 (basic report). Section 4.6. 4. 5 (basic report) discusses 
a "quick-look" assessment of the one pass mission. The results of that 
assessment indicated that the perigee position uncertainties (approxi- 
mately, 35 NM) were similar to those of the longer mission durations. 

One major astrionics concern about the one pass mission was the apogee 
position uncertainty subsequent to the single atmospheric pass 
(approximately 5 NM with use of the horizon sensor). However, as in- 
dicated in Figures 4.6.4. 5-4 and -5 (basic report), while the second 
apogee uncertainty is always greatest, the second perigee uncertainty 
is less than the first. Therefore, apogee uncertainty is not expected 
to significantly impact the 2 pass mission. 

Another major concern was the sensitivity of deboost velocity error to 
initial perigee altitude. The sensitivity of initial perigee altitude 
to deboost velocity is approximately 2700 feet/ft. /sec. For the 2 pass 











INSULATION THICKNESS - . INCHES 


05-17142 



FIGURE A-2.5.3-2 S W:E TUG SIDEWALL INSLUT, 'ON THICKNESS 


A- 43 


CONFIGURATION 

TRAJECTORY 

SIDEWALL TPS WEIGHT (LBS) 

! HEAT SHIELD 
TPS MATERIAL 
WEIGHT (LBS) 

TOTAL TPS 
WEIGHT (LBS) 

RE-RADIATION 

SHIELD 

MICROQUARTZ 

INSULATION 

TOTAL 

Basic (No Flare) 

2-Pass 

247 (1) 

613 

860 (3) 

2380 (4) 

3240 

Short 60° Flare 

2-Pass 

247 (1) 

573 

820 (3) 

2305 (4) 

3125 

Short 60° Flare 

30-Pass 

119 (2) 

348 

467 (3) 

- 

467 

Large Nose Flare 

2-Pass 

0 

0 

0 

- 

i 


NOTES: 


1. Re-radiation shield - L605 Haynes Alloy (0.002 inch thickness) 

2. Re-radiation shield - Titanium (0.002 inch thickness) 

3. Weights include re-radiation shield and microquartz insulation on Tug 
sidewall, astrionics module and payload section. 

4. Heat shield material - ablative type, Martin-Marietta ESA-3560 IIA 


FIGURE A-2.5.3-3 THERMAL PROTECTION SYSTEM WEIGHT SUMMARY 


D5-17142 





D5-17142 


A-2.6 (Continued) 

basic (no flare) configuration, the target perigee altitude is approxi- 
mately 20,000 feet above the altitude resulting in a direct re-entry. 

The other 2 pass configurations have higher initial perigee altitudes. 
Therefore, it appears that velocity could be controlled sufficiently to 
avoid direct re-entry. As discussed in Section 4.6 (basic report), a 
detailed navigation and guidance analysis would be required in a follow- 
on activity to determine the full impact on the astrionics system 
caused by very short duration missions. 

A-2.7 AEROBRAKING KIT MATERIALS 

The materials used for the two pass mission are the same as those de- 
fined in Section 4.7, Aerobraking Kit Materials Selection (basic report). 
The only differences are that the use of ablatives for heat shield 
applications and higher temperature outer foil materials for the side- 
wall insulation system were investigated. 

The maximum stagnation point heating rates (thermal environment) for the 
2 pass missions are shown in Figure A-2.7. 0-1 below. 


CONFIGURATION/PASS NO. 

MAXIMUM HEATING RATE 
(BTU/ft 2 /sec.) 

Basic (No Flare) 


1 

127 

2 

94 

Short 60° Flare 


1 

85 

2 

63 

Large Nose Flare 


1 

5.3 

2 

3.9 


Figure A-2.7. 0-1 2 Pass Maximum Heating Rates 

The total heat input nto the 2 pass basic (no flare) Tug is approxi- 
mately 11,500 BTU's for the first pass and 13,500 BTU's for the second 
pass. These heat inputs are considerably higher than those encountered 
by the Space Shuttle. A low density ablator (approximately 20-40 
pounds per cubic foot, such as those used on the Shuttle) would not be 
satisfactory for this application. Materials selected for further re- 
view included (1) ESA-3560 IIA, (2) SLA-561 , (3) DC-93104, (4) DC-325, 
and (5) DC-93072. 


A-44 


D5-17142 


A-2.7 (Continued) 

The ESA-3560 1 1 A was selected for the ablative heat shield as it offered 
low density (56 pounds/ft3) compared to the 70 pounds/ft3 and 90 pounds/ 
ft3 of the DC-93072 and DC-93104 silicon phenolics. The SLA-561 had the 
desirable low density of 14.5 pounds/ft3; however, its heating rate capa- 
bility is not sufficient for the high heating rates encountered with 
two pass aerobraking. Figures A-2. 7.0-2 and -3 illustrate the properties 
of the selected materials. 

For the large nose flare, the temperatures encountered are approximately 
1350 to 1400°F. A radiative heat shield material was used. Either 
Inconel 718 or Rene 1 41 are acceptable (Rene '41 was used as it has 
higher strength to density properties at the flare temperatures). The 
properties of these materials are shown in Section 4.7 (basic report). 

The short 60° flare configuration's heat shield will encounter tempera- 
tures of approximately 2000°F (30 pass mission). For this configuration 
and mission duration, the TD-nickel -chrome radiative heat shield 
material was used. 

For the sidewall protection system, a microquartz insulation was used 
with a metallic outer foil. For the two pass basic (no flare) and 
short 60° flared Tugs, the sidewall temperatures are higher than en- 
countered in the longer duration missions. These higher temperatures 
necessitate the use of a high temperature L-605, a Haynes Cobalt alloy 
to replace the titanium outer foil previously used. Figure A-2. 7. 0-4 
lists the properties of the L-605. The short 60° flare configuration 
flown in a 30 pass mission can use the titanium outer foil to cover 
the microquartz. 

The flare of the large nose flare configuration shields the body of the 
Tug and the temperatures on the micrometeoroid shield do not exceed the 
capability of the aluminum shield. Therefore, no thermal protection 
was required for the large flare configuration sidewalls. 

The flare material for the short 60° flare does not exceed the capabil- 
ities of the Inconel 718 or Rene' 41 for either the 2 pass (temperature- 
1400“1500°F) or the 30 pass (800-900°F). Either of these materials, 
therefore, can be used for the flare. 

The payload adapter for the two pass mission used the same materials as 
the previous aerobraked configurations. Additional microquartz with 
the L-605 outer foil were required for the sidewall and payload for 
both the two pass basic (no flare) and short 60° flare Tug configura- 
tions. 


A-2. 8 WEIGHTS AND MASS PROPERTIES 

The weights and mass properties for the three configurations studied in 
the add-on effort are presented in this section. The three configura- 
tions were (1) the basic (no flare) Tug (two pass mission), (2) the 
short 60° flare (two and 30 pass missions) and (3) the large nose flare 
(two pass mission). 


A-46 


~<r 


j> 



THERMO PHYSICAL 

PROPERTY 

VIRGIN MATERIAL 
@ 80°F 

CHAR MATERIAL 
@ 1500°F 

SHUTTLE 

APPLICABILITY 

DENSITY (PCF) 

56.0 

2L3 

REQUIRE TEST 
VERIFICATIONS 
FOR SHUTTLE 
ENVIRONMENT 

THERMAL CONDUCTIVITY 
(BTU/IN-SEC-OF) 

1. 9 x 10' 6 

5. 9 X 10' 6 

SPECIFIC HEAT 
(BTU/LB-OF) 

0.26 

0.27 

EMISSIVITY 

0.75 

0.67 

MECHANICAL 

ULT. TENSILE STRENGTH (PSD 

245 

— 

APPLICABLE 

ULT. TENSILE STRAIN 
(IN. /IN.) 

0.062 


MODULUS OF ELASTICITY 
(PSD 

5000 

— 


PROJECT SOURCE: X-15A-2 FLIGHTS 


FIGURE A-2.7.0-2: ESA-3560 IIA PROPERTIES 


D5-17142 




A-47 


r~ ' 




LEADING EDGE LAMINAR FLOW 
DIRECTLY BONDED ABLATOR 
ABLATOR DENSITY = 56.0 LB/FT 3 


ADD: 


ABLATOR 

THICKNESS 

(INCHES) 


0.02 psf HARDCOAT 
0.05 psf BOND 



MAXIMUM STRUCTURAL 
TEMPERATURE = 300^ 

BACKFACE SMEARED 
THICKNESSES 

ALUMINUM* 1 .06” 
TITANIUM** .05” 


Q T ~ TOTAL LOCAL HEAT (ENTRY) (BTU/FT 2 ) 


FIGURE A-2.7.0-3: 


ESA-3560 I I A THICKNESS VS. HEAT INPUT 


D5- 17142 




D5-17142 


(L-605 ALLOY) 


SPECIFICATION: L-605 Sheet and Bar. 


CHARACTERISTICS: L-605 is a heat-resistant material similar to 

Stellite 31. It possesses high strength and 
oxidation resistance up to 2000°F. Ductility 
appears to be superior to the other high 
temperature alloys. Available as sheet bar, 
plate, wire and tubing. 

APPLICATIONS: Primarily for afterburner parts requiring high 

strength up to 2000°F. 


COMPOSITION: 0.15 C max, 19.0 - 21.0 Cr, 14.0 - 16.0 W, 

9.0 - 11.0 Ni, 2.0 Fe, max, 1.0 Si max, 

1. 0-2.0 Mn, bal Co. 

MECHANICAL PROPERTIES: 


Sheet 

Condition Anneal ed 

Tensile, PSI 155,000 

Yield, 0.2% offset 70,000 

% Elongation in 2 inch 55.0 

% Reduction of Area 40-45 

Brinell (3000 Kg) 218-228 


Sheet 
As -rolled 
170,000 
108,500 
40.0 

305-330 


PHYSICAL PROPERTIES: 


Density 

Expansion Coef ( 70-600° F ) 
Thermal Conductivity 
Scaling Temperature 


9.15 g/cc; 0.330 Ib/cu. in. 

7.6 x 10-6 in/in/°F 
100 BTU/sq.ft./hr/°F/in. 
2000° F 


FABRICATION: 

Forming, Good 
Welding, Good 
Machining, Good 


FIGURE A-2.7.0-4 L-605 COBALT ALLOY MATERIAL PROPERTIES 


A-48 


05-17142 


A-2.8 (Continued) 

The mass properties used in the aerodynamic and the control analyses are 
shown in Figure A-2.8, 0-1. The mass properties as shown in the figure 
were calculated for the conditions existing at the start of the aero- 
braking return from geosynchronous orbit. 

The inert weight associated with the implementation of these aero- 
braking return modes of operation are summarized in Figure A-2.8. 0-2. 

The structural weights for the aerobraking kit elements for the basic 
(no flare) Tug and for the short 60° flare Tug were determined by a 
detailed sizing analyses of the heat shield and flares for the 
pressure and thermal environments. The large nose flare was subjected 
to a less detailed sizing; however, the weight estimates were conducted 
in sufficient detail to determine that the large nose flare concept in- 
vestigated has no geosynchronous payload capability. 

The materials used for each of the four configuration aerobraking kits 
are as shown in Figure A-2, 8.0-3. 

The total Tug weight at the start of aerobraking for two of the three 
configurations are spotted on Figure A-2. 8. 0-4 for comparison with the 
previously reported weights data shown in Section 4.8 (Figure 4. 8. 6. 0-1, 
basic report). Note that for the large nose flare, no payload was 
obtained so that this point was not plotted. The two pass mission' data 
shows a rapid increase in aerobraking kit weight (and corresponding 
lower payload capability) with shorter duration missions. The short 
60° flare configuration has approximately 40 pounds more aerobraking 
kit weight for the 30 pass mission than the 30° flare. 

A-3.0 SENSITIVITIES AND PAYLOADS 

This section discusses (1) the revised atmospheric dispersion targeting 
scheme and model, (2) a typical total mission navigation error correc- 
tion burn sensitivity, (3) the payload capabilities of the configurations 
studies in the add-on activity, and (4) the performance sensitivity 
comparisons between the initial MSFC Point Design Tug and the Aerobraked 
Tug. The first two topics listed above are presented to provide initial 
insight into certain problem areas identified in Section 5 of the basic 
report. The third topic is a continuation of the add-on activity report. 
The fourth topic complements the conventional -aerobraked Tug comparisons 
of Section 5.3 of the basic report. 

A-3.1 REVISED ATMOSPHERIC DISPERSIONS 

Figure 5. 2. 1.0-2 of the basic report shows the range of atmospheric per- 
turbations used in the main study. This density range was approximately 
+50% and -40% from the 1962 Standard Atmosphere. The expected variations 
during a mission or between individual passes were not given in the NASA 
Atmospheric Model. This gap in the atmospheric model data was discussed 
in the basic report and with the Space Environment personnel of MSFC. 

In response, the revised atmosphere model shown in Figure A-3.1. 0-1 was 


D5-17142 



BASIC 
(NO FLARE) 

SHORT 
60° FLARE 

LARGE 

FLARE 

WEIGHT (POUNDS) 

14,430 

14,430 

14,430 

CENTER OF GRAVITY X (INCHES) 

243.3 

233.5 

158.1 

CENTER OF GRAVITY Y (INCHES) 

0 

0 

0 

CENTER OF GRAVITY Z (INCHES) 

0 

0 

0 

ROLL MOMENT OF INERTIA (SL.-FT 2 ) 

10,388 

12,700 

64,000 

PITCH MOMENT OF INERTIA (SL.-FT 2 ) 

101,686 

94,000 

101,000 

YAW MOMENT OF INERTIA (SL.-FT 2 ) 

101,686 

94,000 

101,000 


FIGURE A-2.8.0-1 MASS PROPERTIES FOR ADD-ON ACTIVITY 


D5-77142 


COMPONENT 

CONFIGURATION WEIGHT (POUNDS) 


BASIC 



LARGE 


{NO FLARE) 

SHORT 60° FLARE 

NOSE FLARE 


2 PASS 

2 PASS 

30 PASS 

2 PASS 

AFT HEAT SHIELD 

2785 

2700 

480 

400 

FLARE 

— 

510 

380 

6580 

*S I DEWALL INSULATION 

925 

881 

513 

— 

PAY LOAD/ FLARE ADAPTER 

350 

390 

390 

350 

AST RI ON ICS PENALTY 

25 

25 

325 

25 

ADDED RCS INERTS 

15 

— 

— 

— 

RCS PROPELLANT 

516 

256 

271 

6 

TOTAL 

4616 

4762 

2359 

7361 


INCLUDES PAYLOAD INSULATION AND PAYLOAD CAP ACTUATION DEVICE 


FIGURE A-2. 8.0-2 AEROBRAKING KIT WEIGHTS 








A-52 


9 




CONFIGURATION 

COMPONENT 

BASIC 

SHORT 60° FLARE 

LARGE NOSE 
FLARE 


2 PASS 

2 PASS 

30 PASS 

2 PASS 

HEAT SHIELD STRUCTURE 

TITANIUM 

TITANIUM 

TD-NICKEL CHROME 

RENE' 41 

HEAT SHIELD TPS 

ESA-3560 I I A 

ESA-3560 II A 

- 


SIDEWALL, PAYLOAD ADAPTER 
AND CLOSURE TPS 

MICROQUARTZ 
L-605 OUTER 
FOIL 

MICROQUARTZ 
L-605 OUTER 
FOIL 

MICROQUARTZ 
TITANIUM OUTER 
FOIL 

“ 

PAYLOAD ADAPTER 

ALUMINUM 

ALUMINUM 

ALUMINUM 

ALUMINUM 

ASTRIONICS 

FUEL AND 
REDUNDANCY 

FUEL AND 
REDUNDANCY 

FUEL AND 
REDUNDANCY 

FUEL AND 
REDUNDANCY 

RCS SYSTEM 

INCREASE 
THRUSTER 
SIZE AND 
FUEL 

FUEL 

FUEL 

FUEL 


FIGURE A-2.8.0-3 MATERIALS USED FOR AEROBRAKING KIT COMPONENTS 


D5-17142 



TOTAL TUG WEIGHT AT START AEROBRAKING (LBS. X 10 


D5-17142 


OO 

! 


14.5 


14 


13 


12 









— 








K- 

= SHOR 

T 60° 
C {NO 

FLAK 

FLAR 

E 









i 


-BASI 

E) 












r 






* 














+3 a ATM 

DENSITY V 

— ■ a n* ^ k.iAP>* r 

OSPH 

ARIAT 

* Ana— 

ERIC 

ION 









WIC: LAnU E H 

HAS NEG 
LOAD AN 

U» rLAKC 

ATIVE PAY- 
D NOT PLOTTED 



l 














| 





























r 

J-60 

FLARE 

1 










L ' 

y 

f 

45' FI 

.ARE 






- 




V 






. 

— 

— 





\ 

Nj 

V 




r SH 

ORT 6 

- 

0° FL 

1 

ARE . 

Lb ^* 1 






l 

\ 

>- 



1/-J 







F* 






L 

f 

-30® 1 

: LARE 


-NO FLARE 

1 1 












BASIC TUG AT START AEROBRA 

(POUND. 

ppnpin cimi IMCDTC R DM 

KING 

5) 








REACTION CONTROL 480 

ASTRIONICS INERTS 1,960 

PAYLOAD ADAPTER 200 

PROPELLANT 1,690 

TOTAL (WITHOUT 9,718 

AEROBRAKING) 

1 L_J 1 1 L 








t 


i 




■4WHI 


02 


10 


20 30 40 

NUMBER OF PASSES 


50 


60 


FIGURE A-2. 8.0-4: TOTAL TUG WEIGHT VS. NUMBER OF PASSES 

A- 53 


i 


D5-17142 




A-3.1 (Continued) 

furnished by MSFC. This revised model is applicable to the equatorial 
regions in which the aerobraking perigees are expected to occur. The 
total variations are one-half or less of those in the original model. 
The maximum rate of change is 3% per hour of flight time for short 
time periods (approximately 4 hours). For comparison purposes, the 
1962 Standard Atmosphere density values are approximately: (1) 85 KM - 

7.96 x 10-6 kg/m3, (2) 90 KM - 3.17 x 10-6 k g/m3, and (3) 95 KM - 
1.21 x 10-6 kg / m 3. 

The two preliminary targeting schemes discussed in Section 4.3 (basic 
report) were based on the invariancy of the apogee decay rate and 
used either burns at apogee or burns at atmospheric exit to achieve 
the desired decay rates. The former scheme used apogee burns to up 
the perigee (assuming a constant high density atmosphere) for correc- 
tion. The latter scheme used burns at exit to increase the apogee 
to the desired apogee. 

The final targeting scheme was based on making the atmospheric correc- 
tion burns at or near entry where better trajectory knowledge would be 
available. A series of trajectory were flown using atmospheric 
correction burns at entry to verify the feasibility of this scheme. 

Figure A-3.1. 0-2 shows the apogee decay rates for the 10 pass mission. 
The basic (no flare) was flown under four atmospheric conditions with 
the final targeting scheme utilized. The first atmospheric condition 
was the 1962 Standard Atmosphere to determine the nominal decay line 
(this nominal line is only dependent on mission duration and is 
identical to other 10-pass configurations). The effects of using a 
constant (+) or (-) density (initial atmosphere dispersion model] are 
shown and resulted in relatively large errors in the apogee subsequent 
to the first perigee pass. The apogee errors then rapidly converge to 
near-nominal for the remainder of the passes. Using the Varying At- 
mosphere of Figure A-3.1. 0-1, the initial apogee error is greatly 
reduced. Although the decay rate convergence is not as rapid because 
of the time variance in the atmosphere, equivalent convergence occurs 
by the fourth pass. 

Figure A-3.1. 0-3 shows the final apogee error comparison between the 
two atmospheric models. This figure (at the 10-pass absicssa) is an 
expanded plot of the final pass decay rate data shown in Figure 
A-3.1. 0-2 above. The Varying Atmosphere with the correction-at-entry 
technique resulted in an insignificant final apogee error of about 
3 n.m. (10 pass mission). The 30 pass varying atmosphere mission had 
a final apogee error of 13 n.m. The constant atmospheres also had 
small final apogee errors (8-29 n.m.). This range of final apogee 
errors is less than those shown in Figures 4.3.4. 1-7 and -8 (basic 
report) for the other targeting schemes. 


A- 54 


ATMOSPHERIC DENSITY (KG/M 3 ) 


D5-17142 




AEROBRAKING FLIGHT TIME AFTER FIRST PERIGEE (HOURS) 

FIGURE A-3. 1.0-1: ATMOSPHERIC DENSITY VARIATIONS FOR SPACE TUG AEROBRAKING STUDIES 








APOGEE ALT - 10° FT 


Cb-17142 







FINAL APOGEF. ALT ERROR ~llr FT 


r 


> 



NUMBER OF PASSES TO 270 NM 

FIGURE A-3. 1.0-3: SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - FINAL APOGEE 

ERROR VERSUS NUMBER OF PASSES TO 270 NM 


D5-17142 


D5-17142 


A-3.1 (Continued) 

Figure A-3.1. 0-4 compares the atmospheric correction delta velocities 
used to maintain the decay rates shown in Figure A-3.1. 0-2 and to 
attain the desired final apogee. The constant atmospheres (+ and -) 
required approximately 200-225 ft/sec during the mission and this value 
is relatively insensitive to number of passes. This insensitivity is 
contrasted to the large delta velocities for the short duration missions 
using the other two targeting schemes (Figures 4.3.4. 1-9 and -10 of 
basic report). 

The Varying Atmosphere required approximately 100 ft/sec delta velocity 
for the 10 pass mission and less than 200 ft/sec for the 30 pass 
mission. The Varying Atmosphere trajectory was flown by correcting each 
pass back to the nominal apogee decay line. Because of the limited re- 
sources available during the add-on activity, no optimization concerning 
numbers of correction burns was made. The Varying Atmosphere has cyclic 
characteristics tending to equalize the high and low peaks during a 
mission. Therefore, the Varying Atmosphere may require no atmospheric 
corrections during the early passes since the atmosphere may provide 
decay rates approximating the desired rates. It is also probable that 
bounds or limits of deviations from the nominal decay rate could be 
established and burns made only when these limits are exceeded. These 
optimization studies would be performed in follow-on activities. 

A-3.2 NAVIGATION ERROR CORRECTIONS 

Sections 4.6 and 5 (basic report) discussed the requirement for and the 
timing of navigation error correction burns. The data showed the de- 
sirability of performing these burns after the landmark tracker had 
made sufficient readings to converge the navigation uncertainties to a 
relatively small value. This occurred after 500-1300 seconds of land- 
mark tracker operation (just prior to entry). Using the navigational 
uncertainties (3 sigma values) during the total mission, one trajectory 
was flown to determine the optimum location of the navigational correc- 
tion burns. The results for a basic (no flare) 10-pass mission are 
shown in Figure A-3.2. 0-1. The minimum total mission navigation 
correction delta velocity occurs at a true anomoly of -30 degrees. 
However, making the navigation burns ac entry each pass insignificantly 
increased the delta velocity requirement and is also compatible with 
the atmospheric correction burn location discussed in Section A-3.1 
above. The possibility of combining these two independent burns to . 
reduce the total delta velocity requirement should be investigated. 

As discussed in Section 5 of the basic report, an in-depth analysis of 
this problem area is still required in a follow-on activity. However, 
this single test case has indicated a promising approach to achieve the 
desirable objectives of (1) maximizing perigee accuracies, (2) mini- 
mizing correction delta velocities, and (3) simplifying operational 
procedures . 


500 


> 

U1 

ID 


CO 

Q_ 

U_ 


> 

•o 


400 


300 


200 


100 


BASIC (NO FLARE} CONFIGURATION 
CORRECTION & V APPLIED AT ENTRY 


-CONST AN 


gk VARYING 
ATMOSPHERE 


T(-) DENSITY 



10 20 

NUMBER OF PASSES TO 270 NM 


VARYING 

ATMOSPHERE 


30 


FIGURE A-3. 1.0-4: SPACE TUG AEROBRAKING RETURN FROM SYNCHRONOUS ORBIT - INCREASE IN AV 

TO GO TO 270 NM FOR OFF NOMINAL ATMOSPHERE 


D5-17142 



COURSE CORRECTION AM (MEAN + 3 a ) - (FT/SEC) 


\ 


D5-17142 



TRUE ANOMALY OF COURSE CORRECTION (DEGREES) 


FIGURE A-3.2,0-7 : SENSITIVITY OF TOTAL MISSION NAVIGATION ERROR CORRECTION 

DELTA VELOCITY TO CORRECTION BURN TIME 

A-60 



A-3.3 


PAYLOAD CAPABILITIES 


The payload capabilities of the two pass configurations and 30 pass 
short 60° flare configuration were computed in the same manner as 
discussed in Section 5 (basic report). The effects of atmospheric 
dispersions and navigational errors were not specifically addressed. 
However, a delta velocity budget of 800 ft/sec was allotted for 
docking in synchronous orbit and for making correction burns. This is 
the same budget utilized in Section 5.1.2 (basic report). 

Figure A-3.3. 0-1 shows the fully fueled weight statements for the 
selected add-on activity configurations. The aerobraking kit weights 
are the same as shown in prior Section A-2.8 (Weights and Mass Pro- 
perties). The total mission delta velocities include the equivalent 
velocities furnished by the RCS. All three of the two pass configura- 
tions have low mass fractions due to their heavy thermal protection 
systems or large flare areas. 

The geosynchronous round trip payload capabilities of these configura- 
tions are shown in Figure A-3.3. 0-2. The two pass basic (no flare) 
and short 60° flare configurations have approximately equivalent 
capabilities. As discussed Section 5 (basic report), changing the mode 
from departing 100 n.m., circularizing at 270 n.m. and transferring 
back to 100 n.m. (270-100 n.m. mode) to a mode consisting of departing 
from and circularizing at 200 n.m. (200 n.m, mode) increases the pay- 
load capability approximately 1200 pounds. This increase is reflected 
in the two pass payloads shown in the figure. 

The two pass large nose flare has a negative payload capability, even 
when flown from 200 n.m. The inert weights of this configuration 
would have to be reduced by approximately 600 pounds to have a zero 
payload on the 200 n.m. mode. Because the gross weight of this con- 
figuration is slightly over 60,000 pounds, little if any additional 
assistance could be expected from the Shuttle. Therefore, it was con- 
cluded that this particular configuration has little applicability In 
the geosynchronous mission role. 

The 30 pass short 60° flare's payload capability is approximately 100 
pounds less than the 30 pass 30° flare's capability (Figures 5. 1.2. 6-1 
and -2, basic report). This is caused by two factors: (1) the equiva- 

lent mission delta velocity for the short 60° flare is T 50 ft/sec 
greater (The short 60° flare is not stable. RCS fuel is required for 
stability); and (2) the short 60° flare configuration required a com- 
pletely- insulated payload which negated most of the decrease in flare 
weight. 

The short 60° flare (flare slant height = 4,9') discussed above is the 
baseline it 2 configuration (prior Section A-2.1). This flare has the 
same equivalent W/C Q A as the 30° flare, however, it is statically un- 
stable. The short 60° flare #3 configuration (flare slant height = 
8.68') is neutrally stable. It has larger coefficients of drag and 
requires higher perigee altitudes (for same mission durations) as the 


COMPONENT 


CONFIGURATIONX 


NO. OF MISSION 
PASSES 


BASIC (NO FLARE) 


SHORT 60° FLARE § 2 


25 56,928 0.790 21,369 


57,334 0.785 

54,976 0.819 


60,183 0.748 20,915 


FIGURE A-3.3.0-1 FULLY FUELED AEROBRAKING TUG WEIGHT STATEMENTS 






































D5-17142 


A-3. 3 (Continued) 

#2 configuration. These configuration #3 characteristics could result 
in a larger payload capability. 

Figure A-3. 3.0-3 shows the fully fueled aerobraked Tug weight statement 
for the 30 pass short 60° flare configuration #3. The flare design 
concept is identical to that shown in Figure 4. 2. 2. 3-1 (basic report) 
for the 30°, 45°, and 60° flare configurations. The other aerobraking 
kit elements are identical to those discussed in the basic report. 

The # 3 configuration mass fraction for the 30 pass mission is 0.820 
while the § 2 configuration mass fraction for the 30 pass mission is 
0.819 (prior Figure A-3. 3.0-1). The aerobraking kit element weight re- 
duction was achieved by the decreased total sidewall insulation require- 
ment of the #3 configuration. Therefore, without considering the 
decreased delta velocity requirements of the neutrally stable #3 con- 
figuration, this configuration has a greater payload capability than 
does the #2 configuration. 

The 30 pass 30° flare configuration has a total aerobraking kit weight 
of 2238 pounds and a fully fueled mass fraction of 0.817 (Figure 
5. 1.2. 0-2, basic report). The 30° flare has the same mission delta 
velocity requirements as the short 60° flare #3 because of their 
equivalent stability characteristics. Therefore, the short 60° flare 
# 3 has a 30 pass round trip payload capability almost 200 pounds 
greater than the 30° flare. This neutrally stable short 60° flare Tug 
has the largest round trip payload potential of all the configurations 
examined in this study. 

Prior Figure A-2. 1,0-1 showed the preliminary flare configuration/net 
payloads sensitivity estimate. The preliminary estimate indicated 
that a moderately unstable configuration might maximize the payload 
capability. The short 60° flare payload analysis in this section in- 
dicates that a near-neutrally stable vehicle will maximize the payload. 
Based on these results, Figure A-3. 3.0-4 shows a revised estimate. The 
data points computed in this study are shown as small circles. Except 
for the 80° large nose flare (W/C 0 A ® 2), these data points represent 
30 pass data. The large nose flare's data point is for two passes. 

The trend lines shown on Figure A-3. 3. 0-4 represent estimates from the 
study results. Follow-on activity is required to verify/modify these 
trend lines. These follow-on activities include flow separation effects, 
flare designs, and dynamic stability impacts. Completing the data for 
this figure would have several advantages including: 

o Establishment of flared configuration scaling laws 

o Optimization of flare s;ze and stability characteristics 

o Establishment of configuration selection criteria based on 
allowable mission durations 


D5-17142 


COMPONENT 
TUG INERTS 
PROPELLANT 

TOTAL AEROBRAKING KIT 

AFT HEAT SHIELD 450 

FLARE 610 

SIDEWALL INSULATION 270 

PAYLOAD/ FLARE ADAPTER 390 

ASTRIONICS MODULE PENALTY 325 

FULLY FUELED TUG 
MASS FRACTION 


WEIGHT (LBS) 
7,828 
45,000 
2,045 


54,873 

0.820 


FIGURE A-3. 3.0-3 FULLY FUELED WEIGHT STATEMENT FOR 30 PASS 
SHORT 60° FLARE TUG (#3) 


A-65 


7 




* 


3=* 

I 

Cn 

cn 


CO 

a 


rD 

o 

Q_ 


co 

O 

<c 

co 


Q 

C 

o 

—I 

<c 

O- 

O- 

C£ 


O 

or 



FIGURE A-3. 3.0-4: FLARE CONFIGURATION/NET PAYLOAD SENSITIVITY ESTIMATE 


D5-17142 


D5-17142 


A-3.4 CONVENTIONAL/AEROBRAKING TUG PERFORMANCE COMPARISON 

The performance comparisons shown in Section 5.3 of the basic report 
were based on: (1) The conventional Tug's inert weight trends and 

mission delta velocity requirements developed in the prior Boeing Pre- 
Phase A Tug Study (Reference 1.1. 0.0-1) and (2) the aerobraked Tug's 
inert weight penalties and mission delta velocity requirements de- 
veloped in this study. Subsequent to the conclusion of the basic 
portions of this study, the MSFC Point Design Tug concept was de- 
veloped under another set of groundrules. This section compares the 
payload sensitivities of the initial MSFC Point Design Tug with the 
aerobraked 30° flare configuration (30 pass mission). The groundrules 
used for this comparison were those established for the MSFC Point 
Design Tug Studies and differ in some respects to those used in either 
Boeing's Pre-Phase A Tug Study or in the Aerobraking Study, These 
groundrules were used in this section to provide a common basis for 
analysis. The reader should not make direct comparisons of the aero- 
braking results in this section with any other in the Aerobraking 
Study without first rationalizing the differences in groundrules 
utilized. 

Figure A-3. 4.0-1 shows the round trip payload capabilities of the con- 
ventional and aerobraked Tugs. The initial MSFC Point Design goal of 
3000 pounds payload {A'= 0.895, Isp = 470 sec, total usable propellant 
weight = 55,552 lbs.) is shown at the top of the figure. If the main 
engine Isp were to be degraded to 460 seconds, approximately 500 pounds 
of payload capability would be sacrificed as evidenced by the Isp » 460 
line. As the conventional stage mass fraction is degraded (propellant 
loading constant), the payload capability is decreased until the capa- 
bility is reduced to zero when the mass fraction is approximately 0.855 
(Isp = 470) or 0.861 (Isp = 460). 

The conventional Tug used as a “Starting Point" for this aerobraking 
study has a mass fraction of 0.852 without the payload adapter (total 
usable propellant weight = 45,000 lbs). The 3000 pound round trip pay- 
load could be achieved by the aerobraked 30° flare configuration (30 
pass mission) with the current stage mass fraction (without the aero- 
braking kit) and with an uprated main engine having an Isp of 470 
seconds. Using the current engine (Isp = 460 seconds), the stage would 
require a mass fraction of 0.862 to attain the 3000 pound payload capa- 
bility. If the Aerobraked Tug stage were designed similar to the Point 
Design Tug and with a reasonable scaling factor to account for the 
differences in propellant loading (45,000 vs. 55,552 lbs.), the Aero- 
braked Tug might have a mass fraction of 0.875. This would provide a 
payload capability of approximately 4400 pounds (Isp = 470 sec,). 

The conventional Point Design Tug with 3000 pounds of payload, has a 
total gross weight in the Shuttle of 65,000 pounds, the maximum Shuttle 
capability at 28.5°/100 n.m. The Aerobraked Tug, with the same 3000 
pounds of payload, has approximately 7000 pounds less gross weight in 
the Shuttle. Therefore, the Aerobraked Tug is not as sensitive to 
possible degradations in Shuttle payload capability . 


MASS FRACTION (A') 



30° FLARE CONFIGURATION - 30 PASS MISSION 
100- 270-100 N.M. AEROBRAKING MODE 
MSFC POINT DESIGN GROUNDRULES USED FOR 
ALL NON-AEROBRAKING MANEUVERS ( A V’s, 
THROTTLED l sp , RCS l jp , ETC.) 

STANDARD ATMOSPHERE 
400 FT/SEC INCLUDED FOR AEROBRAKING 
NAVIGATION 

AEROBRAKING MODIFICATION NOT INCLUDED 
IN STAGE MASS FRACTION 
W Q = TUG & PAYLOAD WT g 100 N.M. OUTBOUNDl 

• PROPELLANT LOADING CONSTANT 


ROUND TRIP PAYLOAD (THOUSANDS OF POUNDS) 


FIGURE A-3. 4.0-1: ROUND TRIP PAYLOAD CAPABILITIES OF CONVENTIONAL AND AEROBRAKED TUGS 


D5-17142 



D5-17142 


i 


A-3,4 (Continued) 

Figure A-3.4.0-2 shows similar data for the payload placement mission. 
The Initial Point Design Tug has a placement capability in excess of 
8000 pounds. The Point Design Tug is constrained in the mission 
because of the 65,000 pound Shuttle limit. This limit forces the Point 
Design Tug to fly with approximately 5000 pounds of propellant off- 
loaded. The conventional Tug's payload sensitivities to possible de- 
gradations in both stage mass fraction and specific impulse are also 
shown. 

The smaller Aerobraked Tug does not require propellant off-loading to 
remain within the Shuttle's 65,000 pound constraint. Therefore, the 
Aerobraked Tug again has a significant advantage over the larger con- 
ventional Tug even though this particular mission takes the least 
advantage of the aerobraking potential. To place the same 8000 pound 
payload, the Aerobraked Tug would require a stage mass fraction (with- 
out aerobraking kit) of 0.8605 (Isp = 470 seconds) or 0.869 (Isp = 460 
seconds). The Aerobraked Tug's payload sensitivities to changes in 
stage mass fraction are less than for the conventional Tug because of 
the propellant off-loading factor. To achieve a placement capability 
of 10,000 pounds, the Aerobraked Tug would require technology similar 
to that of the Point Design Tug. 

Figure A-3.4.0-3 shows the retrieval mission data. The Initial Point 
Design Tug has a payload retrieval capability of 4160 pounds as shown 
in the figure. This capability is decreased as the design parameters 
are degraded. 

The aerobraking mode is most effective in the retrieval mission because 
the payload is carried only during the reduced propulsive requirement 
leg of the mission. Therefore, the Aer< braked Tug has a large retrie- 
val potential as shown in the figure. Even with the current state-of- 
art technology as used in Boeing's Tug ( A* = 0.852, Isp = 460 
seconds), the Aerobraked Tug significantly out performs the conventional 
Tug in this retrieval mission. 


A-69 


ot-v 

MASS FRACTION ( X ) 



0.90 
0.89 
0.88 
0.87 
0.86 
0.85 

0 1 2 3 4 5 6 7 8 9 10 11 12 

PAYLOAD PLACEMENT (THOUSANDS OF POUNDS) 

FIGURE A-3.4.0-2: PAYLOAD PLACEMENT CAPABILITIES OF CONVENTIONAL AND AEROBRAKED TUGS 



D5-17142 


MASS FRACTION ( A.' ) 


BASIS: 


CONVENTIONAL 
TUGS 


-MSFC POINT DESIGN - 
{ A' = 0.895, l sp = 470) 


30° FLARE CONFIGURATION - 30 PASS MISSION 
100-270-100 N.M. AEROBRAKINGMODE 
MSFC POINT DESIGN GROUNDRULES USED FOR 
ALL NON-AEROBRAKING MANEUVERS ( AY’s, 
THROTTLE & RCS I ’s, ETC.) 

STANDARD ATMOSPHERE 

400 FT/SEC INCLUDED FOR AEROBRAKING 

NAVIGATION 

AEROBRAKING MODIFICATIONS NGT INCLUDED 
STAGE MASS FRACTIONS . 

PROPELLANT LOADING CONSTANT S 


-EXCEEDS SHUTTLE 
CAPABILITY OF 65,000 LBS. 


AEROBRAKED 
TUGS- 


RETRIEVED PAYLOAD (THOUSANDS OF POUNDS) 


FIGURE A-3.4.0-3: PAYLOAD RETRIEVAL CAPABILITIES OF CONVENTIONAL AND AEROBRAKED TUGS 


D5-17142 


APPENDIX B 

BRIEF DESCRIPTION OF KALMAN FILTERING 


To understand the Kalman filter, it is helpful to examine the differences 
between two basic techniques for estimation. These are "least squares" 
and "maximum likelihood" methods. 

Least squares estimation involves fitting a curve to the available data 
in such a way that the sum square of the residuals to the data points is 
minimized. The least squares curve fit is simply the aver age of measure- 
ment data readings. The RMS error in the estimate is a//N, where N is 
the number of readings. In general, no least squares curve fit is possible 
until there are at least as many equations as there are unknowns. 

Maximum likelihood estimation introduces the new factor of "weighting" into 
the estimation process, in order to make allowances for the accuracy of the 
measuring instrument, as well as the variation in the quantity being 
measured. Thus, statistics concerning both the quantity being estimated 
and the errors in measurement are utilized. 

The RMS error for both the Least Squares and Maximum Likelihood concepts 
is shown in Figure B-l. 



FIGURE B-l. COMPARISON OF LEAST SQUARES AND 
MAXIMUM LIKELIHOOD ESTIMATION 


PRECEDING PAGE BLANK NOT FILM** 


B-l 



D5-17142 


A concept brought forth by Dr. R. E. Kalman in 1970, made feasible a prac- 
tical maximum likelihood estimation technique for use on digital computers 
in a real-time application. Kalman developed a recursive or iterative 
procedure that applies a weighting factor b to measured data y, to yield a 
new estimate ft. The equation involved is summarized as shown. 


b n = 


IT 0 ; 


M a\ 

2 

n-1 


n-1 
+ 0 


Xn - Xn-1 + bn (yn - MX n _-|) 


a n = 


" D n Vl 

where 

y 

2 

measured data 

A 

X 

- 

estimated value of measurement 

M 

= 

scale factor 

2 

CT X 

— 

variance or mean square deviation of estimate 

2 

a e 

- 

mean square error in measurement 

b 

- 

weighting factor 

2 

a 

- 

update mean square error in estimate 


A comparison of navigation using a Kalman filter with navigation using straight 
position fixes is shown in Figure B-2. In the case where straight position 
fixes are implemented, the inertial navigation position is adjusted to agree 
with the externally indicated position. The position update is as accurate 
as the external position fix measurement. However, the navigation errors 
grow unabated between position fixes. On the other hand, the Kalman filter 
position fix updates continually improve the estimate of position and velocity 
so that the RMS position error gradually converges to a steady state error 
much smaller than that achieved using straight position fixes only. 

In general, when properly mechanized, the Kalman filter will not introduce 
errors which are statistically larger than the errors which existed prior 
to making the estimate, This is true regardless of how poor the reference 
data are. It is definitely not true of least squares where one poor 


B-2 



D5-17142 



D5-17142 


measurement can greatly affect the mean average of the data. 

In summary, a Kalman filter makes possible real-time on-board navigation. 

The filter accepts various external data and makes corrections to the 
system "state". This state vector estimate improves with each succeeding 
measurement, provided the accuracy of the measurement is better than the 
previous estimate of the state. The Kalman technique applies in a practical 
sense principally to linear systems and requires considerable computing 
capability. 


D5-17142 


APPENDIX C 
ANALYSIS PROGRAMS 

1.0 TUG INTEGRATION PROGRAM 

1.1 GENERAL DESCRIPTION 

This program is a six dimensional orbital simulator that generates orbital 
trajectories by integrating a given set of initial conditions for as many 
orbits as required to satisfy one of three end conditions: Time of flight; 

flight-path-angle; or apogee altitude. Also, the program can calculate 
the required delta velocity to regulate perigee altitude and apply velocity 
increments at any point from apogee to the beginning of the atmosphere. The 
delta velocity is calculated as a function of semimajor axis, eccentricity, 
eccentric anomaly and velocity and is applied along the current velocity 
vector. 

1.2 SUBROUTINE DESCRIPTION 

a. Main - The function of MAIN is to control the overall operation of 
the program. It first reads the input data and initiates the program 
by callinq subroutine VAROP. 

b. VAROP - Subroutine VAROP monitors the steps necessary to perform an 
Encke ntegration of a near earth trajectory. 

c. EDITOR - EDITOR monitors the integration of the trajectory to determine 
when the desired terminal conditions have been met. Also, delta velocity 
corrections to adjust perigee are executed in this routine. 

d. MI SC - This subroutine is a collection of entry points, each of which 
performs some calculation related to conic trajectories. The calcu- 
lations that can be performed are as follows: 

o Conic radius vector as a function of time 
o Conic velocity vector as a function of time 
o Coast time between two points on a conic 

o Calculation of the elliptical or hyperbolic eccentric anomaly from 

the true anomaly 

o Solution of Kepler's equation 

o Calculation of the true anomaly of the desired altitude to perform 
a delta velocity correction 
o Calculation of altitude as a function of latitude 

o Calculation of argument of perigee and orbital period 

e. ORB I TP - This subroutine utilizes the position and velocity vectors 

to calculate the orbital parameters needed for use in other subroutines. 
The quantities calculated are angular momentum, semi-latus rectum. 


C-l 



D5-17142 


1.2 (Continued) 


specific energy* eccentricity, true anomaly, semi-major axis, and 
eccentric anomaly. 


f. RUN K UT - This subroutine uses a 6th order Runge-Kutta integration formula, 

with' step-size control, to calculate the disturbed velocity and position 
of the vehicle. 


9* FCALC - This subroutine calculates the disturbing acceleration acting 
on a vehicle with respect to an Earth centered coordinate system. In 
particular, this subroutine includes the second, third and fourth 
earth gravitational harmonics; calculation of the Encke acceleration; 
and calculation of drag accelerations. 

h. PARA62 - This routine is the 1962 standard atmosphere. 

i. INTRPT - Subroutine INTRPT is an iteration routine which calculates 

the current drag coefficient (C D ) as a function of current altitude and 
velocity. u 


\ 


j- 


BLKDATA - Block data is an initialization routine for constants used 
in the Tug program. 


2.0 ANS PROGRAM DESCRIPTION 

2.1 GENERAL DESCRIPTION 

2.1.1 Introduction 


The Autonomous Navigation Simulation Program (ANS) simulates autonomous 
navigation along a Keplerian orbit as performed by a Kalman filter using 
data from selected combinations of horizon sensor, star tracker, landmark 
tracking telescope, radar altimeter and range (laser) measurements. Au- 
tonomous navigation is the process of determining the orbit of a spacecraft 
from on-board the space vehicle. The program simulates the use of a Kalman 
filter to process a sequence of stellar referenced or earth referenced 
measurements. The latest measurement data and the current best estimate 
of the orbit are processed by a Kalman filter to obtain a differential 
correction to the estimated orbit, i.e., a more accurate estimate of six 
orbit ephemeris parameters - three components of spacecraft position and 
three components of spacecraft velocity. 

The program contains an environment simulation which generates true and 
estimated spacecraft position and velocity, landmark location and star 
sighting data. Given a set of initial conditions and an observation schedule, 
the program simulates the navigation process, producing a time history of 
RSS (root-sum-square) position error, RMS (root-mean-square) position error 
and other related quantities. The RSS position error, based on the dif- 
ference between true and estimated spacecraft position, is a measure of the 
accuracy actually achieved by the Kalman filter estimation process. The 
RMS error, based on the estimation error covariance matrix, is an ensemble 


C-2 



D5-17142 


2.1.1 (Continued) 

statistic derived from a linearized model of the navigation process and is 
a measure of the theoretical or predicted accuracy. RSS position error is 
the length of the vector difference between true and estimated spacecraft 
position. RMS position error is the square root of the diagonal terms of 
the covariance matrix. 

The program may be run in either of two modes: error analysis mode or full 

simulation mode. In the error analysis mode, which may be used if only 
navigation accuracy is of interest, the environment simulation generates 
true orbit parameters only. A time history of RMS position error and other 
covariance matrix-related quantities reflecting navigation accuracy is pro- 
duced. The calculation of the estimated orbit and its correction, and RSS 
position error, which are features of the full simulation mode, are omitted 
for error analysis mode. Available options include the addition of earth 
oblateness and air drag effects upon spacecraft motion. 

2.1.2 Navigation Systems Included 

The following categories of navigation systems can be simulated (the 
observables measured by each system are indicated in brackets): 

a. For known or unknown landmarks 

o Two star trackers and landmark tracking telescope [measures 

pi tch and roll angle orientation of landmark line-of-sight in a 
known stellar orientation frame], 

b. For known landmarks only 

o Range (laser) measuring device [range, which is the length of the 
line-of-sight vector from the spacecraft to the landmark], 
o Two star trackers, landmark tracking telescope, and range (laser) 
measuring device [pitch, roll, range], 
o Single star tracker and landmark tracking telescope [star- 
landmark angle]. 

o Horizon sensor and landmark tracking telescope [landmark-vertical 
angle]. 

c. Other systems 

o Two star trackers and horizon sensor [star-vertical angle], 
o Radar altimeter [altitude, or the radial distance from space- 
craft to earth's center]. 

For these modes, the Kalman filter estimates only the six orbit ephemeris 
parameters. 

The categories of navigation systems and their combinations simulated in the 
ANS program are depicted in Figure C-l. 


C-3 



05-17142 


Name of Mode 

Navigation Sensors 

Star-Landmark: 
Submode 1 

2 Star Trackers, 
Landmark Telescope 

Star-Landmark: 
Submode 2 

Range (Laser) 
Measuring Device 

Star-Landmark : 
Submode 3 

2 Star Trackers, 
Landmark Telescope, 
Range (Laser) 
Measuring Device 

Star-Landmark: 
Submode 4 

Single Star Tracker 
Landmark Telescope 

Landmark- 
Verti cal 

Horizon Sensor, 
Landmark Telescope 

Star-Vertical 

2 Star Trackers, 
Horizon Sensor 

Altimeter 

Radar Altimeter 

Tracking 

MSFN and/or DSNT 
Tracking Stations 


No. of 


Landmarks 

Observable(s) 

States 

Known , 

Pitch, 

9 

Unknown 

Roll 


Known 

Range 

9 

Known 

Pitch , 

Roll, 

Range 

9 

Known 

Landmark - 
Single Star 
Angle 

9 

Known 

Landmark- 
Vertical Angle 

10 

N/A 

Star-Vertical 

Angle 

6 

N/A 

Altitude 

6 

N/A 

Range 

Range Rate 

Azimuth 

Elevation 

6 


Figure C-l 

AUTONOMOUS NAVIGATION MODES 



D5-17142 


2.1.3 Input Description 

Program input consists of the initial conditions {initial time, true and 
estimated spacecraft position and velocity, and covariance matrix), 
information pertaining to the options to be included in the simulation, and 
the observation schedule. Specified for each observation are the time and 
type of observation, and the instrument pointing accuracy. For landmark 
tracking systems, landmark acquisition angle (pitch and roll) and the land- 
mark survey accuracy, are also specified. For two systems, star acquisition 
angle(s) are additionally specified (star azimuth for the two star trackers - 
horizon sensor system; star azimuth and star elevation for the single star 
tracker-landmark tracking telescope system). For landmarks, a sequence of 
equally-spaced times of observations on the landmark, instead of a single 
observation time, is specified. For ground-based tracking systems, a sequence 
of equally-spaced times of observations by the tracker is specified. 

2.1.4 Kalman Filter Operation 

At each observation time, for each observable appropriate to the navigation 
system observation type, the Kalman filter equation sequence is used to 
generate updates to the estimated orbit (simulation mode only) and the 
covariance matrix. 

The sequence of equations providing the weighting or gain matrix W(K), 
the updated estimated orbit a(K+ 1) (simulation mode only), and the updated 
covariance matrix P.(K+) at the Kth observation time are; 

o W(K) = P(K-) H(K) CH T (K)P(K-)+Q]’ 1 

o X(K+) = X(K-)+W(K) [A(K)-A(K)] 

o P(K+) = P(K-)-W(K)H T (K)P(K-) 

~ A 

A(K) and A(K) are the actual and estimated observables, H(K) is the observable- 
related gradient vector, and Q is the variance of instrument noise errors. 

The above update sequence is executed for each of the observables appropriate 
to the navigation system observation type (the observables, which are anqles, 
or vector lenqths, were listed above for the various navigation systems). 
a 

X.(K-) t s the estimated orbit before update, for an observable under con- 
sideration. If this observable is the first or only observable for the 
navigation system, x(K-) is the estimated orbit after propagation from the 
preceding (K-l) observation time. A If the observable is not the first in a 
set of more than one observable, X.(K-) is the estimated orFit- after update 
for the preceding observable in the set. 

Similarly, £_( K- ) is the covariance matrix before update. If the observable 
under consideration is the first or only observable for the navigation system 
P_(K- ) is the covariance matrix after map up from the preceding (K-l) ob- 
servation time. A state transition matrix (Jacobian matrix, computed in 


C-5 


D5-17142 


2.1 .4 (Continued) 

closed form) along a Keplerian orbit is used to map up the covariance matrix. 
A program option is available which provides a machine noise modification to 
the Kalman filter to prevent filter divergence caused by modeling and com- 
puter round-off errors. The form of this filter modification is an additive 
term to the covariance matrix. 

For landmark tracking systems a sequence of observations is made on the 
landmark (the sequence timing is specified in the observation schedule 
input data). The first sighting is used by the environment to compute 
landmark position coordinates from the acquisition time and acquisition 
pointing angles which form part of the input data. Subsequent sightinqs to 
this landmark are then treated as navigation observations, i.e., data is 
processed by the Kalman filter into updates of orbit and landmark parameters 
and of the corresponding covariance matrix. 

Orbit propagation and covariance matrix map up occur between observations. 
Filter modification occurs only when the landmark is spotted. 

2.1.5 Output Description 

Three types of output are available: 

o Printout (one of three types, which include standard items plus 
varying amounts of additional items, may be selected), 
o Cal-Comp Plots (optional), 

o Punched Covariance Matrix (optional) 


D5-17142 


APPENDIX D 

NAVIGATION COMPONENTS 


1 .0 NAVIGATION COMPONENT DESCRIPTION 

This section presents a description of all the hardware elements that were 
considered as candidates for the Space Tug aerobraking navigation. 

1.1 IMU 

The platform has four gimbals with appropriate synchros, resolvers and torque 
motors for each gimbal. The angular sequence starting with the inner gimbal 
is pitch, roll and yaw with the fourth gimbal providing redundant roll. The 
stable element (inner member) contains two-degree-of -freedom gyros with their 
spin reference axes directed along the pitch and yaw gimbal axes. One gyro 
controls the roll and yaw platform gimbals while the other controls the 
pitch gimbal axis. The platform baseline is typified by a Kearfott KT-70. 

1.2 STAR TRACKER 

The star tracker is a strapped down optical sensor using electronic gimbalinq 
to determine star positions within the eight degree diameter field-of- 
view (FOV). The acquisition mode results in a scan of the entire FOV after 
which the brightness object is selected. The tracker then enters a tracking 
mode in which the selected object is scanned over a very small FOV, on the 
order of 16 arc minutes. The position of the object is measured in two axes 
with respect to the bores iqht of the tracker. 

The star tracker baseline is typified by the ITT Dual Mode star tracker. 

1.3 HORIZON SENSOR 

Lockheed Missiles and Space Company (LMSC) has developed and qualified a 
high accuracy, horizon sensor for use at synchronous altitude. This sensor 
has two scanning mirrors, one in the pitch axis and one in the roll axis, 
and will operate over large displacements in either axis. The sensor is a 
4 pound, 5 watt earth sensor which will perform for over 5 years in orbit 
with an accuracy of 0.05 degrees (3 sigma). 

1.4 AUTOMATIC LANDMARK TRACKER 

The automatic landmark tracker or automatic earth feature sensor (AEF5) is 
similar to the star tracker in that it consists of a sensor unit and an 
electronics unit. The entire sensor unit consisting of a vidicon and optics 
is gimballed in pitch and roll. Gimballing and optical magnification is 
required to obtain the desired accuracy. Gimballing is required for initial 
image motion compensation for acquisition and then final tracking; optical 
magnification of 2 or 3 power is required to overcome the TV line limitation 
for resolution. The gimbal angle readouts requirements are rouqhly 15 arc 
seconds requiring a simple inductosyn resolver and a 17 bit encoder. 


D-l 


D5-17142 


(Continued) 

Sequence of Operation 

o Known Target Mode - The field-of-view is directed to the estimated 
location and tracked with estimated image motion rates. The 
shutter is opened, exposure made, and the video readout examined 
to set the circuits and logic levels. A second exposure is made 
and the video readout is processed to compute the x value. A 
third exposure is made and read out with the scanning lines 
orthogonal to the previous readout. The video is processed to 
compute y. These values define the roll and pitch pointinq errors. 
The gimbals are corrected by the distance of x + y from boresiqht. 

A second sight is taken (two exposures) and a new x, y, value 
obtained. A successsion of sights, taking two to four seconds 
per cycle will allow both pointing accuracy to one resolution 
element, and image motion compensation (IMC) to one resolution 
element per cycle time. 

The only information which must be stored to identify a known 
landmark is the geodetic position. Target separation, size, 
contrast and other pertinent characteristics are accounted for 
in the selection process. 

During all coasting phase navigation, an extrapolation of position 
and velocity by numerical integration of the equations of motion 
is required. The integration scheme implemented dictates the 
integration increment required. Therefore, the number and 
location of landmarks must be chosen to maintain the required 
integration accuracy. 

o Unknown Target Mode - The unknown mode of operation is the same 
as the known mode, except that the sensor is pointed maximum 
forward along track and the estimated IMC applied if necessary. 

The video is examined to determine if a trackable tarqet is 
within the field-of-view. When a trackable target appears, the 
operation is identical to the known mode. 



D5-17142 


APPENDIX E 

RECOMMENDED AEROBRAKING FOLLOW-ON ACTIVITIES 


E-l FOLLOW-ON TUG AEROBRAKING PROGRAM ACTIVITIES 

This initial program was conducted to determine the feasibility of aero- 
braking as applied to the Space Tug and to identify the technical considera- 
tions associated with the aerobraking mode. From this study, the significant 
potential advantages of Space Tug aerobraking entry were indicated. This 
study, however, did not examine the concept in sufficient depth to assess 
its full potential. Further studies are required to identify the missions 
and mission capabilities, the operational modes. Tug aerobraking economics 
and their impact on overall space program costs, Tug interfaces with other 
space transportation systems such as payloads, etc., and the impact of a 
Space Tug modified for aerobraking on the Space Shuttle. Each of these 
studies are discussed below. 

The additional Tug aerobraking activities should be done in sufficient 
depth to bring the level of knowledge on Space Tug aerobraking to that 
obtained by the ongoing conventional Space Tug activities. Further, when 
these levels are comparable, comparisons should be made to fully assess 
the advantages and disadvantages of applying the aerobraking system as a 
major consideration for future Tug studies. 

E-l . 1 MISSIONS 

This study primarily investigated round trip mission modes. Placement 
and retrieval modes were not considered in the technical trades. As 
the initial missions in the space program will most likely be placement 
only missions, it is desirable to determine the trajectory modes for 
payload placement and return of the^ empty Tug and their effect on the 
Shuttle and its operational modes. Further, this study investigated only 
synchronous orbit missions. Aerobraking should be investigated for other 
missions. For example, significant payload advantages may be obtained 
when aerobraking is applied to other high altitude earth orbital missions. 
Also, for interplanetary type missions, where recovery of the Tug is 
desirable, aerobraking may offer the potential for reducing the Space Tug 
propulsive requirements for the return mode and therefore increasing the 
interplanetary payload capability. 

E-l. 2 OPERATIONAL MODES 

The primary operational mode examined in this study was an off optimum . 
one wherein the Space Tug returns from synchronous orbit to a low 
(approximately 50 n.m.) perigee by 270 n.m. apogee orbit. The Tug then 
circularizes at 270 n.m. and after proper phasing between it and the 
Shuttle, in the lower 100 n.m. orbit, the Tug accomplishes propulsive 
burns to transfer from 270 n.m. to 100 n.m. The Tug circularizes at 
the 100 n.m. orbit. Then the Tug provides the necessary rendezvous and 
docking maneuvers. This mode of operation severely penalizes the Tug 
payload capability. The Space Shuttle has the capability to perform 
many of these operations including ascent to somewhat higher orbits with 


E-l 



D5- 17142 


E-1.2 (Continued) 

little penalty to the Shuttle capability. Therefore, it is highly desirable 
that trades be accomplished on a mission by mission basis to identify 

(1) which of the interfacing vehicles should perform the various operations, 

(2) the Tug deployment and retrieval altitudes, and (3) the rendezvous and 
docking operations. The navigation timelines for performing all of these 
operations should be defined. 

E-1.3 ECONOMICS 

The economics of the Space Tug aerobraking has not been defined. R&D and 
operational costs for the aerobraking kits should be developed to define 
the impact on the overall Space Shuttle/Space Tug R&D and operational bud- 
gets. It is believed that the R&D costs for the aerobraking components 
will be small. Since the components may be designed for refurbishment 
and reuse during the operational phases, the overall costs of applying the 
aerobraking kit options can be minimal. An important economic considera- 
tion is the advantage of applying aerobraking to reduce the required number 
of Shuttle flights and to negate the requirements for orbital fuel trans- 
fer and assembly operations. This economic advantage is so significant 
that it should be defined as soon as possible in order to allow the 
\ aerobraking concept to have the required effect on the Space Shuttle/Space 

Tug design and operational planning. 

E-1.4 INTERFACE ANALYSES 

The Space Tug is designed to interface with all of the other space trans- 
portation systems including the Space Shuttle, satellites, interplanetary 
payloads, Space Stations, Nuclear Shuttles, etc. These interfaces will 
impose constraints on the configuration of the Space Tug and will impose 
further constraints on the operations that can be performed during aero- 
braking and on the environments which are permissible. An analysis of 
these interfaces should be undertaken to determine their impact on the 
Space Tug aerobraking concept. 

E-1.5 SPACE SHUTTLE/AEROBRAKING SPACE TUG INTERACTIONS 

The Shuttle and Tug must operate as a team for the majority of the space 
program missions. As a result, these two elements have a 'significant 
impact upon each other. The Shuttle has certain acceleration, thermal, 
handling, timeline, operational, and other constraints which will impact 
the aerobraking Tug design. Similarly, the aerobraking Tug will impose 
certain constraints upon the Shuttle, for example, the use of the aero- 
braking kit will impact hard points for stowage, deployment and retrieval 
mechanisms, etc. Investigation of these areas should be undertaken to 
minimize the aerobraking Tug impact on the Shuttle. 




E-2 



D5-1714? 


E-2 SUPPORTING RESEARCH AND TECHNOLOGY/ADVANCED RESEARCH AND 

TECHNOLOGY IMPLICATIONS 

This study was specifically conducted to define the feasibility and practi- 
cality of the aerobraking return trajectory mode for the Space Tug. The 
baseline Tug configuration utilized for this activity was a scale-up of 
the Tug configuration developed by The Boeing Company for NASA under a 
prior contract (prior Reference 1.1. 0.0-1). Groundrules for this pre- 
vious study specified utilization of current and projected state-of-the-art 
technology for the Space Tug configurations. The previous study and this 
current study, therefore, did not consider advanced technology in the design 
of structures, components, engine systems, astrionics, etc. As such, this 
study has presented the payload capabilities of the aerobraking mode con- 
servatively. Increased payload capability for the aerobraking mode should 
therefore result from designs applying advanced technologies. Further, 
the limited scope of this study precluded in-depth investigation of the 
anticipated aerobraking environments, their impact and the optimal material 
and design solutions for overcoming the environments. Therefore, attractive 
advanced technology areas requiring further attention were defined and are 
outlined below. 

E-2.1 AERONAUTICS 

Aerodynamic Properties of Aerobraking Tug Options - The design of an optimum 
aerodynamic configuration cannot be adequately defined without a wind tunnel 
test program to ascertain the aerodynamic properties of potential configura- 
tion options. These options include nose bluntness, cylindrical length, 
flare angle, and angle of attack versus drag. The effects t>f real gas 
aerodynamics on drag is required as the flare is embedded in the flow field. 
The impact of angle of attack on lift coefficient, pitching moment coeffi- 
cient and heat transfer rate distribution will provide substantiation for 
computation technique to be used in future studies. 

Flow Field and Flow Separation - The effects of flow separation were not 
investigated within the limited scope of this study. Disassociated flow 
field effects will change the shock position and the boundary layer 
characteristics and can significantly alter the aerodynamic characteristics. 
Further, boundary layer height and separation effects were assumed to have 
a negligible influence on the aerodynamics characteristics. This needs 
verification by further analysis. 

The effects of non-diffuse reflection and imperfect thermal accommodation 
in the free molecular flow regime have not been investigated and need 
analysis. 

Rarefield flow field effects on the flared configurations depend on local 
flew field properties approaching the flare. These effects were approxi- 
mated but need further analytical verification or experimental verification. 


E-3 


D5-17142 


E-2 (Continued) 

Aerodynamic Characteristics of Alternate Drag and Stability Configurations - 

The aerobraking configurations investigated in this study were selected to 
provide a range of resulting data to establish feasibility and trends. They 
do not necessarily represent the more desirable systems. In future 
activities, consideration should be given. to the aerodynamic characteristics 
of alternative configurations for drag and stability such as (1) split 
flares, (2) a forward drag brake, (3) different entry nose shapes, 

(4) forward facing engine exhausting parallel to the flow, (5) "lift" 
flare configurations, and (6) other options such as forward and entry 
(payload first), sideways entry, tumbling entry and use of extendable 
micrometeoroid shielding to provide drag brakes. The effect of flare 
semi vertex angle and flare length should be parameterized. 

Dynamic Stability Effects - No dynamic stability effects were included in 
the aerodynamics study. The dynamic instability may severely affect the 
reaction control system and therefore should be a part of follow-on studies. 

Angle of Attack - The effects of angle of attack and lift-to-drag ratios 
were not investigated as part of this study. The use of applied angle of 
attack as a method of attaining the desired perigee altitude should be 
studied. The impact of angle of attack, in turn, upon the thermal environ- 
ments, thermal protection system requirements and control (RCS fuel) 
requirements should be investigated. The maximum correction capability 
(planar and lateral) should be determined. 

Flow Field Effects on Aerothermal Environments - Flow conditions behind 
the Space Tug payload section (basic no-flare configuration) and behind 
the aerodynamic flare (flare configurations) should be investigated in 
order to better define aerothermal environments in the area. This activity 
could be augmented by a series of wind tunnel tests. 

Aerothermal Environmental Effects on Proturberances and Surface Discontinui- 
ties - The effects of proturberances and surface discontinuities' should be 
investi gated. Gaps occurring at the dome closure point and between adjacent 
hot structures must be evaluated for thermal design adequacy. 

E-2. 2 MATERIALS AND STRUCTURES 

High Temperature Materials - This study generally assumed TD-nickel -chrome, 
with the maximum temperature capability of 2000°F, as an upper radiative 
temperature limit material. This restriction limited the configuration options, 
and operational and mission modes significantly. The use of higher tempera- 
ture materials in the 2000° F to 3500°F range, such as the various columbiums 
and Fansteels (tantalum alloys) may offer reduced trip times with equivalent 
payloads. These materials, however, may have physical properties, lack of 
availability, or cost which will restrict their use. A survey of potential 
materials, their characteristics and their potential availability and cos.ts 
should be made and recommendations made as to which of these materials would 
enhance aerobraking and should therefore be pursued. 


E-4 


E-2.2 


(Continued) 


Alternative Thermal Protection Systems - The materials used for aerobraking 
kit components concentrated on the use of radiative type materials for the 
aft heat shield and the aerodynamic flare. Limited analyses of ablatives 
for the heat shield was undertaken in the add-on study effort. An investi- 
gation should be conducted considering alternative radiative materials 
and/or alternative thermal protection techniques which could, for example, 
utilize active cooling, transpiration cooling, exhaust gas cooling (with 
Tug engine operating in idle mode) and/or combinations of these systems. 

Integrated Thermal /Structures/Micrometeoroid Systems - The use of an 
integrated structure offers a means of reducing the lug inert weights. 

The weight of the Tug sidewall structures may be substantially reduced by 
use of an integrated thermal /structure/micrometeoroid system. Analytical 
and experimental methods should be developed to verify the integrated 
design approach, e. g., dimensional analysis and modeling techniques can 
be developed for studying the effects of meteoroid impact. 

Alternative Design Concepts - This s‘;uly investigated point designs for the 
aft heat shield and the aerodynamic flare. Alternative design concepts 
should be investigated to determine ways of reducing aerobraking kit 
weights, simplifying the aerobraking kit actuation systems, improving 
reliability and reducing risk. Alternative design options could include 
forward drag flares, inflatable flares, split flares, "lift" flares, 
deployable micrometeoroid shielding for drag brakes, etc. 

Materials Environmental Capabilities - The environmental design data and 
design criteria should be developed for potential missions. The capability 
of non-metallic materials that will be exposed to vacuum and radiation 
environments should be determined. For example, the application of advanced 
composites such as graphite-epoxy and boron-epoxy to integrated system 
design should be considered as a method of reducing Tug inert weight. 

Materials Processing - Potential material and process problems associated 
with fabrication and/or refurbishment of aerobraking and thermal shielding 
components should be investigated. 

E-2.3 ELECTRONICS AND CONTROL 

Reaction Control System Operating Concepts and Modes - The reaction control 
system design and operating mode used for this aerobraking study were 
those developed for the Conventional Space Tug and are not considered as 
optimum for the aerobraking mode. The desirable sizing of the thrusters, 
the minimum and maximum operating times, accuracy, stabilization periods, 
the limit cycle requirements, and the specific type of reaction control 
system should be investigated in more depth. The impact of these require- 
ments on RCS fuel consumption should be assessed. The use of gaseous 
hydrogen and oxygen which will result from boiling of the main tank 
residuals should also be considered for augmentation of the reaction con- 
trol system. 



05-17142 


E-2,3 (Continued) 

Reaction Control System Requirements for Lift Trajectory Modes - The 
control system sizing and propellant requirements for configurations 
utilizing lift for atmospheric trajectory corrections should be evalua- 
ted. 

Alternative Reaction Control Systems - The previous study selected a 
gaseous LOX/LH? RCS based upon the application of the then envisioned 
Space Shuttle technology. The current Space Shuttle RCS is based on the 
use of storables. The applicability of storables to the Space Tug both 
with and without aerobraking requires assessment. Criteria for evaluation 
should include performance, weight, mission life required, variation of 
impulse with pulse width, and funding requirements. 

Stability Margin Effects - During the aerobraking missions, it may be 
desirable to employ lift to attain target perigees, to accomplish minor 
plane changes and/or to adjust perigee to compensate for the effects of 
atmospheric anomalies. The required controlling moments relative to 
the aerodynamic moments for the proper angle of attack for the desired 
lift-to-drag ratio, will require further study. Variable methods of 
achieving stability and the stability margins should be developed. 

Update Capability Versus Control Requirements - An assessment should be 
made of the optimum sensors and update times to minimize RCS fuel re- 
quirements for error correction while at the same time minimizing perigee 
uncertainty. By proper selection of the time in which the guidance 
avionics are updated, the control fuel requirements may be reduced by a 
factor of 5 or more. Update timing which minimizes control requirements 
tends, however, to increase measurement uncertainty. 

This recommended study effort would be an expansion of the effort included 
in the navigation analysis. The effort would perform detailed trades to 
determine the updating capability of onboard navigation components versus 
the required control penalties. This would include attitude control dead- 
bands and requirements during sensor observations js well as the accuracy 
versus RCS penalties to perform navigation update burns for various times 
during the aerobraking orbit prior to perigee. 

Astrionic System Configuration Analysis - The analysis performed for the 
present study was an updating of the initial Space Tug astrionic system 
design using the latest Shuttle concepts and components to perform only 
the synchronous mission instead of the broad spectrum originally studied. 
Present Shuttle emphasis appears to be in the area of low cost without 
weight being a pacing item. For the Space Tug, weight and cost would be 
pacing items. Therefore, additional effort should be expended to integrate 
Shuttle astrionic system concepts and components into the Tug while main- 
taining minimum weight where possible. Although this is not an aerobraking 
analysis per se, it is important to provide a well-defined baseline system 
configuration as a basis for future aerobraking study effort. 



D5-17142 


E-2.3 (Continued) 

Redundancy Analysis - The redundancy effort to date has defined the typical 
redundancy weight deltas to be expected for aerobraking mission. Additional 
effort should be expended to provide a more detailed redundancy analysis 
using the updated astrionic system components, and to look individually at 
each component to determine methods of reliability and coverage enhance- 
ment. In addition, the risks associated with redundancy management for 
long duration missions should be addressed. 

Navigation Timeline Analysis - The navigation analysis to date selected a 
typical navigation update timeline which provided satisfactory update 
accuracy. However, more analysis is required to perform a detailed opera- 
tion analysis, using attitude constraints, sensor acquisition and reacquisition 
constraints, burn perturbations, length and frequency of updates, etc., to 
insure that the autonomous navigation during aerobraking is indeed feasible. 

Advanced Sensors Systems - The ability to accurately predict the location 
of the Tug during the aerobraking operations is dependent upon the 
accuracy of the sensors systems. The present sensors systems have not been 
subjected to a detailed analysis as to their performance capabilities 
and/or reliabilities. The use of advanced sensors which represent Shuttle 
era technology should be likewise examined in detail for performance and 
reliability capabilities. The reconmended study would look at existing and 
potential autonomous navigation sensors. One aspect of the study would be 
improvement of autonomous space navigation by using new technology sensors 
and/or concepts. A second part of the study would include tndepth analysis 
of navigation sensor hardware to determine means of enhancinq reliability 
and to determine operating modes and limitations of the hardware. 

Redundancy Implementation - The successful completion of the Tug aerobraking 
mission depends on an operational astrionics system which in turn depends 
on "coverage" (failure identification) to "keyin' redundancy components 
at the proper time. Current technology depends heavily on BITE (Built- 
In-Test-Equipment), off-line dynamic response testing, voting or comparison 
or reasonableness testing to identify failed components. These methods 
have had limited success with electromechanical sensors and hence the 
development of a new component evaluation technology is desirable. 

The basic evaluation limitations can be overcome by using random or pseudo- 
random noise input and correlation techniques. The advantages of using 
correlation identification techniques are (1) the system may be checked 
out "on-line", (2) test signals can be kept small and will not interfere 
with normal operation, (3) results can be obtained in the presence of 
random noise and parameter drifts, and (4) the technique can be easily 
applied to existing hardware as correlation is inherently a simple digital 
process. 

The objective of such a new technology study would be to define the 
suitability and application of "on-line" system evaluation by digital 
methods . 


E-7 



D5-17142 


E-2.3 (Continued) 

Navigation Sensor Integration - The navigation analysis performed in this 
study utilized an optimal filter (Kalman) implemented on a general purpose 
floating point computer (IBM 360/75). In a space application, the sensor 
integration routines must be programmed in a limited memory machine in 
fixed point arithmetic. A new technology task is recommended to define 
a suboptimal filter routine in 16 bit fixed point arithmetic to provide 
integration of the landmark, horizon, and star sensors and the IMU. 

Navigation and Guidance Analysis - The study effort to date has evaluated 
navigation uncertainties and preliminary evaluations of burn corrections 
and associated updates. To enhance the above effort, a study should be 
made considering the navigation system configuration, navigation accuracy, 
attitude pointing accuracy, targeting schemes and predictive problems 
associated with orbit disturbances that result from uncertainties in at- 
mospheric density, navigation sensor, and vehicle dynamics. The study 
would involve development of a guidance law to predict future orbital 
variations based on past inputs (accelerations, state vectors), compute 
new orbit perigees to achieve mission objectives, and control vehicle 
forces during atmospheric braking to achieve the desired end conditions. 

The guidance law would then be tested using available simulations to verify 
its operation under a variety of disturbances from nominal performance. 

The midcourse velocity corrections would be defined and the subsystem 
weight penalties determined. 

Radiation Analysis - The radiation analysis for this study was a "quick 
look" to identify any significant astrionic system impacts because of 
repeated passes through the Van Allen radiation belt. Additional study 
effort would include a more accurate determination of the elliptical orbit 
profiles using the Burrell Orbital Flux and Spectra (BOFES) code. The 
impacts of both single aerobraking missions and repeated aerobraking mis- 
sions for a ground based Tug would be evaluated. Sensitivities of various 
components for various shielding and radiation doses would be addressed. 
Also, the effort could be expanded to include the impacts of artificial 
radiation environments if deemed desirable. 

Astrionic System New Technology Component Analysis - As mentioned pre- 
viously, weight is the pacing item for the Space Tug. The astrionic 
system weight can be reduced if new technology components (in lieu of 
Shuttle components) are considered. However, development costs and de- 
velopment risks would be associated with the new technology components. 

A study effort is required to evaluate the relative merits of using new 
technology components in the astrionic system of the Tug. This would 
include a survey of potential new technology components (such as LSI 
computers) and the relative development progress of each. 


E-8 


05-17142 


APPENDIX F 

TWO PASS LIGHT WEIGHT LARGE FLARE 
F-1.0 LIGHT WEIGHT LARGE FLARE CONCEPT 

The large nose flare concept failed to have any payload capability due 
to the flare weight. One of the attractive alternative concepts in- 
vestigated to decrease the flare weight employs an inflatable torus to 
extend the flare. This concept (see Figure F-1.0. 0-1) integrates the 
radiative heat shield design (heat shield functions as described in 
Section 4. 2. 2.1 of basic report) with the forward flare. The flare 
attaches to the heat shield at the dome/cylindrical section interface 
as shown in the figure. The flare and torus would be fabricated of 
"AIRMAT," a "fabric" woven on the U. S. Air Force Loom operated by 
Goodyear Aerospace Corp. in Akron, Ohio. The Airmat "cloth" would be 
woven from Haynes 188 (Cobalt alloy) .0005 inch diameter wire and 
operate at a maximum temperature of 2000°F. This Airmat material has 
a porosity of 1% which means that 99% of the flare area would serve as 
a decelerator. When extended by the torus, the Airmat would result in 
a decelerator area of 50 feet in diameter. At the 50 foot diameter, 
the Airmat wraps around the torus. To reduce the torus temperatures 
from 2000 to 1500°F at the tip of the flare, a 1/16 inch thick coating 
\ of silicone rubber was used as noted. The torus is made from the same 

Airmat material and is sealed by a spray coating of silicone rubber. 
The torus is pressurized by helium from helium bottles housed under 
the L0X TANKS. The inflated torus shape is held rigid by 91 Rene' 41 
cables spaced 20 inches apart. The cables are extended/retracted by 
electrically driven cable drums. Two methods were designed for pack- 
aging the retracted flare. If the Tug's diameter is fixed at 14 feet, 
then the flare would retract alongside the Tug sidewall as shown in 
the lower part of the figure. This approach would necessitate that 
the RCS system be moved aft into the astrionics module region. The 
other method would take advantage of the L0X tanks smaller size and 
would taper the Tug sidewalls to the smaller diameter. The flare 
could then retract into the recessed region as shown in the figure. 
This concept allows the RCS system to stay at the desired station. 

A protective door can be placed on the Tug to restrain the retracted 
flare. 

The torus as shown consists of one large bag. A non-col lapsable 
flexible hose would be used to supply/remove the helium from the 
torus. The configuration was designed for 20 missions, with creep- 
to-rupture at 2000* F as the design criteria. The weight of the 
system is 2703 pounds as compared to the 6580 pounds of the large 
nose flare. The two pass weight statement of this configuration is 
shown in Figure F-1.0. 0-2. 


F-l 


D5-17142 


COMPONENT 

WEIGHT (POUNDS) 

AFT HEAT SHIELD 


450 

FLARE 


2703 

FLARE CLOTH 

1086 


TORUS 

801 


SILICONE RUBBER 

420 


CABLES 

45 


PRESSURIZATION/DEPLOYMENT 

105 


CONTINGENCY 

246 


PAYLOAD/FLARE ADAPTER 


390 

ASTRIONICS PENALTY 


25 

RCS PROPELLANT 


6 


TOTAL 

3574 


FIGURE F-l. 0.0-2 AEROBRAKING KIT WEIGHT STATEMENT 

FOR TWO PASS AIRMAT FLARED CONFIGURATION 


F- 7.0 (Continued) 

The candidate aerodynamic decelerator configurations considered in this 
appendix are shown in Figure F-l .0. 0-3. As described above, design, 
stress and weight analyses were only performed for Configuration #1 
(50' diameter, W/C^a ^ 4.2). It is believed that the concept shown 
as Configuration §2 (120* diameter, W/C d a «- 0.74) would weigh less 
than that shown in Figure F-l. 0.0-2 for Configuration #1 (AIRMAT 
Flare). The larger flare (#2) would have a maximum equilibrium temp- 
erature of approximately 800°F. Therefore, a polyimide film could be 
used instead of the high temperature Haynes 188 cloth used in the de- 
sign shown in Figure F-l. 0.0-1, The polyimide film has approximately 
the same strength at 800°F as the Haynes 188 alloy has at 2000°F while 
the polymide's density is approximately 1 /7th that of the Haynes. An 
estimated weight savings of approximately 1300 pounds might be achieved 
by using the larger polymide flare. This potential weight savings is 
directly convertible into round trip payload. Therefore, it is rec- 
ommended that further large polyimide flare design analysis be conducted 
in follow-on activities. 


F-2 



— £0** SPACING 




•1/16" THK* silicone rubber 


-SPRAY COATED WITH 
SILICONE RUBBER - .DOS T^CK 


£ 

^BASIC PACKING 
CONCEPT 


0 


© 

E> 


T ) GOODYEAR AIRHAT MATERIAL; 

HAYNES IBS ALLOY: (11 POROSITY) 

WARP: &4 WIRES PER INCH, 

3GO-.QOG5 IN. OIA. FILAMENTS 
PER WIRE 

WOOF: 112 WIRES PER IN.* 

90-.GOG5 FILAMENTS PER INCH 
CABLE: 100 TWISTED WIRES. RENfc 41 

300’. 0005 IKCH OIA. FILAMENTS PER WIRE 
TORUS SPRAY COATED WITH 
SILICONE RUBBER TO 
PREVENT LEAKAGE. 


AIRhAT WT. * .384 lb/ft z 



SECTION A -A 


£ 02 . 




PAYLOAD 




AERO -BRAKING TUG CONCEPT 
MAX. TEW. * 2000*F 


FIGURE F-l. 0.0-1 LIGHT WEIGHT LARGE FLARE CONCEPT 


D5-17142 



NOTE: ALT. CONFIGURATIONS HAY HAVE 

SLIGHTLY HIGHER MAX. EQ. 

STAG TEMP 9 NOSE BUT TEMP Tuay k 800° F 

OVER MAJORITY OF AREA SHOULD ** 

BE ~ Tmax. p LOCAL~ 2 PS f 

(N/CqA as .74) 


.FIGURE F-bQ.Q-3,. CJWDIMFE LIGHT HEIGHT URGE FLARE CONCEPTS 



D5- 17142 


F-2.0 TWO PASS LIGHT WEIGHT LARGE FLARE PAYLOAD CAPABILITIES 

The two pass payload capability of the AIRMAT flare (50' diameter) was 
computed with the same groundrules utilized for the large nose flare 
(Appendix A) In addition, the estimated payload capability of the 
larger poly imide flare (120* diameter) was computed in the same manner. 
Figure F-2.0. 0-1 shows the capabilities in both the nominal 270 n.m. 
recovery mode and the Shuttle assisted 200 n.m. recovery mode. 



ROUND TRIP PAYLOAD (POUNDS) 

CONFIGURATION 

270 n.m. Mode 

200 n.m. Mode 

AIRMAT Flare (50') 

1800 

3000 

Polymide Flare (120') 

3100 

4300 


FIGURE F-2.0. 0-1 ROUND TRIP PAYLOAD CAPABILITIES 


In either operational mode, these large flares have significant pay- 
load capabilities when used in the two pass mission. Compared to 
those payload capabilities shown in Appendix A for the other two pass 
configurations, this general concept is far superior. Follow-on 
study activities should investigate this type of aerodynamic decelera- 
tion in detail in order to determine all of its ramifications and 
potential . 


F-5