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JPj_ Publication 88-26 



The State of the Art of do Power 
Distribution Systems/Components 
for Space Applications 

S. Krauthamer 


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July 1988 


NASA 

National Aeronautics and 
Space Administration 

Jet Propulsion Laboratory 

California Institute of Technology 
Pasadena, California 


tn to 



JPL Publication 88-26 


The State of the Art of dc Power 
Distribution Systems/Components 
for Space Applications 

S. Krauthamer 


July 1988 


NASA 

National Aeronautics and 
Space Administration 

Jet Propulsion Laboratory 

California Institute of Technology 
Pasadena, California 



The research described in this publication was carried out by the Jet Propulsion 
Laboratory, California Institute of Technology, under a contract with the National 
Aeronautics and Space Administration. 

Reference herein to any specific commercial product, process, or service by trade 
name, trademark, manufacturer, or otherwise, does not constitute or imply its 
endorsement by the United States Government or the Jet Propulsion Laboratory, 
California Institute of Technology. 



ABSTRACT 


This report is a survey of the state of the art of high voltage dc 
systems and components. This information can be used for consideration of an 
alternative secondary distribution (120 Vdc) system for the Space Station. All 
HVdc components have been prototyped or developed for terrestrial* aircraft* and 
space applications, and are applicable for space application with appropriate 
modification and qualification. HVdc systems offer a safe, reliable, low mass, 
high efficiency and low EMI alternative for Space Station secondary distribution. 


iii 



ACKNOWLEDGMENTS 


This work was done for the Program Studies Division, Program 
Requirements and Assessment Office of the NASA Space Station Level II 
Organization, 

The contribution of the following individuals at JPL is 
acknowledged, Linda Palmieri provided the inputs for Chapter 10. John Klein, 
Radhe Das, and Albert Whittlesey reviewed the document and provided helpful 
critiques. Elly Hardy and Esther Graham did an excellent typing job; Pat South 
was the document editor. 

S. Cuk of the Power Electronics Group at the California Institute of 
Technology provided the inputs on the State of the Art of dc/dc Converter, which 
is included in Appendix C. 

Many individuals and organizations provided valuable information 
regarding their activities in high voltage dc distribution systems and 
components. The names of the organizations providing information are listed in 
Appendix A. 


Ed White and Bob Digirolano of the Naval Air Development Center 
provided extensive technical information reflecting their pioneering work on 
High Voltage dc Electrical Distribution Systems for Advance Aircraft. 

Don Teague of the Space Station Program Office provided the relevant 
data from the European Space Agency (ESA) and the National Space Development 
Agency of Japan. Thanks are due to all these individuals and organizations. 

Mukund Gangal Stan Kraut hamer 

Technical Manager Member Technical Staff 

Program Studies Div. Power Electronics Group 

Jet Propulsion Laboratory 

Robert Edelson 

Director 

Program Studies Div. 

Program Requirements and 
Assessments Office 

NASA Space Station 


PRECEDING PAGE BLANK NOT FILMED 






ABBREVIATIONS AND ACRONYMS 


A 

ac 

ALPHA 

BDC 

BETA 

CSD 

dc 

DCSU 

di/dt 

DSVRS 

EMI 

ENG 

EPS 

ESA 

GCU 

GVR 

h 

HVIC 

HVdc 

Hz 

IGBT 

IOC 

JEM 

kHz 

kW 

LVdc 

MBSU 

MCT 

MHz 

MIL-STD 

MILSTAR 

MIU 

MOSFET 

ms 

\iS 

MTBF 

NADC 

NASA 

NASDA 

NM 

NPCU 

NSTS 

NSU 

PDA 

PDCA 

PDCU 

PDCU Module 

PMAD 

POP 

PPS 

PV 


amperes 

alternating current 

Represents the gimbal through which power is transmitted 
Bidirectional Converter 

Represents the gimbal through which power is transmitted 
Constant Speed Drive 
direct current 

Direct Current Switching Unit 
Change of current with respect to time 
Deep Sea Research Vehicle 
Electromagnetic Interference 
Eng ine 

Electric Power System 
European Space Agency 
Generator Control Unit 
Generator Voltage Regulator 
hours 

High Voltage Integrated Circuit 
High Voltage dc (100-300 Vdc) 
hertz 

Insulated Gate Bipolar Transistor 
Integrated Operations Configuration 
Japan Experimental Module 
kilohertz 
k i 1 owa 1 1 s 

Low Voltage dc (Below 100 Vdc) - NASA Std. 28 Vdc 
Main Bus Switching Unit 

Metal Oxide Silicon Controlled Thyristor 

Megahertz 

Military Standard 

Military Strategic Tactical Relay 
Main Inverter Unit 

Metal Oxide Field Effect Transistor 

millisecond 

microsecond 

Mean Time Between Failure 
Naval Air Development Center 

National Aeronautics and Space Administration 

National Space Development Agency of Japan 

Nautical Miles 

Node Power Control Unit 

National Space Transportation System 

Node Switching Unit 

Power Distribution Assembly 

Power Distribution and Control Assembly 

Power Distribution and Control Unit 

Power Distribution and Control Unit in the module 

Power Management and Distribution 

Polar Orbiting Platform 

Power Processor Subsystem 

Photovoltaic 


PAGt y 1 

vii 

PRECEDING PAGE BLANK NOT FILMED 



PWM 

RBI 

REC 

RPC 

RTG 

s 

SAE 

SASU 

SOC 

SPST 

SS 

STC-DBS 

TU 

TWT 

Vdc 

VF 

VT 

W 

W / in^ 


Pulse Width Modulation 

Remote Bus Isolator (interrupter) 

Rectifier 

Remote Power Controller 

Radioisotope Thermal Electric Generator 

second 

Society of Automotive Engineers 
Solar Array Switching Unit 
State of Charge 
Single Pole Single Throw 
Space Station 

Satellite Television Corp. - Direct Broadcast System, RCA (GE) 

Astro-Electronics 

Transformer Unit 

Traveling Wave Tube 

Volts dc 

Variable Frequency 

Voltage-Temperature 

Watts 

Power Density — watts/cubic inches 


viii 



CONTENTS 


EXECUTIVE SUMMARY ES-1 

1. INTRODUCTION 1-1 

1.1 DESCRIPTION OF BASELINE SPACE STATION POWER SYSTEM 1-1 

1.2 ALTERNATE SECONDARY POWER DISTRIBUTION SYSTEM 1-1 

1.3 DESIRED CHARACTERISTICS OF A SECONDARY DISTRIBUTION SYSTEM . 1-5 

1.4 SUMMARY 1-6 

2. BACKGROUND 2-1 

2.1 SECONDARY POWER DISTRIBUTION CONSIDERATIONS 2-1 

2.2 dc EXPERIENCE ON AMERICAN SPACECRAFT 2-1 

2.3 SYNERGISTIC DEVELOPMENT 2-5 

2.3.1 100 Vdc Spacecraft Bus 2-5 

2.3.2 270 Vdc Bus Electrical Systems for Aircraft 2-5 

2.3.3 High Voltage dc Bus Electrical Systems for Submarines . 2-8 

2.3.4 270 Vdc Bus Electrical Systems for Tanks 2-8 

3. SPACECRAFT HIGH VOLTAGE dc SYSTEM TOPOLOGY 3-1 

3.1 SPACE STATION SECONDARY DISTRIBUTION 3-1 

3.2 PLATFORM POWER DISTRIBUTION 3-2 

4. SPACECRAFT COMPONENT CONSIDERATIONS FOR HIGH VOLTAGE DC SYSTEMS . 4-1 

4.1 APPLICATION CONSIDERATIONS 4-1 

4.2 REMOTE BUS ISOLATOR (RBI) 4-2 

4.3 REMOTE POWER CONTROLLERS (RPCs) 4-2 

4.4 POWER CONDITIONERS 4-3 

4.5 ac MOTOR DRIVES 4-3 

ix 


4.6 dc BUS REGULATOR 


4-4 


4.7 CURRENT SENSORS /GROUND FAULT SENSORS 4-4 

5. COMPONENTS TO IMPLEMENT AN HVdc DISTRIBUTION SYSTEM 5-1 

5.1 REMOTE BUS ISOLATORS (RBIs) 5-1 

5.1.1 Contactors 5-1 

5.1.2 Hybrid Power Interrupters 5-2 

5.1.3 Solid State Power Interrupters (RBI) 5-4 

5.2 REMOTE POWER CONTROLLERS (RPCs) 5-8 

5.3 POWER CONDITIONERS 5-8 

5.4 dc BUS REGULATOR 5-13 

5.5 ac MOTOR DRIVES 5-13 

5.6 SENSORS 5-17 

5.6.1 Current Sensors 5-17 

5.6.2 Ground Fault Sensors 5-18 

6. FLAT CABLE IN ELECTRICAL POWER DISTRIBUTION SYSTEMS 6-1 

7. RELIABILITY 7-1 

8. EMI CONSIDERATIONS 8_1 

9. SAFETY CONSIDERATIONS IN HIGH VOLTAGE dc SYSTEMS 9-1 

10. POWER MANAGEMENT AND DISTRIBUTION (PMAD) CONTROL SYSTEM 10-1 

10.1 LOW VOLTAGE dc SPACECRAFT 10-1 

10.2 POWER GENERATION AND STORAGE 10-1 

10.3 POWER REGULATION AND DISTRIBUTION 10-3 

10.4 MICROPROCESSOR FUNCTIONS 10-3 

10.5 CONCLUSIONS 10-5 


X 



11. CONCLUSIONS 11-1 

12. REFERENCES 12-1 

13. BIBLIOGRAPHY 13-1 

APPENDIXES 

A. ORGANIZATIONS INVOLVED IN HVdc COMPONENTS /SYSTEMS A-l 

B. POWER CONTROLLER EFFICIENCY DATA COMPARISON B-l 

C. STATE-OF-THE-ART TECHNOLOGY IN dc/dc CONVERTERS C-l 

Figures 

1-1. Space Station Configuration . 1-2 

1-2. Space Station Electric Power System Functional 

Block Diagram 1-3 

1- 3. SS Manned Core Loads 1-4 

2- 1. Power and Voltage of Various Spacecraft (Flight 

Hardware and Developmental Spacecraft) 2-4 

2-2. Simplified EPS Block Diagram RCA - (STC-DBS) 2-6 

2-3a. Existing Aircraft Electric Power System 2-7 

2- 3b. Proposed Advanced Aircraft Electric Power System 2-7 

3- 1. Space Station Baseline Power System Block Diagram 3-1 

3-2. Space Station Secondary Power Distribution, HVdc Option . . . 3-2 

3-3. Free Flyer Platform Power Distribution (Space Station 

Baseline 20 kHz ac) 3-2 

3-4. Free Flyer Platform Power Distribution (HVdc Alternative 

to Space Station Baseline) 3-3 

5-1. Functional Block Diagram of Hybrid Interrupter 5-3 

5-2. Hybrid Power Interrupter Block Diagram 5-3 

5-3. Typical Solid-State Power Interrupter Block Diagram 5-6 

5-4. RBI Trip Characteristic 5-6 

xi 



5-5. Trip Characteristics for RPC 5-11 

5-6. TESLAco 100 W, 270 V to 5 V Switching Power Supply 5-12 

5-7. Boeing Electronics Company 50 W, 28 V to 5 V dc-to-dc 

Converter Hybridized 5-12 

5-8. dc Bus Regulator (buck type) 5-14 

5-9. Basic Inverter Circuit Diagram 5-16 

5-10. Integrated Actuator System 5-17 

10-1. PPS Functional Block Diagram 10-2 

10-2. Hybrid Switch Block Diagram 10-4 

Tables 

1- 1. Space Station Load Mix 1-5 

2- 1. Spacecraft Bus Voltage Summary 2-2 

3- 1. Secondary Power Distribution Components 3-3 

4- 1. High Voltage dc Distribution Systems Design Objectives . . . 4-1 

5- 1. High Voltage dc Contactors 5-2 

5-2. 270 Vdc Hybrid Power Interrupters (RBI) 5-4 

5-3. Solid-State Power Interrupters (RBI) 5-5 

5-4. Remote Power Controllers (RPC) 5-9 

5-5. Power Conditioners (dc/dc Converters) 5-11 

5-6. Characteristics of dc Bus Regulators 5-15 

5-7. ac Motor Drives 5-15 

5-8. dc Isolated Current Sensors 5-18 

7-1. Component Reliability 7-1 

9-1. Effect of Electrical Currents on Cells and Tissue ...... 9-1 

10-1. Power/Pyro Subsystem - Software Code Estimates 10-6 


xii 



EXECUTIVE SUMMARY 


HIGH VOLTAGE dc (HVdc)* POWER FOR SPACECRAFT 

This report is a survey of the state of the art of HVdc systems and 
components. This information can be used for consideration of an alternate 
secondary distribution system (120 to 160 Vdc) for the Space Station (SS), and 
for platforms. The Space Station requires substantially more power than any 
previous space mission j therefore, previously developed low voltage (28 to 
42 Vdc) systems are not practical, due to mass and efficiency considerations. 

The SS baseline specifications call for a 20 kHz primary and 
secondary distribution system. One hundred and sixty Vdc, generated by photo- 
voltaics and stored in the batteries, will be inverted to single phase, 440 Vac, 
20 kHz for primary distribution, then transformed to 208 Vac for secondary 
distribution, supplying up to 2000 loads on the 208 V bus (Figure ES-1). 

Some users such as the European Space Agency (ESA) and the National 
Space Development Agency of Japan (NASDA) prefer to use 120 Vdc for secondary 
distribution. To implement such a system, the distribution voltage is selected 
within the constraints unique to the spacecraft environment including electrical 
discharge (corona), plasma effects, and parts rating. With these constraints a 
range of 120 to 160 Vdc appears to be practical, thus 120 Vdc has been selected, 
at this time, for consideration as secondary power distribution in the Space 
Station modules. 

HVdc designs for spacecraft may take advantage of previous experience 
with HVdc systems for spacecraft, aircraft, tanks, and submarines, as well as 
commercial work for power supplies and controllers for computers and industrial 
controls. In addition, the control experience with lower voltage (28 to 42 V) dc 
(LVdc ) spacecraft will be useful to the HVdc design. An example of a practical 
implementation of an HVdc spacecraft is the GE (RCA) STC-DBS spacecraft. This 
spacecraft uses a dual bus with one bus operating at 100 Vdc and a 2 kW power- 
rating. This experience places the state of the art close to the Space Station 
platforms requirements. 

All critical components for HVdc distribution and control have been 
developed for terrestrial and aerospace applications. European and Japanese 
spacecraft designers have proposed to use 120 Vdc distribution in their models 
because of the improved efficiency, lower levels of electromagnetic interference 
(EMI), and lower mass. The components for HVdc space power systems have been 
developed to a point where space environment related development and 
qualification can be undertaken. 


ARCHITECTURE / TOPOLOGY 

The proposed architecture for a high voltage dc power secondary 
distribution system for Space Station is shown in Figure ES— 2. Key components 
shown in the figure include remote bus interrupters (RBIs) to interrupt the bus 
under the control of the power management and distribution (PMAD) system, remote 
power controller (RPC) power switches, motor drives, and dc/dc controllers. 

PMAD is not shown. 


*100 to 300 Vdc 


ES-1 


m 

w 

i 

ho 



DCSU 

• DC switching 

• Battery charge/ 
discharge 
control 

• Data interface 


MIU 

• Main inverter 
converts 
160 Vdc 
to 440 V 
20 kHz 


MBSU 

• Power 
processor 

• Remote bus 
isolator 

• Connection and 
distribution 


PDCU 

• Power 
distribution 
and control 

PDCA 

• Power 
distribution 
and control 



TU 

• Transformers 
reduce voltage 
from 440 to 
208 V for power 
inside modules 


NODE 

• Power 
management 

• Switching 

• Fault isolation 
and control 

• DMS interfaces 

• Docking 
interface 



MODULE PDCU 

• Converters 

• Transformers 

• Regulators 

• Fault arrestors 


Figure ES-1. Space Station Electric Power System Functional Block Diagram 





20 kHz 
PRIMARY 
DISTRIBUTION 



USER LOADS 


•See debus regulator 

Figure ES-2. SS Secondary Power Distribution, HVdc Option 


DISTRIBUTION BUS 



Figure ES-3. Free Flyer Platform Power Distribution (HVdc) 


The power distribution HVdc option for the platform is shown in 
Figure ES-3. A dc bus regulator is shown in Figure ES-3 and may be used if the 
dc bus (120 Vdc ) is lower than the array platform voltage. 


SPECIFIC COMPONENTS 

Various components to implement an HVdc secondary distribution 
system are shown in Table ES— 1 amd Figure ES— 2. Components such as batteries 
and photovoltaic arrays are beyond the scope of this report. 


Remote Bus Interrupters (RBI) 

Remote bus interrupters (RBIs) are remotely operated devices for the 
connection and interruption of secondary power distribution. They are controlled 
by a central power management and distribution (PMAD) system. Interruption 
allows routing of power to loads as required and protects against faults and 
overloads. Remote bus interrupters are of three classes: electromechanical 
contactors, solid-state interrupters, and hybrids. 


ES-3 















Table ES-1. Secondary Power Distribution Components 


Component 

Comments 

Flat cable 

See Section 6 

Power conditioners 

dc/dc converters (electrical isolation) 

Load controllers 

Resistive load control (dc/dc con- 
verter, none isolated) 

ac Motor drives 

Inverters (including brushless dc motor 
control) 

Ground fault detector 

Senses bus to ground faults 

Current sensors 

Electrically isolated dc current sensors 

Remote bus isolators (RBI) 

Solid state switch for power bus control 

Remote power controllers (RPC) 

Solid state switch for load control 
switching 


Electromechanical Contractors 


HVdc contactors are produced by Hartman Electric Co. and Eaton Corp., 
Cutler Hammer Division. Models produced are 5 types and 1 type respectively. 
Sizes range from 15 in 3 and 12 oz, to 150 in 3 and 7 lb. Various models are 
current rated between 80 A and 650 A. 


Hybrid Power Interrupters 

Hybrid interrupters incorporate electromechanical contactors with 
solid state devices across the contacts. Four models are produced by Lockheed 
Advanced Marine Systems, Westinghouse, Teledyne Kinetics Division and Eaton 
Corp., Cutler Hammer Division, ranging in size from 30 in 3 and 12 oz, to 
512 in 3 and 14 lb. Various models are rated for currents between 80 A and 
400 A. 


Solid State Power Controllers 


Solid state power controllers have switching times ranging between 
1 ps and 2 ms. Current limits may be preset or variable. Switching life is 
several orders of magnitude greater than contactors or circuit breakers. 
Capabilities which may be designed into a solid state power controller are 
built-in diagnostics, MIL-STD-1553B Data Bus Interface, status feedback, 
bidirectional current limiting and operation, as well as limits on the rate of 
rise and fall of overload current (di/dt). These devices may also match I 2 t 
characteristics to system cables, and monitor current level and leakage. Two 
models are produced by Westinghouse and Lockheed Advanced Marine Systems. One 
device is 863 in 3 and 15 lb. The other is an integral part of an undersea 
vehicle. Current ratings range from 117 A to 150 A. 


ES-4 




Remo te Power Controllers ( RPC ) 


Remote power controllers are devices which replace conventional 
electromechanical relays in the 1 to 30 A range. They are similar in function 
to solid state power controllers (RBI). Mass, volume, efficiency and switching 
life are of critical importance. Some significant characteristics are: a 

controlled rate of current rise and fall, short circuit protection, current 
limiting to prevent load transients, wide temperature range, electrical 
isolation of control and status signals from the power bus, high speed trip-out 
response characteristics, diagnostic capability. Teledyne Solid State, Leach 
Relay, Rockwell, Kilovac, and Westinghouse produce 20 models of RPCs. Sizes 
range from a 0.15 in 3 hybrid, to a unit 16.5 in 3 weighing 14 oz. 


HVdc dc/dc Power Conditioners 

High voltage dc/dc power conditioners for spacecraft are available 
in the 100- to 300- V range. These power conditioners are available with 
efficiencies from 85 percent at 5 V to 90 percent at 15 V and 95 percent above 
100 V. Off-line, high efficiency power conditioners are well characterized, and 
the techniques of their design are applicable to HVdc spacecraft power 
conditioner design. Frequencies >400 kHz allow reductions in power supply mass 
and volume. These low mass converters allow the use of multiple power supplies, 
located at point-of-use . Experience with LVdc spacecraft devices may also be 
applied. Three models are available from TESLAco, Inland Motor, and Space Power 
Inc. Sizes range from 3 to 167 in 3 , and power densities range from 30 to 
300 W/in 3 . 


a c Motor Drives 

The ac motor drives are required to provide appropriate voltage and 
frequency to a motor load such as an actuator on a large spacecraft. They also 
effect conversion of source energy to single- and three-phase power. Waveform 
synthesized, pulse width modulated (PWM) inverter-type motor drives are used 
extensively in commercial, military, and spacecraft applications. These 
inverters are noted for their bilateral characteristics (energy may be fed to 
the motor or returned to the source). These inverters are noted for their high 
efficiency and their minimum usage of power semiconductors. The efficiency 
range of these motor controllers is 92 to 97.5 percent, and their masses are 
low. Three different models have been built by Aeroenvironment Inc., General 
Electric, and JPL. Sizes range from 96 in 3 and 8 lb, to 1152 in and 42 lb, 
while ratings vary from 5 to 35 kW. 


HVdc Bus Regulators 


The dc bus regulators not only control voltage, but can also act as 
dc/dc transformers. These HVdc bus regulators can regulate the spacecraft dc 
bus, provide regulated and adjustable voltage to a heater load, and may be used 
as charge/discharge controllers for a battery subsystem. They provide an 
interface between a source voltage and that of a secondary power distribution 


ES-5 


bus. There are 2 models of dc bus regulators from Aeroenvironment Inc. and 
Space Power Inc. They span a weight range of 0.6 lb to 5 lb, and range in power 
rating from 150 W to 20 kW. Efficiency ranges from 98.5 percent to 95 percent. 


Sensors 


Sensors of two types are examined. Current sensors are required for 
dc bus current monitoring and sensing for protective devices. Ground fault 
sensors identify faults between the power bus and the spacecraft. Three 
manufacturers of dc isolated current sensors offer three models; the smallest 
device being 0.326 in 3 and 0.75 oz, and the largest being 2.55 in 3 and 
1.3 oz. Current ratings vary between 1 A and 1000 A. 

Such devices have been used in terrestrial photovoltaic systems and 
other military applications. 


Flat Cables 


Replacement of round wire with flat cable can reduce cable mass by 
40 percent and reduce temperature rise. Such cable is now available to meet 
spacecraft requirements. 


SYSTEM DESIGN CONSIDERATIONS 

Reliability 

High reliability is a major requirement of an HVdc secondary 
distribution system. Since the high voltage dc spacecraft secondary 
distribution bus has similarity with its low voltage counterparts and HVdc 
aircraft, it can be expected that by using the same ground rules of design, 
reliability will be equivalent or better. 

Because of the limited production of HVdc components, mean time 
between failure (MTBF) data is limited in availability. Available data for HVdc 
components is shown in Table ES-2. 


Table ES-2. Component Reliability 


HVdc Components 

Reliability Characteristics 

Solid State Switches 
Hybrid Switches 
Power Conditioners 

10^-3 x 10^ Switching Cycles 
10^-5 x 10^ Switching Cycles 
10 5 MTBF (h) 


ES-6 




EMI Considerations 


EMI levels that meet standards such as MIL-STD-461B help assure 
trouble-free spacecraft operation, 

MIL-STD-461B specifies allowable current levels over a frequency 
band. It can be inferred that the conducted ripple current level of a power 
condition operating at 28 Vdc and 1 kW, will be similar to a 4.5 kW power 
conditioner at 126 Vdc. This is because the bus current and the power switches 
in the power conditioner will process similar current levels. Ripple voltages 
on the spacecraft bus will be higher because the bus source impedance will rise 
at approximately the same ratio as the voltage. Ripple voltage as a percentage 
of dc bus voltage will remain the same. If additional ripple voltage reduction 
is required i additional filtering can be added. With the usage of flat dc bus 
cable* which represents itself as a transmission line, additional conducted EMI 
suppression is provided due to the cable distributed capacitance. The issue of 
whether the radiated noise from the power conditioner will meet MIL— STD— 461B is 
not known at this time, but the techniques of suppression should be similar to 
that used in the lower voltage designs. Thus, the heritage of EMI control and 
the levels in HVdc systems are expected to be similar to the methods used in the 
LVdc counterparts. 


SAFETY CONSIDERATIONS 

Shock level is related to current passing through the body. Current 
is dependent on voltage level, body resistance, and frequency. Secondary 
distribution voltage level of 120 Vdc is sufficiently low and should not present 
a major hazard to personnel, when used in conjunction with ground fault sensors. 

From a safety point of view, voltage and frequency do have an effect 
on perception and paralysis of humans. At similar voltage levels, the 
sensitivity to dc is approximately equivalent to that at 10 kHz. No unusual 
safety problems at the 120 Vdc level were reported in the literature. It should 
be noted that a 120 Vdc bus and a 20 kHz, 208 Vdc bus appear to be comparable 
with regard to shock hazard. 


Power Management and Distribution 

Power Management and Distribution is well developed for low voltage 
dc space applications which have applications in HVdc power management. Control 
of power sources, batteries, controllers, converters and remote load controllers 
is similar for both LVdc and HVdc, thus LVdc experience is applicable. The LVdc 
experience will also be directly applicable to the unmanned platforms. The 
control of a manned system is obviously more complex. 


CONCLUSIONS 


o Current technology will allow the development of an HVdc secondary 
distribution system for space application. 


ES-7 


o A large number of vendors have the technology and space qualification 

background to produce the required components and design of HVdc systems 
competitively. RCA has built and qualified a spacecraft with a 100 Vdc bus. 

o All components to build HVdc systems have been prototyped or developed for 
terrestrial and aircraft applications. Previous deficiencies in dc 
switching have been resolved. Those HVdc components developed for aircraft 
applications are compact, low mass, high efficiency, and suitable for space 
environment development. 

o Qualification of most HVdc components for space applications is still 
required. 

o Because of the use of HVdc for array and battery subsystems, a number of 
HVdc components have already been developed for space applications, 
including brassboards for Space Station. 

o There exists a substantial 28 Vdc to 50 Vdc spacecraft experience in the 
design of power management and distribution (PMAD) control systems. The 
prototype developments using MIL-STD-1553B control bus for aerospace 
applications have paved the way for HVdc PMAD systems for space station 
applications . 

o The extensive space experience in low voltage dc design heritage allows the 
development of procedures and components for EMI control, system stability, 
and grounding for HVdc systems. 

o Safety issues relating to HVdc are well understood. No unusual safety 
problems were reported in the literature. 

o Users are familiar with design techniques and applications of dc/dc con- 
verters and dc power controls which simplify user interface. 

o The simplicity, low parts count, and previous space qualification history 
of dc systems offer the potential of high reliability for space 
environments . 


ES-8 



SECTION 1 


INTRODUCTION 


1.1 DESCRIPTION OF BASELINE SPACE STATION POWER SYSTEM 

The Space Station will require substantially more power than any 
previous space mission. A conceptual drawing of the Space Station configuration 
is shown in Figure 1-1. The design is baselined [1-1]* for 75 kW average 
capacity with provisions for growth scarring to 300 kW. The three elements of 
the Electric Power System (EPS) shown in Figure 1-2 are: power generation and 

storage, primary and secondary distribution, and utilization. Photovoltaic (PV) 
arrays, shunt regulation, battery charge/discharge controller, dc switching, and 
data interfaces are shown as the functional blocks in the generation and storage 
element. Batteries are used to store power for use during the eclipse period as 
well as during peak demand periods. 

For a number of reasons, including mass considerations, the National 
Aeronautics and Space Administration (NASA) has decided to use 20 kHz frequency 
for primary and secondary distribution [1-2]. An inverter will invert 160 Vdc 
input from a photovoltaic array/battery subsystem to single phase, 440 V, 20 kHz 
alternating current (ac) for primary distribution. A transformer will step down 
the voltage to 208 Vac for secondary distribution. Users will plug into a 208 
Vac bus and convert the power to the appropriate form and level. 

There are a number of different types of loads connected to the EPS. 
Figure 1-3 shows a distribution of loads for housekeeping as well as for end 
user’s applications. The important thing to note is that there may be as many as 
2000, each requiring a mix of transformers, converters, regulators, fault 
isolators, and power factor correction reactors. 


1.2 ALTERNATE SECONDARY POWER DISTRIBUTION SYSTEM 

Some designers [1-3, 1-4] would like to substitute a 120 Vdc system 
for the 208 Vac secondary distribution system in the modules. Voltages of 
100 Vdc to 300 Vdc will be referred to as "high voltage" in this report. They 
claim that the use of High Voltage direct current (HVdc) in secondary 
distribution will improve efficiency, reduce mass, and reduce development risk. 
It is difficult to directly compare the ac versus dc approaches because there is 
little space experience with either system in the range of power required. 
Therefore, the purpose of this report is to survey the state of the art of HVdc 
systems and components, and to provide to the reader, up-to-date and 
comprehensive data. 

A second issue involves the proliferation of dc throughout the Space 
Station. While distribution of the power will be ac, the primary source and 
regulation functions will be dc. In addition, it is anticipated that all 
emergency power will utilize dc components. And finally, the preponderance of 


* Bracketed numbers indicate references listed in Section 12 


1-1 





t 

U) 



DCSU MIU MBSU PDCU 

• DC switching * Main inverter • Power • Power 

• Battery charge/ converts processor distribution 

discharge 160 Vdc •Remote bus and control 

control to 440 V isolator 

• Data interface 20 kHz • Connection and PDCA 

distribution . Power 

distribution 
and control 



TU 

• Transformers 


NODE 
• Power 


MODULE PDCU 
• Converters 


reduce voltage management * Transformers 

from 440 to * Switching • Regulators 

208 V for power # Fault isolation • Fault arrestors 

inside modules and control 

• DMS interfaces 

• Docking 
interface 



Figure 1-2. Space Station Electric Power System Functional Block Diagram 








Power in Watts 


Figure 1-3. SS Manned Core Loads 


loads projected for the Space Station will convert the 208 Vac to regulated dc. 

Dc usage is prevalent throughout the system (Table 1-1). Many electronic end 
loads use low voltage dc. Motors generally use low frequency alternating current 
(ac). Heaters and furnaces can be either ac or dc, depending upon the design. 
Thus, a review of the state-of-the-art components will aid designers throughout 
the system to better utilize available or near term components in their subsystem. 

A third reason is that HVdc distribution may offer mass and 
efficiency advantages for Polar Orbiting Platforms (POP) and future coorbiting 
platforms used in conjunction with the Space Station. Platforms are more like 
conventional spacecraft, with power capacity in the range of 2 kW to 3.8 kW. 
Historically, dc power systems have been chosen over ac systems [1-1]. The 
information provided here may be useful for platform design. 

A fourth reason is that the elements of the emergency power system on 
the Space Station may also utilize dc components. The information collected in 
this report may be useful for that design effort. 


1-4 



Table 1-1. Space Station Load Mix 


Type 

Voltage, 

Volts 

Regulation, 

% 

Frequency/ 

Phases 

Percent of 
Total Load 

Load Type 

1 

120 

5 

400/3 

i 

Lights, small motors 

2 

240 

5 

400/3 

7 

Pumps, motors 

3 

— 

— 

Variable 

20 

Induction heaters 

4 

— 

— 

Any 

frequency 

21 

Heating devices 

5 

5 

2 

dc 

10 

Electrical proc. and 
controls 

6 

15 

2 

' 

dc 

16 

Electronic/ inst . 
devices 

7 

50 

5 

dc 

5 

Controls, devices 

8 

28 

10 

dc 

13 

Critical devices, 
controls 

9 

150 

2 

dc 

3 

Battery processes 

10 

400 

5 

dc 

4 

Transmitters 


1.3 DESIRED CHARACTERISTICS OF A SECONDARY DISTRIBUTION SYSTEM 

Capacity, mass, efficiency, electrical noise (conducted and 
radiated), reliability, maintainability, audio noise, safety, and cost are the 
principal design considerations of the EPS. Table 2-1 shows the current 
experience with different power transmission-distribution systems used on 
spacecraft. It shows that there is a trend toward increased distribution 
voltage. Spacecraft such as the Shuttle (NSTS) have used 400 Hz power systems, 
similar to aircraft power technology, when larger power capacity was needed. 
Unfortunately, the noise environment generated by a 400 Hz system will not be 
acceptable to the Space Station experimenters [1-5]. Therefore, the 400 Hz 
distribution approach will not be considered in this report. 

The military, motivated by reliability and weight considerations, has 
been developing 270 Vdc systems for aircraft [1-6], tanks, and submarines. 
Replacing hydraulic components with electromechanical actuators is attractive 
because it reduces weight while it improves reliability. For similar reasons, 
commercial aircraft manufacturers have also been interested in the 270 Vdc 


1-5 










effort. The systems and components developed for aircraft applications hold 
promise for space applications with proper derating to 120 to 160 V. However, 
only a few of these components have been qualified for the space environment. 

A number of new developments such as smart power conditioners and 
power integrated circuits are leading to very compact high efficiency, dc/dc con- 
verters [1-7]. Internally, these dc/dc converters use an intermediate high 
frequency stage, with isolation transformer and rectifier. Because of the weight 
of the magnetic circuit elements, frequencies have trended upward in order to 
reduce converter mass. The current state-of-the-art converter frequency is in 
the 200 kHz to 2 MHz range. In computers and control applications, truly 
"distributed" power conditioners which can be incorporated onto individual, 
printed circuit boards will improve reliability and reduce costs of 
industrial-commercial equipment. The smart power business is increasing 
steadily and the trend is toward continued growth [1-8] . Dc-based smart motor 
controllers are another development which will have a profound impact on a wide 
range of products [1-7]. Although space environments are far more demanding, 
this commercial-industrial effort toward compact, efficient, and distributed 
power conditioners provides a large pool of technological building blocks and 
know-how to draw upon for designs for space power applications. These efforts 
should result in designs that are cost effective, reliable, and have low 
development risks . 


1.4 SUMMARY 

To summarize, the power capacity needs of the Space Station require 
the development of higher voltage distribution than the voltage levels of exist- 
ing spacecraft. The proponents of dc for secondary distribution claim to have 
technical and cost advantages compared to the baseline ac system, and synergy 
with other industrial and military systems. In the following sections, we shall 
survey the specific developments and components available for 120 to 160 Vdc 
secondary power distribution and control. 


1-6 



SECTION 2 


BACKGROUND 


2.1 SECONDARY POWER DISTRIBUTION CONSIDERATIONS 

The advances in solid state devices and methods of power management 
and control have resulted in the creation of improved methods of power distribu- 
tion systems. Secondary power distribution systems can be either ac or dc. A 
20 kHz primary and secondary power distribution system has been chosen for the 
Space Station [2-1]. An alternative to 20 kHz secondary distribution is the use 
of 120 Vdc distribution for modules and platforms, as described in Section 3. 

A multiplicity of HVdc commercial components is currently available 
in prototype or in development form that can be used to implement an HVdc 
secondary distribution system for space applications. 

It should be noted that at the distribution voltage levels for HVdc 
operation or 20 kHz operation, only few components have been space qualified. 


2.2 dc EXPERIENCE ON AMERICAN SPACECRAFT 

The majority of American spacecraft have had dc bus voltages. Most 
spacecraft, to date, have had power levels below 1 kW. The evolution of most 
spacecraft dc bus voltage generally centered at the NASA standard bus voltage of 
28 Vdc. While some spacecraft utilized a regulated dc bus, others utilized an 
unregulated bus • 

Even with the NASA standard, other spacecraft have had voltage levels 
up to 100 Vdc [2-2]. An American spacecraft bus voltage summary is shown in 
Table 2-1 [2-2]. It is evident that there is a tendency for bus voltages to 
increase as power levels increase. This generally results in lower cable mass 
and improved power conditioner efficiency. Information concerning power and 
voltage of various spacecraft is shown in Figure 2—1; it includes operational 
spacecraft and development types [2—3]. 

European spacecraft (selected types) have flown with a nominal 50 Vdc 
bus. Both European and Japanese spacecraft designers are presently directing 
their development and hardware efforts to a 120 to 160 Vdc spacecraft bus with a 
focus on secondary distribution for the Space Station [1—3]. The higher bus 
voltages have been designed to reduce cable mass, improve power conditioner 
efficiency, and reduce electromagnetic interference (EMI), because of reduced bus 
current . 

Different loads on a spacecraft require different voltage levels. 
These vary from 3 Vdc for logic circuits to levels up to 10,000 Vdc for traveling 
wave tubes (TWTs). Data for Space Station loads are presented in Figure 1-3 and 
Table 1-1. 


2-1 



Table 2—1. Spacecraft Bus Voltage Summary [2-2] 


TRW 

Vdc 

DSCS II (s) 

23.7 to 33 

DSP (s) 

24.25 to 33 

Classified 

24 to 32.5 

Program (s) 
FSC 

20 to 70 

Pioneer 10/11* (s) 

28 +0.6 

HEAO (s) 

23 to 33 

OGO (s) 

23.5 to 33.5 

INTELSAT III (s) 

21 to 31 

TDRS 

22 to 40 


Hughes 

Vdc 

ANIK (s) 

28 to 36 

0S0I (s) 

! 23 to 33 

INTELSAT IV 

23.8 to 48 

WESTAR (s) 

28 to 36 

MARISAT (s) 

28 to 38 

TACSAT (s) 

25 to 31.5 

HASPS (s) 

24 to 33 

LEAS AT (s) 

28.7 to 31.9 

ANIK-C (s) 

28 to 42.5 

SBS (s) 

28 to 42.5 


Ford Aerospace 
and 

Communications 

Vdc 

INSAT I (s) 
SMS* (s) 
IDCSP-A* (s) 
NATO III* (s) 
INTELSAT V (s) 

26.5 to 42.5 
29.4 ±0.2 
29 ±0.6 

29.4 ±0.2 

26.5 to 42.5 


RCA 

Vdc 

IT0S-D (s) 
DMSP* (s) 
Satcom (s) 

STC-DBS 

-26.5 to -36.8 
28 

23.5 to 35.5 
100 ±1V 
23.5-35.5 


GE 

Vdc 

NIMBUS/ERTS (s) 

-24.5 ±0.5 

Japanese Broad- 


cast Sat* (s) 

29 ±1% 

DSCS III* (s) 

28 ±1% 


Lockheed 

Vdc 

LST Study (R) 
Classified Prog (R) 
Type XI (R) 

Agena (Z) 

24 to 32 
24 to 32 
22.5 to 29.5 
24 to 29.25 


GSFC 

Vdc 

MMS (PRU) 
0A0 (PRU) 

21 to 35 
21 to 33 


— 

MSFC 

Vdc 

Sky lab (PRU) 
PEGASUS (Z) 

28 Regulated 
28 Regulated 


2-2 


























Table 2-1. Spacecraft Bus Voltage Summary (Cont'd) 


Fairchild 

Vdc 

ATSF and G* (s) 

30.5 +0.6Vdc 

Shuttle/IUS 

Interface 

24-32 Vdc at 1.7 

kW 

400 Hz, 115/200 Vac, 3$, 390 VA 


JPL** 


Vdc 


MM 71 (Z) 

50 Vac, 30 

Vdc 

MVM 73 (Z) 

50 Vac, 30 

Vdc 

Viking Orbiter 

(Z) 2400 

Hz 50 

Vac, 


30 

Vdc 


Galileo 

2400 

Hz 50 

Vac, 


30 

Vdc 


Voyager 

2400 

Hz 50 

Vac, 


30 

Vdc 



Legend: s = shunt 

Z = zener diodes 

R = relay switching of array 


* = regulated bus DET system 

PRU = array series regulator 
power ** = Mariner TWT operate from 

unregulated dc bus 


With the need of higher power capacity, higher distribution voltages 
are being considered for spacecraft such as Space Station, SP-100 [2—3] , and 
projected electric propulsion systems. The selection of the upper limit voltage 
is determined by phenomena associated with space environments and components 
available in that voltage range. These phenomena [1-4] are: 

( 1 ) Electrical discharge . 

Electrical discharge can cause shorting between conductors and 
subsystems at voltage levels above 280 to 330 V, under space 
vacuum conditions. 

(2) Plasma discharge . 

Plasma discharge can cause power loss in the array at 250 run 
altitude. The effect is that the plasma acts as a resistor 
shunting the photovoltaic array. Power losses at 400 Vdc may 
be one percent or more of rated array power. Power loss begins 
to become noticeable above 160 Vdc. 

(3) Parts . 

Certain critical parts that are voltage limited include tan- 
talum capacitors, semi-conductors, and power devices. Problems 
occur in parts availability for operation above 160 V. How- 
ever, design methods are available to overcome the voltage 
limitations of parts. 

Considering these problems, selection of a voltage of 160 Vdc for 
the SS, the array, and the batteries, and 120 Vdc for secondary distribution 
inside the modules is considered to be conservative. 


2-3 





maximum out fur power 


100,000 


10,000 


1000 


100 


18 kV PEAK 


/*- 


P 3 (180V OUTPUT) TO J7SV 


MA#n«/Nsrc 

P 3 (T?kW AVERAGE 1 


‘ ) 


TRW IRAQ ?5KW POWER STSTEH 
TCC 6.< KW BREADBOARD 


\ 


PUCHT PROVEH/FLIGW 
READY HA8QWARE 


GSFC/MMS IATTERY 
CHARGER 

j f 

(34 V OUTPUT) 

^ SHUTTLE/ 1 US* 
(30V OUTPUT) 


SKYLA* AIRLOCK I 
MODULE LINE REGULATOR 
(78 V OUTPUT) j 


MASS BOARDS 


/ 


/ 



T*WA«RC ION ENGINE 
POWER PROCESSOR 
MUTIPLE HI-VOLTAGE OUTPUTS 

fcT“ 

400 V 


/ 




• STC-DBC (RCA) 
(100V OUTPUT) 


/ 


APOLLO TELESCOPE MOUNT 
CHARGER ROOSTER REGULATOR MOOULE 

a * 1 -t t 

(28 V OUTPUT) / 

, MARINER LINE REGULATOR 

(SO V OUTPUT) 


/ 


FSC TRANSMITTER CONVERTERS 
(28 V OUTPUT) 

/ 

/ 

/ I 


_L 


100 200 
input voaAGt (vars do 


300 


350 


Figure 2-1. Power and Voltage of Various Spacecraft [2-3] 
(Flight Hardware and Developmental Spacecraft) 


2-4 


2.3 


SYNERGISTIC DEVELOPMENT 


Various activities are in progress in developing HVdc distribution 
systems [2-4, 2-5] for spacecraft, aircraft, submarines, and tanks. HVdc 
systems require smaller cables and lighter weight components because of lower 
currents. Also, in many instances, bulky hydraulic components can be replaced 
with lighter weight electrical components. Electrical systems provi e lg er 
efficiency than hydraulic systems. Moreover, reliability improves wi 
electrical actuation [2-4] and control, and has been shown to be better with dc 
systems because of fewer components. These efforts have resulted in the 
technology and components which can be applied to HVdc distribution systems o 
space applications [2-4]. Development of system and components for HVd ^ SY 8 ^ 1 " 5 
has taken place over the past 15 years, mostly in the last six years. ’ 

270 Vdc distribution systems are being developed for spacecraft, aircraft, 
ground-based applications. Much of the material in this report summarizes the 
state of the art of these advanced distribution components developed for HVdc 
applications. A brief synopsis of these synergistic activities which have 
future application in spacecraft systems, follows. 


2.3.1 


100 Vdc Spacecraft Bus 


RCA has developed a spacecraft (STC-DBS) using a dual bus voltage 
(100 Vdc and 35 Vdc). The topology of this spacecraft is shown in Figure 2 1 
[2-2]. The 100 Vdc bus powers 6 TWTAs. Relays are used to switch the TWTs and 
heater loads. The 100 Vdc array generates 2 kW. RCA has qualifiedthe eac 
KCL-02A relay rated at 25 A and 35 Vdc to 100 Vdc operation. The TWT power 
supplies were also qualified at the 100 Vdc level. The spacecraft has been 
completely tested and qualified for space flight, and is ready to be launc ^ d ; 
The important consideration is that space-qualified relays and o er compo 
have been developed and qualified for the STC-DBS spacecraft. 

2 . 3.2 270 Vdc Bus Electrical Systems for Aircraft 

Aircraft electrical power distribution has changed very little ° v ® r 
the past number of years, even though substantial advancements have occurred in 
computers and communications equipment. The components used in the aircratt 
power distribution system and the methods of switching and protection, 
especially the protection of cable and wire, have had very few characteristic 
changes over the past two decades* 

The physical size of modern aircraft requires long cable runs 
between the control (manual switches) and remote loads. A hydraulic-type 
constant speed drive, driving a 400 Hz generator, has been used to provide fi 
frequency, three-phase aircraft power. Most military and commercial aircraft 
have used this method of secondary power generation (Figure 2-3a). The use o 
remote control circuit breakers has resulted in the reduction of cable mass and 
voltage drop. Circuit breakers are selected to have trip characteristics that 
allow handling of worst case overloads, transient inrush current, transient 
overloads, and cable protection without nuisance tripping. 


2-5 



2-6 


NORTH 

| 

NORTH 

ARRAY ORIVE 


ARRAY DRIVE 

ELECTRONICS 


ASSEMBLY 


Ho 


100 V 
SOLAR 
ARRAY 


-to 


hto 


35V 

SOLAR 

ARRAY 


--am 

hs 


E 


100V BUS 



1 

100V 

- I 

PARTIAL 


SHUNT 

I 

ASSEMBLY 

1 

■ 


Ol— E4S 


100 V 
SHUNT 
CONTROL 
ELECTRONICS 


-L_^°±J u 


I 


TYP. 1 OF 6 


£ 


35V 

PARTIAL 

SHUNT 

ASSEMBLY 

1 

II 


n 





*1 


LOW VOLTAGE (35V) BUS 



TWTA 

WARM-UP 

HEATER 


3SV 
SHUNT 
CONTROL 
ELECTRONICS 


SOUTH 


SOUTH 

array drive 


ARRAY ORIVE 

ELECTRONICS 

1 

ASSEMBLY 


i 


BATTERY A 
CHARGE 
CONTROLLERS 


i 


BATTERY A 


i 




BATTERY B 
CHARGE 
CONTROLLERS 


BATTERY Bl 


J 




PYRO 
BUS A 


J 


PYRO 
BUS B 


TO LOW 
VOLTAGE BUS 
LOADS 


Figure 2-2i Simplified EPS Block Diagram RCA - (STC DBS) 








Elimination of the hydraulic constant speed drive (Figure 2-3a) 
results in a system in which a variable frequency, regulated voltage is 
generated. This three-phase voltage, when rectified, produces a 270 HVdc bus. 
The configuration for the aircraft high voltage dc secondary distribution system 
is shown in Figure 2-3b. 

The desire for reduction in aircraft weight, replacement of 
mechanical switches with solid-state devices, elimination of the hydraulic 
constant speed drive, as well as future computer control concepts, has been the 
goal of this ongoing development activity [2-5 through 2—10]. 

The scope of activities in the HVdc area encompasses the planning 
phase of new aircraft, breadboarding of HVdc power distribution systems, test 
and development of systems and subcomponents. The Society of Automotive 
Engineers, SAE standard No. AS1931 [2-7] defines the requirements for HVdc 
systems for aircraft. In addition, extensive design, development, and hardware 
efforts have been made in the areas of power system stability analysis, flat 
cable, safety issues, EMI, fault protection, power management, and control 
[2_n t 2-12]. At present, no firm decisions have been made to build an aircraft 
using HVdc bus. While technical problems still must be resolved, the main 
impediments to development are costs. The HVdc aircraft activities involve many 
organizations and manufacturers in the development and manufacturing of 
components and systems (APPENDIX A). 


CONSTANT SPEED DRIVE 


ENG 




cso 

GEN 





i 

* 

GCU 







115V, 400 Hz 


Figure 2-3a. Existing Aircraft Electric Power System 


HIGH VOLTAGE DC 



270 VDC 


Figure 2-3b. Proposed Advanced Aircraft Electric Power System 


2-7 









2.3.3 


High Voltage dc Bus Electrical Systems for Submarines 


Submarines have used high voltage dc for high power distribution 
systems and have developed the technology to implement these designs. 

Switchgear, contactors, and other parts of the system utilize standard Navy-type 
devices. Most of the equipment utilized would not be applicable to space, due 
to the large mass and size of components. Another class of power distribution 
is being developed in the 100 kW power range. A unit of this type is being 
built for the Navy at the Newport News Shipbuilding Company [2-13, 2-14]. This 
distribution system operates at 155 Vdc and has been designed to operate lower 
power submarine electric loads. The system is designed with an ungrounded bus 
where cables, circuit breakers, and contactors are standard Navy-type 
components. Ground fault sensors are used in order to detect ground fault. 

Dc/dc converters are used to interface with loads. 

A Navy standard, "Military Standard Interface for Shipboard Systems 
MIL— STD— 1399, " has been issued to define the 155 Vdc operating bus for shipboard 
systems [2-15] . 

2.3.4 270 Vdc Bus Electrical Systems for Tanks 

There is an ongoing activity in developing tanks with an HVdc bus. 
The HVdc power distribution system for the Electric Gun and Turret Drive (EG/TD) 
has been breadboarded and tested at General Dynamics Land Systems Division 
[2-16]. The system is currently in the brassboard phase. Power rating will be 
in the range of 45 to 60 kW. The distribution system is similar to that used in 
the design of HVdc aircraft buses. The rationale for utilizing HVdc for this 
tank was to reduce mass of the distribution system components and cable. 

In this distribution system, three-phase, ac generator output is 
rectified to obtain an HVdc distribution bus. Solid state remote bus isolators 
(RBIs) are used in the generator output lines in order to replace 
electromechanical components. Solid state RBIs and remote power controllers 
(RPCs) will be used for switching and protection functions. Dc/dc converters 
will be used for regulating load voltage. 


2-8 



SECTION 3 


SPACECRAFT HIGH VOLTAGE dc SYSTEM TOPOLOGY 


3.1 SPACE STATION SECONDARY DISTRIBUTION 

The planned primary distribution of power on the space station will 
be at 440 V, single phase, and 20 kHz [1-3, 2-1]. The planned baseline 
secondary distribution is 208 Vac, single phase. The Space Station power 
system, including the primary and secondary power distribution systems, is shown 
in Figure 3-1. The array voltage in Figure 3-1 is distributed to a 20 kHz 
inverter and to a battery subsystem and controller at 160 Vdc (the 6ame level as 
discussed in Section 2). The battery subsystem charge controller is used to 
charge the batteries. The array shunt regulator and the battery discharge 
controller are used to regulate the dc bus voltage at 160 Vdc. All the 
equipment developed for the array/battery subsystem including charge/discharge 
controller, RBIs, and dc/dc power conditioners, can be used in the HVdc 
secondary power distribution option (208 V, 20 kHz, single phase) in a modified 
form, or as is. 




$ SECONDARY DISTRIBUTION 
208 Vac 


LOAD 

INTERFACE 


Figure 3-1. Space Station Baseline Power System Block Diagram 


A detailed configuration of the space station HVdc secondary 
distribution option is shown in Figure 3-2. The selection of a spacecraft 
secondary distribution dc bus voltage is determined by a number of factors 
[3-1]. Many of the components and systems to implement this secondary 
distribution bus are detailed later in this report. On the array side of the 
spacecraft, as previously discussed, voltage selection is made at a level that 
will tolerate the effects of plasma discharge and electrical discharge and allow 
the use of available electronic parts. A level below 200 Vdc is considered 
safe, in terms of plasma discharge. Space Station nominal array voltage and 
power generation is 160 Vdc and does not exceed 200 Vdc under extremes of 
environmental conditions or load operation. On the secondary distribution side, 
120 Vdc is considered a conservative voltage selection. At that voltage, 
electrical discharge (corona) from outgassing effects does not become a 
problem. Also, the required components are readily available at the 120 Vdc 
level. In addition, 120 Vdc is considered safe from a human factors safety 
point of view [3-2]. 


3-1 





20 kHz 
PRIMARY 
DISTRIBUTION 


REDUNDANT dc FLAT 

POWER BUS REGULATOR CABLE 



USER LOADS 


*Sm debut regulator 


Figure 3-2. Space Station Secondary Power Distribution, HVdc Option 


Proper grounding practice must be exercised in HVdc for modules and 
platforms. HVdc platform grounding practice should follow those procedures 
developed for LVdc spacecraft. Grounding practice for HVdc modules is 
simplified by using a dedicated isolation transformer between the primary 
distribution bus and the module. This allows for single point grounding of all 
the modules and will satisfy the requirements of the Space Station Grounding 
Standard [3-3] . 

3.2 PLATFORM POWER DISTRIBUTION 

208 Vac 20 kHz power distribution is baselined for free flying 
platforms [Figure 3-3] . It uses an architecture similar to that of the Space 
Station. 


DISTRIBUTION BUS 



Figure 3-3. Free Flyer Platform Power Distribution 
(Space Station Baseline 20 kHz ac) 


3-2 














An HVdc alternative to the baseline is shown in Figure 3-4. It 
shows power generation and distribution at 120 Vdc. If a 160 Vdc array vo age 
is used as for the primary power system, then a dc/dc converter wi e requ 
to reduce the 160 Vdc array voltage to the desired 120 Vdc level. 

Remote bus interrupters are used to disconnect sections of the 
distribution system and subcomponents for maintenance. In addition, a c us 
regulator is shown as a means of matching the bulk array voltage to the power 
distribution voltage in order to provide tight bus voltage regulation. 


DISTRIBUTION BUS 



Figure 3-4. Free Flyer Platform Power Distribution (HVdc 
Alternative to Space Station Baseline) 


The battery charge/discharge controller subsystem has been detailed 
in previous studies for Space Station, and since the components developed are 
applicable to the module and the platform power distribution system, y 
not be discussed. The remaining components, other than user oa s, are 
Table 3-1. Techniques of power management and control are described in 

Section 10. 


Table 3-1. Secondary Power Distribution Components 


Component 

Comments 

Flat cable 
Power conditioners 
Load controllers 

ac Motor drives 

Ground fault detector 
Current sensors 
Remote bus isolators (RBI) 
Remote power controllers (RPC) 

See Section 6 

dc/dc converters (electrical isolation) 
Resistive load control (dc/dc con- 
verter, none isolated) 

Inverters (including brushless dc motor 
control ) 

Senses bus to ground faults 
Electrically isolated dc current sensors 
Solid state switch for power bus control 
Solid state switch for load control and 
switching 


3-3 













The types of loads used will determine the types of electrical 
components used and their power ratings. Secondary distribution systems will be 
in the modules and in the platforms. Table 1-1 provides the percent 
distribution of the types of loads in the Space Station, including modules and 
platforms. Components are sized to the required load applications. Component 
details are discussed in Section 5. 


3-4 


SECTION 4 


SPACECRAFT COMPONENT CONSIDERATIONS FOR HIGH VOLTAGE DC SYSTEMS 


APPLICATION CONSIDERATIONS 


The topology of a typical HVdc secondary power system was shown in 
Figure 3-2, and the components used to implement this topology were shown xn 
Table 3-1. Each of the component groups tabulated performs par icu a 
and eenerally has a range of operational characteristics associated with the . 
Some of the important design objectives to implement a highvoitagedc secondary 
distribution system are shown in Table 4-1. Examples of “'=“"8 HVdc 

up to 25 kW. 

Different types of loads require different power controllers. Three 
types of power controllers are discussed in this report. First, there are power 
conditioners which are dc/dc converters with electrical isolation and P™^ de 
the required regulated output voltage. Second, there are d ' ^°“ Vdc 

which are used to interface a high voltage bus to a lower voltage bus (160 Vdc 
to 120 Vdc), typically without electrical isolation. They are also used 
control heater^and resistor loads. Third, there are motor controllers which are 
used to control the speed of three-phase and single-phase ac motors, typica y 
400 Hz types. The Space Station load mix is shown in Table l-i. 


Table 4-1. High Voltage dc Distribution System Design Objectives 


o High Reliability (low parts count) 

o High Efficiency o System 

o Power conditioners and 
other components 

O Low EMI (Based upon previous low voltage spacecraft experience, 
MIL-STD-461B plausible) 

o Solid State power switching 

o Safety 

o Flat cable for minimum mass, dc voltage selection for low cable mass 

o Module Operation; emergency power /uninterruptible power supply to back up 
critical loads 

o Effects of plasma discharge and electrical discharge should not alter 
system efficiency and reliable operation 


4-1 




4.2 


REMOTE BUS ISOLATOR (RBI) 


The term RBI defines an interrupter that can be remotely controlled 
to connect and interrupt secondary distribution power. Historically, 
interruption of distribution power was accomplished with circuit breakers. They 
were designed to interrupt ac and dc loads. Circuit breakers are being replaced 
by new classes of electronically controlled power switching devices. The 
requirement of this class of solid state, hybrid or electromechanical 
interrupters is that they can be controlled by a central power management and 
distribution computer (PMAD), typically interconnected to a MIL-STD-1553B data 
bus [4-1]. Current levels for the secondary power distribution systems may vary 
from 25 A to 200 A, and voltage levels from 120 Vdc to 160 Vdc, according to the 
module or platform rating. The interrupter will allow the connection or 
disconnection of different sections and components of the distribution system. 
Topologies using RBIs are shown in Figures 3-2 and 3-4. The RBI will permit the 
maintenance of parts of the distribution system that are redundant and will 
allow the rerouting of distribution power to loads, as required. 

Interruption of fault current and overload current is desired in 
order to protect the interrupting devices, cables, generation sources, storage 
components, and critical load components. 

The Space Station array/battery subsystem HVdc distribution system 
uses solid state power interrupters. High reliability is essential and low 
power losses are needed, in order to maintain high systems efficiency. Mass and 
volume should be low and consistent with spacecraft performance requirements. 

There are three classes of remote bus isolators (RBI) that can 
connect and interrupt dc bus power. These are: 

(1) Contactors. 

(2) Hybrid Power Interrupters (Solid State and Electromechanical 
combinations). 

(3) Solid State Power Interrupters. 

Each of the above classes of isolators/interrupters has its own 
particular characteristics, which are described in Section 5.1. 


4.3 REMOTE POWER CONTROLLERS (RPCs) 

The term RPC defines an interrupter that can be remotely controlled 
to connect and interrupt a load. Solid state RPCs are similar to solid state 
RBIs except for differences in function and current rating. Historically, 
interruption of both dc and ac loads was accomplished by relays. 

Relays (some of which are space qualified for HVdc [2-2]) are being 
replaced by new classes of electronically controlled solid state power switching 
devices [4-2 and 4-3]. The requirement of this type of interrupter is that they 
can be controlled by a central power management and distribution (PMAD) 
computer, typically interconnected to a MIL-STD-1553B data bus. Current levels 


4-2 



for loads in the Space Station secondary distribution system may vary from 1 A 
to 25 A, and voltage levels from 120 Vdc to 160 Vdc, according to a module or 
platform voltage rating. RPCs are used to connect and disconnect user power 
conditioners, ac motor drives, and other components to the secondary 
distribution bus. An application of the RPC is shown in Figures 3-2, 3-3 and 
3-4. 


Interruption of fault current and overload current is desired in 
order to protect the interrupting devices from excessive currents, associated 
cable, inrush surges on the distribution bus, and critical load components. 

Detailed characteristics for RPCs are shown in Section 5.2. 


4.4 POWER CONDITIONERS 

Power conditioners are used to interface the secondary distribution 
bus and provide a load with the desired voltage and regulation characteristic. 
They also regulate voltage to the load and provide electrical isolation. Low 
mass and high efficiency are design objectives. Various topologies are used in 
power conditioner design [4—4] . All of the techniques used utilize an internal 
switching frequency. The higher the switching frequency the lower is the mass 
of the magnetics. The optimum band for switching frequency, which is dependent 
on magnetic hysteresis losses, is between 200 kHz and 1 MHz [Appendix C] . Power 
conditioners mass and volume can be reduced compared to 20 kHz designs by 
operating the power conditioner at higher switching frequencies (above 100 kHz) 
while still maintaining high efficiency. Power conditioner design also requires 
low EMI levels consistent with spacecraft EMI requirements (typically 
MIL— STD-461B) . Generally, dc power conditioner designs are similar to those 
used on 28 Vdc buses, with the exception that the components must operate at 
higher voltage levels. Load voltage requirements are generally in the +5 Vdc 
and + 15 Vdc range. Power ranges from 25 W to 50 kW are described in Appendix C 
and Section 5.3. 


4.5 ac MOTOR DRIVES 

Many loads such as actuators on larger spacecraft require ac motor 
drives to provide appropriate voltage and frequency to a motor load [4-5] . The 
efficient conversion of spacecraft primary distribution or secondary distribu- 
tion power to single- and three-phase power is required. Motor control may be 
at a fixed voltage and frequency or it may be at a variable voltage and variable 
frequency. Variable f requency/variable voltage control provides a means of 
accelerating a motor load without large inrush currents. Constant torque 
operation requires a constant volts /hertz motor drive. Waveform synthesis, 
bilateral operation, and regulation of voltage and frequency are important 
parameters in the design of an ac motor drive. Waveform synthesized, pulse 
width modulated (PWM) inverter type motor controllers are used extensively in 
commercial, military, and spacecraft applications. These inverters are noted 
for their bilateral characteristics (energy may be fed to the motor or returned 
to the source). This class of inverters is noted for its high efficiency and 
the minimum usage of power semiconductors. 


4-3 


Electrical isolation from the spacecraft bus is not required for 
driving motor loads because of EMI or safety concerns. Since this class of 
inverters is bilateral in nature, complex power and logic control circuitry and 
components are unnecessary to implement these design topologies. State-of-the- 
art ac motor drives are described in Section 5.5. 


4.6 dc BUS REGULATOR 

The dc bus regulator provides the function of interfacing an array or 
voltage source at one voltage level to a voltage level required by the secondary 
power distribution bus. An example of this is conversion of 160 Vdc to 120 Vdc 
or from a low voltage to a higher voltage as in a battery discharge controller. 
The dc bus regulator is not typically used in 28 Vdc spacecraft. Regulated dc 
voltage is supplied by the array/battery subsystem as shown in Figure 3-2. The 
added use of a bus regulator will provide improved distribution voltage regula- 
tion, generally low source impedance with improved EMI and transient charac- 
teristics. The same type of dc bus regulator can also be used to control resis- 
tive and heater loads. It should be noted that these converters do not have 
electrical isolation since it is not required for the function performed, and as 
a result of eliminating a transformer, efficiency and mass of the regulator is 
improved. State-of-the-art dc bus regulators are described in Section 5.4. 


4.7 CURRENT SENSORS/GROUND FAULT SENSORS 

Electrically isolated current sensors are used to monitor and sense 
the dc current. Such a measurement is critical to protect and control current 
within the distribution and conversion system. Electrical isolation provides a 
means of isolating the secondary power bus from interfacing data and control 
buses . 


Ground fault sensors provide a means of identifying faults from the 
power bus to the spacecraft structure, and typically interface through a 
MIL-STD-1553B data bus to the spacecraft power management and distribution 
system control. State-of-the-art sensors are described in Section 5.6. 


4-4 



SECTION 5 


COMPONENTS TO IMPLEMENT AN HVdc DISTRIBUTION SYSTEM 


This section presents examples of HVdc components that have been 
developed for Space Station, aircraft, and other applications. All the 
described components have the potential to be used in space vehicle electrical 
power distribution systems. The list of devices and manufacturers was developed 
to reflect the broad scope of activities previously and currently in progress. 
The list is representative of available components and does not encompass all 
ratings and manufacturers of these types of components. 

Many of the components described have been designed and tested for 
extremes of environmental and electrical conditions. The components described 
reflect a high level of development including technical and performance capa- 
bilities, and many can be applied to space applications. All will need space 
qualification testing and may need redesign for the space station application. 
Most components appear to be viable candidates for an HVdc secondary 
distribution system [5-1]. The steps required to adapt these components for 
space usage are as follows: 

(1) Modification for operation at the required voltage and current 
levels . 

(2) Modification of the design for thermal conditions and heat 
rejection in the space environment. 

(3) Selection of appropriate space-qualified parts to perform the 
required function. 

(4) Modification, as required, for operation per MIL-STD-461B. 

(5) Qualification to the appropriate reliability level required. 

5.1 REMOTE BUS ISOLATORS (RBI) 

5.1.1 Contactors 

Contactors are electromechanical relay devices that close or open 
contacts to supply load power or interrupt load. Contactors are useful to 
provide a positive disconnect of a particular section of a bus if maintenance is 
required. The contactors needed for space applications differ from the conven- 
tional submarine and locomotive contactors in that the means of suppressing the 
arc generated in opening the contact is accomplished with magnetic blowouts. 
Using magnetic materials such as samarium cobalt results in a low mass and 
volume contactor. For space applications, special means are necessary to 
provide for arc extinguishing [5-2], such as hermetic sealing of the contactor 
with container pressurization. Contactors are not designed to protect 
spacecraft cable or connected loads in the event of a fault. If the purpose of 
the contactor is to provide disconnect for bus maintenance and not for opening 
or closing of current, then special hermetic sealing and pressurization may not 


5-1 



be required. Examples of existing contactors for this application are shown in 
Table 5-1, with current range from 80 to 650 A and voltage levels up to 340 Vdc. 


Table 5-1. High Voltage dc Contactors 


Manufacturer 

Voltage 

Rating 

(Vdc) 

Current 

Rating 

(A) 

Size, in. 

Weight 

(lb) 


326 

80 

5.5 x 4.5 x 4.3 

2 

(Hermetic 

Seal) 


270 

100/600 

surge 

3 x 2 x 2.5 

12 oz 

Hartman Electric Co. 

270 

200/600 

surge 

4 x 2 x 3.25 

20 oz 

[5-2] 

270 

400 

nonbreak 

5.5 x 4 x 4 

2.75 


340 

650 

7.1 x 5.3 x 4 

7 

Eaton Corp.,Cutler 
Hammer Division 
[5-3] 

270 

250 

6x6x4 

6 


5.1.2 Hybrid Power Interrupters 

Hybrid power interrupters incorporate electromechanical contactors 
with solid state devices across the contactor contacts, and are used in 
applications where distribution load currents are switched. A functional block 
diagram of a high voltage dc hybrid power interrupter is shown in Figure 5-1, 
and a block diagram of control and system elements is shown in Figure 5-2 
[5-4]. In this interrupter, the main contacts of an electromagnetic contactor 
are shunted by a power semiconductor switch. Prior to disconnecting a load, 
both the mechanical contacts and the semiconductor switch are on. When a signal 
is provided to turn off the hybrid switch, the mechanical contact is opened and 
the load is carried by the solid state switch. After the mechanical contact is 
open, the solid state switch is turned off. For hybrid power interrupter 
turn-on, the solid state part of the power switch is turned on to carry the 
load, after which the mechanical contactor is energized (closed). The load is 
then transferred from the solid state switch to the mechanical contact. Load 
current is carried by the mechanical contacts in order to reduce power loss. 
These interrupters have provisions for both input and output overvoltage 


5-2 












5-3 

















suppression. They are actuated by a control input signal and provide a failure 
signal output. The interrupter can be interfaced to microprocessor control* 
Internal logic power, control logic, leakage current detector, and diagnostics 
are also shown. 

Other methods of implementation of this design differ according to 
the desired function. Eaton Corp., Cutler Hammer Division allows for 
interruption of multiple thousands of amperes of fault current duplicating in 
some fashion the characteristics of a circuit breaker [5-5]. Lockheed allows 
interruption of a certain level of overload current which is characteristic of 
the contactor rating [5-4]. It should be noted that both of these interrupters 
are readily adaptable to microprocessor control. Examples of the current rating 
and surge rating for four separate manufacturers are shown in Table 5-2. The 
sizes shown range from 80 to 400 A dc at a voltage rating of 270 Vdc. 


Table 5-2. 270 Vdc Hybrid Power Interrupters (RBI) 


Manufacturer 

Current 

Rating 

(A) 

Surge 

Rating 

(A) 

Size, in. 

Weight, 

lb 

Lockheed Advanced 
Marine Systems 
[5-4] 

400 

600 

8x8x8 

14 

Westinghouse 

[5-6] 

80 

100 

5.5 x 3.5 x 6 

3.5 

Teledyne Kinetics 

Division 

[5-7] 

100 

1500 

3 x 2.5 x 4 

3 

Eaton Corp., Cutler 
Hammer Division [5-5] 

100 

1500 

4x4x2 

3 


5.1.3 Solid State Power Interrupters (RBI) 

Solid state power interrupters offer performance capabilities, some 
of which are not found in circuit breakers, contactors, or hybrid interrupters 
[5-8]. Two manufacturers of solid state power interrupters are shown in 
Table 5-3. The two units described have been designed for different 
applications. The Westinghouse unit was designed to be used on the Space 
Station, and the Lockheed unit was designed for undersea applications. Figure 
5-3 shows a typical solid-state power interrupter block diagram. The main 
components of this interrupter consist of a semiconductor power switch, driver, 
current sense, voltage sense, over current trip, monitor circuits, and a data 
(MIL— STD-155 3B ) interface unit. The rating of this interrupter developed by 
Westinghouse Electric Corporation is 150 A at 150 Vdc (22.5 kW). A general 


5-4 







Table 5-3. Solid-State Power Interrupters (RBI) 


Manufacturer 


Westinghouse 


[5-9] 


Current 

Rating 

dc 


150 SPST 


Voltage 

Rating 

dc 


150 


Switch 

Opening 


10 ms 
MAX 
(Fig- 
ure 5-4) 


Size, 

in. 


11.5 

x 

15 

x 

5 


Weight 

lb. 


15 


* 


Comments 


(See 5.1.3, Design 
Characteristics 
listing) 
o Current limit 
level adjustment 
o Analogue moni- 
toring circuits 
V, I, current set 
o Digital monitoring 
circuits 
o switch status 
o trip status 
o current sensing 
polarity 
o I/O interface 
circuits 


Lockheed 

Advanced 

Marine 

Systems 

[5-10] 


117 


150 


300 A 
surge 
<1 s 


Part of DSRVs 
vehicle 


o 

o 

o 

o 


Current limiting 
at 4 x load currenl 
Opening of switch 
at 10 A/ jis 
Closing of switch 
>10 A/ps 

Trip and loss of 
saturation trip 


description of generic RBI characteristics is shown below. Many of the 
characteristics described are those for Westinghouse Electric Corporation Part 
No. ED406490-1 [5-9]. The critical characteristics are current limiting, short 
circuit protection, controlled current turn on and turn off, and low power 
dissipation. These characteristics will differ according to the manufacturer 
and RBI requirements. 


The design characteristics for solid state RBIs are as follows: 


(1) Short Circuit Protection . High speed fault clearing capability 
is provided up to the current and thermal limit of the solid 
state switches used. Typically, under short circuit 
conditions, switch opening time will be one microsecond. At 
other levels of overload current, turn-off time will vary up to 
2 milliseconds. This operating time is coordinated with the 
curve shown in Figure 5-4. The voltage drop across the RBI is 
a function of current through the device. This curve is used 
to define a safe operating mode for the device. 


5-5 




Timm tmi t I immm—%4m ) 

Figure 5-4. RBI Trip Characteristic [5-9] 



61 







(2) Current Limit . Current limit if designed in the interrupter 
will limit current to a prescribed level. This level can e 
pre-set or adjusted by the PMAD computer. Turning the switch 
on into a capacitive or resonant load will limit current to 
adequately charge up a bus capacitor without interrupter trip. 

(3) Switching Life . Switching life is several orders of magnitude 
greater than contactors or circuit breakers. 

(4) Built -in Diagnostics . Internal provision can be made to 
identify component failures. Internal signals can provide 
information as to failure in the solid state interrupter or in 
other parts of the proven distribution system. 

IS) Status Feedback. Status feedback signals will provide 

monitoring of the various operational states of the interrupter. 


( 6 ) 


Bidirectional Current Li miting . The solid state power 
interrupter can be designedTor bidirectional operation with 
bidirectional current limiting (Table 5 2, Westinghouse) . a 
this functionally implies is that current could flow in either 
direction and can be provided with short circuit interrupting 
capability and current limit in both directions. This fea ure 
is useful for system protection when protection is provided for 
multiple batteries and energy is exchanged between battery 
subsystems . 

Bidirectional Operation. Bidirectional operation is useful 
when feeding regenerative loads (such as certain types of motor 
loads). This allows return of energy from the regenerative 
load to a battery or to other loads in the system. 

di/dt Limiting. The rate of rise of fault or overload current 
will be monitored without using inductor limiters to allow for 
high surge capability. This is desired in both the turn on or 
turn off modes to prevent excessive voltage and current stress 
on the power switching element. Controlled di/dt reduces EMI 
in the system. 

In addition to the above, which are generic in nature, the following 
important characteristics can also be incorporated: 

(1) Matching I^t characteristic to protect system cable. 


(7) 


( 8 ) 


(2) Current level monitoring. 

(3) Current leakage monitoring. 

Many of the above characteristics are also incorporated by Lockheed Advanced 
Marine Systems Division in the RBI as shown in Table 5 3. 


5-7 


5.2 


REMOTE POWER CONTROLLERS (RPCs) 


Remote power controllers are devices that functionally replace con- 
ventional electromechanical relays in the 1 to 30 A size. In addition to the 
function of supplying power to a load and interrupting power to the load, the 
functions of the RPCs are similar in characteristic to the solid state power 
interrupters discussed earlier with a difference in current ratings. The Solid 
State Power Interrupter Block Diagram (Figure 5-3) is similar to that for Remote 
Power Controllers (RPCs). A summary of some of the most significant 
characteristics for the RPCs follows [4-3 and 5-12]. 

(1) Controlled rate of rise and fall of current. 

(2) Short circuit protection. 

(3) Current limiting to prevent load transients. 

(4) Wide temperature range. 

(5) Electrical isolation of control and status signals from the 
power bus . 

(6) High-speed trip-out response characteristics. 

(7) Diagnostic capability. 

(8) Load capacitor charge capability. 

(9) Programmable trip characteristics. 

It should be noted that these characteristics are typical for a 
120 Vdc 5 A device (Table 5-4 Westinghouse ) , but many of these characteristics 
apply to other manufacturers and other current levels [5-13]. A trip 
characteristic is shown in Figure 5-5. This trip characteristic is used to 
coordinate cable, load, and RPC device protection. 

There are various manufacturers of Remote Power Controllers. The 
manufacturers and pertinent technical data relating to the RPCs are listed in 
Table 5—4. Detailed specifications are available from the manufacturers. Other 
manufacturers, other than those listed, indicated that they had active, in-house 
development programs for RPCs. 


5.3 POWER CONDITIONERS 

There are numerous suppliers of spacecraft dc/dc power conditioners. 
Most power conditioners for spacecraft usage are designed for input voltages 
within the range of 22 to 50 Vdc. Various power conditioner manufacturers have 
built power conditioners utilizing an HVdc bus for aerospace and terrestrial 
application. The design of off-line high efficiency power conditioners is 
common practice in the industry. The input voltage levels range between 100 to 
300 Vdc. The lessons learned and techniques in off-line (utility power) power 
conditioner design are applicable to HVdc spacecraft power conditioner design. 


5-8 



The design techniques for HVdc power conditioners are similar to the 28 Vdc designs 
used in spacecraft. For HVdc designs, higher voltage input components are require 
in the input side of the power converter. The design techniques used for hig 
voltage, high efficiency power conditioners, particularly in the contro o 
radiated and conducted EMI and the protection of the supply from voltage transients 
and faults, are similar to that for the design of lower voltage power 
conditioners. Examples of both high and low input voltage designs (270 Vdc and 28 
Vdc) are shown in Table 5-5 and Figures 5-6 and 5—7. 


Table 5-4. Remote Power Controllers (RPC) 


Manufacturer 

Current 

Rating 

A 

Features 

Volt- 

age 

Rating 

Vdc 

Size, in. 

Weight , 
oz 

Teledyne 
Solid State 
[5-14, 5-15] 

2,5,10, 

15,20, 

50 

Load switching , 
overload pro- 
section, remote 
control, built- 
in-test 

270 

Eurocard format 

NA 

Leach Relay 

2,5,10, 

Load switching, 

270 

1.7 x 1.6 x 0.4 

1.7 

1.5 

[5-16] 

2, 5, 7. 5 

overload pro- 
section, input- 
TTL compatibil- 
ity, current 
limiting 

150 

1.8 x 1.6 x 0.6 

Rockwell 

[5-17] 

1,2,5 

Overload pro- 
section, load 
switching , 
remote control, 
built-in-test 

270 

1.5 x 0.5 x 0.2 
Hybrid design 

NA 

Kilovac 
[5—18 and 

4 

Overload pro- 
section, load 

270 

3.25 x 1.5 x 2.2 

9 

5-19] 

100 

switching , 
remote con- 
trol, status 
indicated 


3.75 x 2.0 x 2.2 

14 

Westinghouse 
[5-6, 5-20, 

5 

See optional 
requirements 

120 

1 x 1 x 1.71 

3.3 

"7 O 

5-21, 5-22] 

30 


120 

3.24 x 3.64 x 1 

7.2 


1 

Load switching 
projection cur- 

300 

2 D x 1.1 

3.5 

n c. 


2 

rent limit, 
remote control, 
built-in-test 

300 

2 D x 1.1 

3 . b 


5-9 



Table 5-4. Power Controller (RPC) (Cont'd) 


Optional Requirements 

a. Steady state current rating of 5 A at 120 + 12 Vdc. 

b. Voltage drop across the RPC at rated current is to be less than 1.0 V. 

c. Inverse trip time delay, T, must satisfy (I 2 - G 2 )T = 20 A 2 seconds 
+5 percent for I greater than 6 A. 

d. The RPC is to limit the maximum load current to three times current (3x) up 

to 0.1 s before interruption. This 3x level is to be constant with respect 

to supply line (bus) voltage variations* 

e. Output rise time and fall time are to be between 10 to 10,000 ps. 

f. EMI generation and susceptibility are to be minimized. Control and power 

circuits are to be dielectrically isolated. 

g. Operating temperature: -55°C to 100°C. 

h. Power losses are to be minimized.* Off-state leakage current is to be less 
than 5 mA. The RPC is to provide remote status indication of the open or 
closed states. 

i. Weight is to be minimized. 

j. Failsafe I 2 t is to be 625 A 2 - seconds for a current greater than 5x. 


*Power loss in an RPC is important since system efficiency is affected. A voltage 
drop of less than 1 Vdc for a device rated at 120 Vdc has a power loss of less 
than 0.9 percent. This level is achievable. Where FET solid state devices are 
used, lower power losses are achievable. 


Power conditioner efficiency improves in high voltage designs since 
forward drop in semiconductors is a lower percentage of input voltage. In 
addition, higher output voltages result in more efficient converters for the 
same reason. 


Table 5—5 shows several commercially available low mass, high 
efficiency power supplies. Power density is up to 50 W/in 3 and efficiencies 
of up to 85 percent at 5 Vdc output. Figures 5—6 and 5—7 are photographs of 
power conditioners developed at TESLAco and Boeing Electronics Co. Figure 5-6 
shows a 270 Vdc, 100 W converter built by TESLAco for Hughes Aircraft. Figure 
5-7 shows a hybridized design of similar topology built by Boeing Electronics 
Co. for 28 Vdc input and 50 W output. 

At high power levels (50 kW), at an input voltage level of 100 Vdc, 
the efficiency is 95 percent (see Table 5-5, Space Power Inc.). 


5-10 



Figure 5-5. Trip Characteristics for RPC [5-20] 


Table 5-5. Power Conditioners (dc/dc Converters) 


Mfr 

Power 

Rating 

(W) 

Efficiency 

% 

■ 

Output 

(Vdc) 

Output 

Current 

(A) 

Q 

Power 

Density 

W/in^ 

TESLAco 

[4-4] 

100* 

85** 

300 

5 

20 

3 

30 

Inland 

Motor 

[5-23] 

100 

78-83 
depending 
on output 
voltage 

135 

270 

As req'd. 

5,12,15, 

28,48 

Depend- 
ing on 
power 
rating 

4.6 

50 

Space 

Power 

Inc. 

[5-24] 

50 kW 

95 

100 

>100 


167 

300 

* See Appendix C 

** This unit was built for Hughes Aircraft Co. which currently manufactures a 
line of power conditioners using this topology. 


"T- 
•1 ' t* 


m 


-c 




r-i 


5-11 











Figure 5-6. TESLAco 100 W, 270 V to 5 V Switching Power Supply 
(3 in. x 2 in.) 



Figure 5-7. Boeing Electronics Company 50 W, 28 V 
to 5 V dc/dc Converter Hybridized 


5-12 


ORIGINAL PAGE IS 

OE POOR QUALITY 




The values of efficiency for the power 

Appendix C and Table 5-5) are confirmed W out^ voltage o? a converter 

increase^'ef f ic iency iufimpro^due to" semiconductor losses becoming a smaller 
percentage of total power. 

A review of the state of the art of high frequency switching power 
conditioners is provided in Appendix 0. Appendix 

placed on high P^t ^ensity. hrgh rffic * contigur ations. The cost of 

performance parameters, as wei a F , Th-iq review of the state- 

the high power density converters rs that very 

of-the-art technology in the swrtchrng dc/dc sent day components, 

high power density P»»er converters can^be ^ effi * lenclell of 85 to 

These converters approach 50 / » P . fanoroaching 450 W/lb). Increased 

90 percent, and result in ^ghtwergh untts ^ 2 kw t0 20 

rie:d^o f rS C Ln:™t l of cliderabl, improving the bandwidth to 20 XHs, as 
well as step-load transient response settUng time to less ^ 10 ° P d ; lacenent 
addition, the use of switching frequences beyond ™ “* “ sul J s in 

Of electrolytic and tantalum capacitors by film capacitors. This 
higher levels of power conditioner reliability. 

tj ^ dc BUS REGULATOR 

A dc bus regulator can be used to regulate the voltage of an 
unregulated dc bus or to provide ad jus ^section” 5- 3 ' in'* that electrical 
ti f nc 6 used” ‘^inHwi tc"h-^de processing techniques, regulation and 

iransformation of dd ^Uchin/deiices lul as^ th^powei^MOSFET 

enable* the us^of^high switching frequencies - which in turn allows for reduced 
magnetic and filter masses. 

A dc bus regulator rated at 20 kW has been developed by s P^ e Power 
me. to power an electric ^^^to^develope^for Aeroenvironment 

l^Vafu^ed as s, array' shunt regulator for a photovoltaic-propelled electrrc 
vehicle (GM-SUNRACER) [5-25]. 

The power levels shown in Table 5-6 can be scaled to satisfy most 
Space Station requirements. These include P d » et r d ““°“ e " a o°" and'other loads 

shown in Figure 5-8. 


5.5 


ac MOTOR DRIVES 


Various types of ac motor drives have been developed for aerospace 
.. V % „ fil y ? hes e drives generally include an inverter operating with 
applications [5-26]. These drives g^ ^ ^ design consid eration in 

increasing" the" input dc voltage level to the HVdc requirements is similar to 


5-13 



MOSFET 


FILTER 


7 ^ 



VDC | 

INPUT 

-> 


n 

L 


i 

i 

1 

1 

r j 

i : 

• >■ + 

DC 

OUTPUT 


IxjuJ 

■ 4 

► - 


nnm 


PULSE WIOTH 
MODULATION 
CONTROL 


OUTPUT 

VOLTAGE 

SENSE 


Figure 58.* dc Bus Regulator (buck type) 


that discussed in the previous section on power conditioners. This class of 

sinreth 8 T ^ electrical isolation between the dc bus and motor lofd 
since the motor provides the electrical isolation desired for safety reasons.’ 

Th* ► A . cla f s °f inverters has been developed for controllers/actuators 

° rS/C0ntrOllerS are basicall y dc brushless type systems. They are 

terrestrial'an^aerosD With °^ . ele ? trical isolation, £ny ac motor c^ntroL in 
terrestrial and aerospace applications do not require electrical isolation f nr 

performance or safety considerations. The net result of eHmination of the 
1 - 121 °™' (eUCtriCal iSOlati °" ) ** “ i-Ptovement in efficie^ and redaction 


.. . . A aew class of solid state dc/ac inverters using high freauencv 
switching and PWM control has been developed that is applicfble^o the space 
environment as shown m Table 5-7. This table shows what can be achieved usine 
advanced semiconductor devices and advanced technioues in pum caie y ed using 
utilize digital control of voltage and c^re^ u^L M, 

effic? 1 '” “* * h'nation of high voltage integrated circuits and high 
efficiency power switching devices [1-7], Advance packaging techniques reduce 
! S .“ d PrOVide ef£ectiv e transfer. Most notable aspecJof 

low L” 68 "" the hi8h to 97.5 percent) and tLir 


o“ ^9r^ P u\L i ^1 r tra V L r fo 8 ™rL™!‘:L P e r ff"L d n“ <i 

confirmed XXTJl $ £ Ku* ^ 


5-14 




Table 5-6. Characteristics of dc Bus Regulators 


Mfr. 

Power 

Rating 

Input 

Voltage 

(Vdc) 

W/in. 3 

Efficiency, 

% 

Weight , 
lb 

Aeroenviron- 
ment Inc . 
[5-25] 

150 W 

140-250 

20 

98.5 

0.6 

Space Power 
Inc. [5-28] 

20 kW 

120-150 

120 

96.0 

5.0 


Table 5-7. ac Motor Drives 


Manu- 

facturer 

Rating 

kW 

Input 

Voltage 

(Vdc) 

Switching 

Element 

Maximum 
Output 
Freq'y, Hz 

Effi- 

ciency, 

% 

Size, in. 

Weight, 

lb 

Aer ©en- 
vironment 
Inc. 
[5-25] 

5 

120 

FET 

500 

97.5 

14 x 16 x 2 

8 

General 

Electric 

[5-29] 

12 

270 

IGBT 

400 

92 

2x6x8 

8 

JPL 

35 

240 

Bipolar 

400 

97 

16 x 12 x 6 

42 

[5-30] 


. 

Transistor 






The above three designs use a six-element bridge type of inverter. 

A schematic of the inverter of the General Electric design is shown in 
FigurH-9. It should be noted that the six switching elements perform both the 
waveform synthesis function and the voltage control function. 

The switching devices used differ according to the design. All of 
the above designs can be modified to utilize the new “^-controlled thy r^s or 

(MCT) that is currently under development at Genera expected that the 

initial production units soon to be available [5 31J. . P 

usage of the MCT will provide improved inverter characteristics. 


5-15 





Figure 5-10 is a block diagram of the General Electric motor 
controller used to drive an electrostatic hydraulic or electromechanical 
actuator [5-29] . It should be noted that a high level of electronics 
integration of drive and control devices is utilized. The control strategy is 
implemented by using high voltage integrated circuits (HVIC) logic and digital 
signal processing systems. These controllers can be constructed using various 
power semiconductor devices such as IGTs, MCTs and MOSFET devices. The motors 

used can be of the permanent magnet ac type, switched reluctance, or brushless 
dc type. 


^* nce tlie inverter has six switching elements, inverter reliability 
is high (due to low parts count). Tens of thousands of these inverters have 
been built for commercial, aerospace, and military applications. Resonant 
converters that are bilateral have a very high power semiconductor parts count 
®^ ora P aratlvel y complex logic circuitry. Therefore, a comparative study of 
the PWM vs the resonant synthesized waveform inverter should be made to 
determine the optimal design choice. 


Typical motor control/inverter characteristics are as follows: 
o Input dc voltage: up to 300 Vdc operation 

o Power rating: up to 50 kVA 

o Voltage and frequency is regulated 
o Microprocessor control 
o Short circuit proof 


5-16 











VMS 


I 


1 

w i 

i 



i 

1 



MVIC 


MCI 


* — 

Htgh-Sp»«d 

E 


i 


fUtl Tln»« 


MOSrf T3 



DS* 






Switched 

fUlucUnc* 


AVflrtm* 

Function 


Figure 5-10. Integrated Actuator System [5 29] 


o 

o 

o 

o 

o 

o 

o 

o 


Current limiting 
Bilateral operation 

Wide range of leading and lagging power factors 
Constant torque, constant power operation capability 

Reversing 
Dynamic braking 

Interface to spacecraft data bus 
Soft start 


5.6 SENSORS 

Two types of sensors offer provisions 

fault°sensors^ St The^ types '"and" characteristics of various sensors are given held.. 
Current Sensors 

A means of measuring dc bus current that^is ^^trically 

from the main dc bus is desirable for sys em ,^8. The Micro Switch and 

[5-32]. Several available ^peslnd can be used in any 

the Liasons Electroniques sensor current. Their frequency response is 

part of a spacecraft for monitoring dc bus current. 

in the low microsecond region. 


5-17 










Table 5-8. dc Isolated Current Sensors 


Manufacturers 

Current Rating, 
A 

Current Rating 
Size, in. 

Weight, 

oz 

Liaisons Electroniques 
[5-34] 

1-1000* 

1.5 x 0.625 
x 0.75 /50 A 

1 

Micro Switch [5-35] 

1-1000* 

2.26 x 0.5 
x 2.26 /950 A 

1 1.3 

JPL [5-36] 

800 

1 x 0.5 

x 0.652 /800 A 

0.75 

♦Various sizes in this range 




nnwor/ .. . A "° th er type of sensor developed at JPL is integratable in a smart 
power/high voltage integrated module and can be utilized as part of RBI RPC or 
power supply assemblies. P 1 * RPC ’ or 


5.6.2 


Ground Fault Sensors 


round fault sensors for dc systems have 

d^ewi a for S r2™ S ™ nd »° th ? r . mil i tar r * 8™nd fault sensor 

• P p f f, 27 ? Vd Naval Alr Development Center (NADC) aircraft to provide 

The size of £h * tS i manufactured b y Rockwell International Inc. [5-33?. 
The size of the sensor is 2 in. x 4 in. x 4 in., and it weighs 8 oz. Work 

performed for terrestrial photovoltaic systems array applications included a dc 
ground fault detector required for safety purposes [5-37]. 


nave uceu usea 


x a. x 


5-18 




SECTION 6 


FLAT CABLE IN ELECTRICAL POWER DISTRIBUTION SYSTEMS 


Flat cable has been considered by the Naval *it D evelop™nt c enter 
(NADC) for inclusion in aircraft with 270 Vdc operation [6-1, 6-2]. The results 
of a study conducted for NADC indicated the following: 

(1) While additional investigation is required* initial 
investigations indicated that cable shield and components 
available to meet aerospace quality environments. 

(2) Flat conductor cable will tolerate even higher current densities 
than predicted in previous studies. 

(3) Most environmental factors do not have a significant advers 
impact on the thermal performance of flat conductor cable. 
Consideration should, however, be given to current derating fo 
cable shielding, cable folds, and altitude. 

(4) Flat cable harnesses can be manufactured following most standard 
processes, and in a timely manner* 

(5) No significant problems are encountered in installation of a 
complete flat cable harness. 

The aPPli-ion^f Useable * 

e^e cable temperature compared to round cable, and thereby reduce cable maea. 

The choice is dependent on system cost and mass modeling. 

Flat bus cable does not exhibit any EMI characteristics that would 

S^&n^lSTtS cable 1 to -t effect of 

this is to reduce interaction between the cable and other components in 

due to the shunting action of the cable capacitance. 

It should be noted that the capacitance of the cable actually stores 
energy that can be used to feed loads that require high rates of current r • 

In an ac system, the capacitance of the cable draws reactive , nd 

inverter source, resulting in additional power being dissipated 


6-1 



power conversion components. In addition, the inverter and its output 
transformer must be increased in kVA capacity to handle the reactive 
requirements of the bus capacitance. 


Connectors for flat bus cable are being developed for 
usage. Manufacturers such as Gore Inc. and G & H Technology Inc. 
connectors for flat cable. 


Space Station 
produce 


6-2 



SECTION 7 


RELIABILITY 


Since the high voltage dc spacecraft secondary distribution bus has 
similarity with its low voltage counterparts and HVdc aircraft [5-33], it can be 
expected that using the same ground rules of design, the reliability will be 
equal or better. Improved reliability for HVdc systems can be expected 
(compared to present generation spacecraft) due to reduced parts count and 
elimination of electromechanical switches [5-33]. In addition, reliability will 
also be improved by incorporation of high voltage power integrated circuits and 
advanced semiconductors that have been qualified for space application 
(including single event upset phenomena). 

The reliability of a power conversion subsystem is dependent on the 
following elements of design, 

(1) Thermal stress. 

(2) Electrical stress. 

(3) Parts count (complexity) . 

(4) Selection of parts (class of part). 

Improved reliability is dependent upon thermal stress. The 
operating temperature of a power conversion device is dependent on heat 
dissipation, mass, radiating area and cooling methods. Improved reliability 
generally relates to increased size and mass. Establishing an optimum 
relationship between reliability and size and mass is the design goal for a 
power conversion subsystem. 

Because of the limited production of HVdc components, mean time 
between failure (MTBF ) data is limited in availability [7-1]. Available data 
for HVdc components is shown in Table 7—1. 


Table 7-1. Component Reliability 


HVdc Components 

Reliability Characteristics 

Solid State Switches 
Hybrid Switches 
Power Conditioners 

10 5 -3 x 10 5 Switching Cycles 
10^-5 x 10^ Switching Cycles 
10 5 MTBF (h) 


7-1 





SECTION 8 


EMI CONSIDERATIONS 


Low voltage spacecraft have been built, tested, and flown with power 
ratings up to 4 kW. These spacecraft have been designed and tested to specified 
EMI specifications. Many spacecraft have been qualified to MIL-STD-461B . 

The usage of solid state or hybrid switches which can limit rate of 
rise and fall times and eliminate arcing will reduce EMI for dc systems since EMI 
characteristics are partially driven by switching and transients [8-1] . 

A key element is designing for low conducted current ripple in order 
to meet MIL-STD-461B. MIL-STD-461B specifies allowable current levels over a 
frequency band. It can be inferred that the conducted ripple current level of a 
power condition operating at 28 Vdc and 1 kW, will be similar to a 4.5 kW power 
conditioner at 126 Vdc. This is because the bus current and the power switches 
in the power conditioner will process similar current levels. Ripple voltages 
on the spacecraft bus will be higher because the bus source impedance will rise 
at approximately the same ratio as the voltage. Ripple voltage as a percentage 
of dc bus voltage will remain the same. If additional ripple voltage reduction 
is required additional filtering can be added. With the usage of flat dc bus 
cable which represents itself as a transmission line, additional conducted EMI 
suppression is provided due to the cable distributed capacitance [6-3] . It 
should be pointed out that the EMI filter component values for the low and high 
voltage spacecraft will be essentially the same, except for the higher EMI 
capacitor voltage ratings in the high voltage design. The issue of whether the 
radiated noise from the power conditioner will meet MIL-STD-461B is not known at 
this time, but the techniques of suppression should be similar to that used in 
the lower voltage designs [1-4] . The resolution of EMI issues for a high 
voltage dc bus distribution is well known compared to the resolution of EMI 
issues of a 20 kHz power distribution/component system. 

With reference to the issue of radiated fields in an HVdc 
distribution system, it should be noted that with the absence of an inherent 
frequency in a dc system and with an inherently lower source impedance it can be 
expected that the dc system will have low levels of radiated electric and 
magnetic fields. This results from the fact that the distribution system will 
have negligible medium and high frequency signals. Radiated fields in the 
components such as power conditioners will be addressed with the same shielding 
techniques used in low voltage dc designs. 

Electric and magnetic field radiated energy is a problem requiring 
resolution in an ac distribution system. 


8-1 



SECTION 9 


SAFETY CONSIDERATIONS IN HIGH VOLTAGE dc SYSTEMS 


Principal safety hazards are fire, shock, and audible noise. Dc 
systems are typically immune from audible noise due to designs above the audible 
range . 


Fire hazard is common to all distribution systems, irrespective of 
the distribution frequency. The RBIs and RPCs detailed in this report can 
quickly clear major faults and thereby reduce fire hazards. 

Shock level is related to current passing through the body. Current 
is dependent on voltage level, body resistance, and frequency [9-1]. Secondary 
distribution voltage level of 120 Vdc is sufficiently low and should not present 
a major hazard to personnel, when used in conjunction with ground fault sensors. 

From a safety point of view, voltage and frequency do have an effect 
on perception and paralysis of humans. Sensitivity to touching a dc bus versus 
touching an ac bus varies with frequency. The worst sensitivity effects occur 
in the band from 60 Hz to 400 Hz. At similar voltage levels, the sensitivity to 
dc is approximately equivalent to that at 10 kHz. This is shown in Table 9—1, 
"Effect of Electric Currents on Cells and Tissues" [3-3]. No unusual safety 
problems at the 120 Vdc level were reported in the literature. 

It should be noted that a 120 Vdc bus and a 20 kHz 208 Vac bus 
appear to be comparable with regard to shock hazard. 


Table 9-1. Effect of Electric Currents on Cells and Tissue 


Effect on Frequency (Hz) 
(fraction of 

on Perception and Paralysis 
60 Hz effect) 

f (Hz ) 

Effect 

0 

0.2 

10 

0.9 

60 

1.0 

300 

0.8 

1000 

0.6 

10,000 

0.2 

RF 

0.01 


9-1 





SECTION 10 


POWER MANAGEMENT AND DISTRIBUTION (PMAD) CONTROL SYSTEM 
(Contributed by Linda Palmier!) 

PMAD control systems are well {3* b°uT 

[10-11. "t*? 1 ! 64 *n cl 4-31 t: ‘“pMAD^ontrol systems are also well developed for 

protocols [10-2, 2 12, • soacecraft PMAD system includes control of 

LVdc spacecraft. A modern comp P batteries, controllers, 

diverse elements such as power source L^L ^d controllers consisting of 
converters, and a complex networtof re^load^con ^ ^ 

hundreds of solid state relays t cur rent limiting, redundant bus 

detection, fault isolatio , . _ sources, storage and conversion 

management, management and control o P system status. PMAD architecture 

subsystems, data management, communic » RVdc systems, and the best 

- ZZZZZXlZZ ^uTaH^eY a -Bern PMAD system. 


I LOW VOLTAGE dc SPACECRAFT 

. . ... wstA - 

radioisotope thermoelectric genera ^ power from these sources is 

provided by two identical secondary batteries. ^ (typically 28 to 

distributed via branches of a P^^ bus^ circuit breakers) are used for 
30 Vdc) . Solid-state switches power bus. These switches 

connecting and disconnect ng sp telemetry data handling and monitoring 

enhance command capabilities and increase ceie 

requirements • 


10.2 POWER GENERATION AND STORAGE 

. M!S V of the RTG is approximately 310 W. The RTG 
Maximum power capabiiity “ bus th „ u gh an isolation diode. 

is electrically connected to the 30 JL- tored v ia telemetry conditioners in 

The RTG electrical characteristics are y div Led into several 

the Power Control Assembly. The solar a y^ connected in a series- 

segments. Each segment consis J rearranging their configuration to 

parallel configuration that facilita ^ ( S ASU) , g capable of connec ting 

control voltage. A solar array g j. response to a 

or disconnecting any number of segments to or from tne 
control signal generated by the shunt regulator. 

Batteries, within their limitations, provide 
those spacecraft activities where the P°£®^ nower. A bi-directional converter 

exceed the available ch arging°of the battery in response to the control 

(BDC) handles charging and discharging excess RTG/ Solar Array power is 

signal generated by the 8 b«nt regulator. bmttwry by providing 

available, the BDC is designed .to begi 8 ^ the RTG /Solar Array power is 

insufficien^t^keep 1 ^^ 1 power bus Voltage regulated, the BDC discharges the 


10-1 



10-2 


("power 


SOLAR 

ARRAY 

[string 71 


CONTROL 


PPS 

processor 


SOLAR 

ARRAY 

SWITCHING 

UNIT 


l ^reconfiguration"] I 

. ELECT I"*-. 


•U-L 

:L-Ff 



CDS 

I 

I j SOLIO STATE 
T I RELAYS T 


COMMANO/DATA 

MULTIPLEXING 


—I LlOWER_DISTRI Bii TIONJ 


I 

' I ^ 1 

I MAIN 
H — ERROR - 
| AMP 

J SHUNT 
! REGULATOR 


POWER 

SEQUENCE 

STAGES 

4 


FAULT PROTECTION 


COMMANOS 


TELEMETRY 


|_SH UNTRADJATORJ 


Figure 10-1. Power Processor Subsystem (PPS) Functional Block Diagram 
















battery at an appropriate rate to augment the RTG/Solar Array power and maintain 
an acceptable bus voltage. A transient current source (TCS), within the BDC, 
contains low-impedance, fast response circuitry through which the battery can be 
connected directly to the 30 Vdc bus in the event of a transient bus overload. 

A reduction in control voltage causes the BDC to respond using the TCS. 


10.3 


POWER REGULATION AND DISTRIBUTION 


The shunt regulator unit provides regulation of the 30 Vdc bus by . 
shunting excess RTG/Solar Array power from the 30 Vdc bus. This excess P ow ® r J-® 
dissipated partly in power semiconductors within the shunt regulator package and 
partly in resistances external to the electronics equipment bays of the 
spacecraft. The shunt regulator develops a control voltage (CV) which is 
approximately proportional to the 30 Vdc power bus voltage. The CV is used to 
control the BDC, the transient current source, the SASU, and the shun regu 
currents. The shunt regulator is responsible for power regulation; however, the 
microprocessor is used for fault detection and response. That 1 % co ^ r ° 1 ° f 
battery connection/disconnection to the power bus is hardwired controlled by the 
shunt regulator through the BDC. However, battery fault monitoring, detectio , 
and response are managed by the microprocessor using software. 

The Power Distribution Assembly (PDA) is an array of solid-state 
switches that are used to connect engineering/science loads (Figure 10-2). Each 
switch connects or disconnects an electrical load to the 30 Vdc us in response 
to signals received from the microprocessor. The present solid state swi 
configuration requires that both the power bus side (HIGH) and the return side 
(LOW) be simultaneously switched. The switching function must be immune to 
single-point failures that would result in a load being permanently on; 
therefore the switches are series redundant (HIGH and LOW sides). The swi c 
incorporates a '•soft" turn-on feature, load current limiting capability, and 
functions as a programmable circuit breaker. A series redundant current limiter 
and circuit breaker, with a fixed overload trip magnitude, is incorporated as 
part of the switch design. Each user load is supplied with parallel redundant 
switches within the PDA in order to prevent a single switch failure from 
resulting in the loss of connecting or disconnecting that load to the 30 Vdc 
bus. Specifically, there are two independent switches in the PDA that can be 
used to supply power to a single load. The internal microprocessor sends 
commands that are decoded in I/O board circuitry, to drive the solid-state 
switch to connect, disconnect, or reset the current limit value of the 
associated switch. 


10.4 MICROPROCESSOR FUNCTIONS 

A redundant microprocessor manages on-board command processing, 
telemetry collection, and fault detection and response. Command processing 
includes the receipt of a ground command and the forwarding of this comman o 
the appropriate power system component. Examples of power commands include 
battery reconditioning, solid-state switch commands, and solar array 
reconfiguration commands. Telemetry collection involves sampling of all 
engineering data that are obtained/contained within the Power Processor 
Subsystem (PPS) and compressing these data to meet downlink transmission 


10-3 



STATUS 



Figure 10-2. Hybrid Switch Block Diagram 

requirements. PPS data are generated by many transducers located throughout the 
system. Analog sensors are used to measure voltages, temperatures, and 
currents. Digital sensors are used to monitor any PPS specific events. These 
data are used by the microprocessor to determine the state of health of the 
power system and to detect the occurrence of any faults. If a fault condition 
exists, the microprocessor initiates an appropriate fault response action, and 
reports the fault occurrence and its associated response to ground operations. 


The combined use of the solid-state switches and the microprocessor 
results in greater flexibility of the power system. The microprocessor handles 
switch commanding, specifically turn-on and turn-off, and setting a current 
"trip” value. Within the switch electronics, this current value is compared to 
the load current and results in a switch turn-off if the load current exceeds 
the trip value. Telemetry from the switches is used by the microprocessor to 
detect switch related faults and proceed with an appropriate fault recovery 
routine. One example of a fault response is attempting to determine whether a 
switch turned off due to a failed switch, or an excess load current condition, 
or a load failure. 

The microprocessor can be used to monitor battery telemetry, 
including temperature, voltage, and current. The microprocessor monitors these 
data for faults and calculates battery state-of-charge (SOC). SOC is critical 
for ground analysis purposes to determine when battery reconditioning is 
appropriate. The microprocessor has the on-board capability to change the 
voltage-temperature level in the event of certain battery faults. The 
microprocessor also interfaces with the bi-directional converter to command 
various VT levels for battery operations. The microprocessor is interfaced to 


10-4 








the solar array through reconfiguration electronic, to provide 
capability to reconfigure the electrical configuration of the solar array 

segments . 

Table 10-1 lists typical software nodules that ««beincludedina 

microprocessor-based power system. Estimates of the site, in £“* °L£ t e°“re 
rode are also listed. The total number of lines is 1420. These estima 
generated with knowledge gained from prototyping, in a development ’ 

the solid-state switch command functions. Command and telemetry collection 
modules in the power subsystem are less complex than modules oi JiWresp^se 1011 
in the command/data or guidance /control subsystems. Fault det ® C ^°"^ e ^° 
modules are more complex due to the relationship of the subsystems and the 
possibility of one subsystem’s faults affecting another subsystem. An attemp 
to limit propagation effects of faults on other subsystems results in greater 
complexity of the subsystem fault protection software. 

As more on-board capability is added to the power subsystem, the software size 
requirements will increase. On-board enhancements include battery ^Hncrease 
and reconditioning, trend analysis, and power margin management The ^crease 
of software modules not only accommodates for the added functionality 
provides for more complex fault. 


10.5 CONCLUSIONS 

The trend for unmanned planetary spacecraft is to connect/disconnect 
user loads through solid-state switches and to expand the on-board capability o 

th^microprocessor to include functions that have been ^ro^d bl fe;/ 
verified on the ground. Such additional functions may be on-board battery 
reconditioning, Lend and performance analysis of the power sources, and 
expanding the fault detection and response capability. 

The control and fault protection schemes for manned spacecraft 
applications are similar to unmanned LVdc power systems in approach however 
more complex in application. Manned spacecraft systems, in addition to power 
regulationand distribution, must provide greater fault protection through^ 
increased redundancy and provisions for emergency power. Thes q mor ,r 

support theprimary^manned spacecraft goal to provide an operating environment 

that supports human life. 


10-5 



Table 10-1. Power/Pyro Subsystem - Software Code Estimates 


Function 


Lines of Code 


COMMAND MODULES 


Initialization (set-up of Solid-State Switches) 
Command Validation (verify prior to execution) 

Command Expansion 
Solid-State Switch Commands 
ON/OFF 

Set CURRENT LIMIT VALUE 
Battery Reconditioning 
Initiate Pyro Events 

TELEMTRY COLLECTION MODULES 

Sample Solid-State Switch Analog Data 

Sample Solid-State Switch Digital Data 

Sample Battery Telemetry 

Sample Bi-Directional Converter Data 

Sample DC Bus Unbalance 

Sample RTG Data 

Packetize Telemetry Data 

FAULT PROTECTION MODULES 

Microprocessor Self-Diagnostics (health check, status) 
Battery Monitoring (State-of-Charge) 

Solid-State Switch Monitoring 
I/O Board Diagnostics 
Solid-State Switch Trip Response 
Solid-State Switch Failure Response 
Battery Over/Under Current Response 
Battery Over/Under Voltage Response 
Battery Over/Under Temperature Response 
Bi-Directional Converter Failure 
I/O Board Failure 
DC Bus Overload 


s 

s 

s 

s 

s 

m 

s 


s 

s 

m 

s 

s 

s 

m 


s 

m 

s 

s 

s 

s 

s 

s 

s 

s 

s 


SYSTEM FAULT PROTECTION INTERACTION 

Microprocessor OPERATING SYSTEM Implementation Modules 
Interrupt Service Routines 
I/O Driver Routines 
Other 


*1 100 lines of C source code 

m 60 lines of C source code 

s 30 lines of C source code 


10-6 




SECTION 11 


CONCLUSIONS 


Current technology will allow the development of an HVdc secondary 
distribution system for space application. 


A laree number of vendors have the technology and space qualification 
background to produce the required components and design of HVdc 
competitively. RCA has built and qualified a spacecraft with a 100 Vdc bu . 

All components to build HVdc systems have been prototyped or developed for 
terrestrial and aircraft applications. Previous deficiencies^ dc 
switching have been resolved. These HVdc components developed for aircraft 
applications are compact, low mass, high efficiency, and suitable for space 
environment development. 


Qualification of most HVdc components for space application is still 
required. 

Because of the use of HVdc for array and battery subsystems, a number of 
HVdc components have already been developed for space applications, 
including brassboards for Space Station. 


There exists a substantial 28 Vdc to 50 Vdc spacecraft experience m the 
design of power management and distribution (PMAD) control systems. The 
prototype developments using MIL-STD-1553B control bus for aerospace 
applications have paved the way for HVdc PMAD systems for space station 

applications . 

The extensive space experience in low voltage dc design heritage allows the 
development of procedures and components for EMI control, system stabil y, 
and grounding for HVdc systems. 


Safety issues relating to HVdc are well understood. No unusual safety 
problems were reported in the literature. 

Users are familiar with design techniques and applications of dc/dc con 
verters and dc power controls which simplify user interface. 


The simplicity, low parts count, and previous space qualification history 
of dc systems offer the potential of high reliability for space 
environments . 


11-1 



SECTION 12 


REFERENCES 


1 - 1 . 


1 - 2 . 


1-3. 


1-4. 


1-5. 


1 - 6 . 


1-7. 


1 - 8 . 


2 - 1 . 


2 - 2 . 


2-3. 


2-4. 


2-5. 


2 - 6 . 


SPACE STATION PROGRAM DEFINITION REQUIREMENTS DOCUMENT (PDRD) , 
SSP30000, Sect. 3, Rev. F, May 6, 1988 

SPACE STATION ELECTRIC POWER SYSTEM REQUIREMENTS AND DESIGN, 

Fred Teren, 1987, IECEC Paper No. 879003 

SPACE STATION INTERNAL COMMUNICATION FROM ESA, June 1986, 

D.M. O’Sullivan, J.E. Haines 

JEM POWER SYSTEMS STUDY SPACE STATION INTERNAL COMMUNICATION, From 
National Space Development Agency of Japan (NASDA) , April 1988 

LETTER REPORT (Draft) PANEL ON ELECTRIC POWER SYSTEM OF AD HOC 
COMMITTEE ON SPACE STATION ENGINEERING AND TECHNOLOGY DEVELOPMENT, 
AERONAUTICS AND ENGINEERING BOARD, NATIONAL RESEARCH COUNCIL, 

June 1986 

ANALYSIS OF THE IMPACT OF A 270 VDC POWER SOURCE ON THE AVIONIC POWER 
SUPPLIES IN THE S-3A AIRCRAFT, Naval Air Development Center Contract 
No. N62269-78-C-0007, Lockheed-Calif . Co., LR 28780, 

November 27, 1978 

SMART POWER/POWER INTEGRATED CIRCUITS, Proceedings of Workshop Held 
May 1987 at California Institute of Technology, "Smart Power 
Applications in Power Actuation" by Dr. M Paparo 

SMART POWER CHANGING THE FACE OF POWER CONTROL ELECTRONICS, 

April 28, 1988 

POWER SYSTEM DESCRIPTION DOCUMENT (PSDD), (Station and Platform 
Systems) September 11, 1987 

STC-DBS ELECTRICAL POWER SUBSYSTEM, S. R. Peck, T. Callen, P. 

Pierce, T. Wylie - RCA Astro-Electronics, Proceedings, 19th IECEC, 
1984, Volume 1, pages 428-441 

BATTERY INTERFACE DESIGN AND SIZING, D. Rusta, TRW, Proceeding of 
Annual Battery Conference on Applications and Advances, California 
State Univ., Dept, of EE, Long Beach, January 1986 

TECHNICAL SPECIFICATION FOR THE SP-100 SPACE REACTOR POWER SYSTEM, 
Revision 7, September 18, 1987, by the SP— 100 Program Office 

SAE-AE7 PROCEEDING, April 20, 1988 

ANALYSIS OF THE IMPACT OF A 270 VOLT dc POWER SOURCE ON THE AVIONICS 
POWER SUPPLIES IN THE P-3 AIRCRAFT AND THE EFFECT ON AIRCRAFT 
PERFORMANCE, Lockheed-Calif ornia Co. Report No. NADC -81 156-60, 
February 28, 1983 


12-1 



2-7. 

2 - 8 . 

2-9. 

2 - 10 . 

2 - 11 . 

2 - 12 . 

2-13. 

2-14. 

2-15. 

2-16. 

2- 17. 

3- 1. 
3-2. 

3- 3. 

4- 1. 
4-2. 


270 VDC AIRCRAFT ELECTRICAL SYSTEM CONFIGURATION /DESIGN STUDY, Naval 
Air Development Center Report, Grumman Aircraft Systems Division, 
Interim Report 3EG-REEES-R— 07, November 1987 

SAE — AEROSPACE STANDARD No. AS 1931, "Electrical Power, High 
Voltage Direct Current, Aircraft, Characteristics and Utilization of" 

INTEGRATED DIGITAL/ELECTRIC AIRCRAFT CONCEPTS STUDY, M. Cronin, 
et al., Lockheed-Calif . Co., NASA Contractor Report No. 3841, 

January 1985 

SYSTEMS STUDY FOR AN INTEGRATED DIGITAL/ELECTRIC AIRCRAFT (IDEA), 
Boeing Commercial Airline Co. NASA Contractor Report No. 3840 - 
G.E. Tagge, L.A. Irish, and A.R. Bailey, January 1985 

MEETING THE ELECTRIC POWER NEEDS FOR AIRCRAFT DIGITAL ELECTRONICS, 

I. Mehdi, Boeing Military Airplane Co., NAECON 1984 

DESIGNING A FAULT TOLERANT ELECTRICAL POWER SYSTEM, M. Dige, 

P. Leong, D. Summer, NAECON, 1987, IEEE No. CH2450-5/87/0000-1483 

ADVANCED ELECTRICAL POWER SYSTEMS TECHNOLOGY FOR THE ALL ELECTRIC 
AIRCRAFT, R. Finke, G. Sundberg, NAECON 1983 

Personal communication with J. Fischer of NAVSEA and S. Krauthamer, 
JPL, March 28, 1988 

Personal Communication, S. Krauthamer, JPL and P. Mizele of Newport 
News Shipbuilding Co., March 28, 1988 

MILITARY STANDARD INTERFACE STANDARD FOR SHIPBOARD SYSTEMS, 

DOD-STD 1399 (NAVY) Section 390, Department of Defense, USA, 

November 20, 1986 

Personal Communication, S. Krauthamer, JPL and D. Farmer of General 
Dynamics, Land Systems Division, May 1988 

SECONDARY POWER DISTRIBUTION FOR SS, DOCUMENT ST-E-1000, National 
Space Development Agency of Japan, March 5, 1986 

NASA-CR-1205 [1], Compendium of Human Responses to the Aerospace 
Environment, Vol. 1, Sec. 5 

SPACE STATION GROUNDING STANDARD - JSC No. 30240, Space Station 
Program Office, July 2, 1986 

CONCEPTUAL DESIGN OF AN ADVANCED AIRCRAFT ELECTRICAL SYSTEM, 

W. Owens, et al., NAECON 1987, IEEE No. CH 2450/87/0000-0441 

Personal Communication - D. Pelhank of Douglas Aircraft Co. and 
S. Krauthamer of JPL, Re: Westinghouse Report "Smart Power 

Controllers for 270 Vdc Electric Power System," written for Douglas 
Aircraft Co., March 24, 1988 


12-2 



4-3. 

4-4. 

4- 5. 

5- 1. 

5-2. 

5-3. 

5-4. 

5-5. 

5-6. 

5-7. 

5-8. 

5-9. 

5-10. 

5-11. 

5 - 12 . 

5-13. 

5-14. 


MILITARY SPECIFICATION MIL-P-81653C, PRELIMINARY, "General 
Specification for Power Controller," July 1986 

LOW PROFILE, 50, W/in 3 , 500 kHz INTEGRATED MAGNETICS PWM CUK 
CONVERTER, S. Cuk, California Institute of Technology; Z. Zhang, 
Lateef A. Kajouke, TESLAco-High Frequency Power Conversion 
Conference, May 1988 

MILITARY SPACE POWER SYSTEMS TECHNOLOGY TRENDS AND ISSUES - 

R. R. Barthelemy and L.D. Massie, IECEC 1985, Paper No. 859035 

BACKGROUND AND EXPERIENCE IN THE DEVELOPMENT AND APPLICATION OF 
RPCs - REMOTE POWER CONTROLLERS, Westinghouse Electric Corporation 

AN INTRODUCTION TO HIGH VOLTAGE dc CONTACTORS, Hartman Electrical Co. 

Written Communication - From P. Theisen, Eaton Corp., Cutler Hammer 
Division to S. Krauthamer, JPL, Re; Contactors, May 1988 

ADVANCED LINE CONTACTOR FINAL REPORT, ID No. 6Y128, C.B. Hassan, 
Lockheed Advanced Marine Systems, San Diego, June 3, 1986 

270-V dc HYBRID SWITCH, P. Theisen, S. Krstic, Eaton Corporation, 

IEEE Transactions, Vol. CHMT-9, No. 1, March 1986 

HIGH VOLTAGE dc SWITCHGEAR DEVELOPMENT FOR MULTI-KW SPACE POWER 
SYSTEMS WAED, Report 81-05E, NASA-CR-1654137, Westinghouse 
Electrical Corporation Electrical Systems Division, November 1981 

Written Communication — from R. Johnson, Teledyne Kinetics to 

S. Krauthamer, JPL, Re: Hybrid Interrupters, April 31, 1988 

HIGH VOLTAGE dc SWITCHGEAR DEVELOPMENT FOR MULTI-KW SPACE POWER 
SYSTEMS - W. Billings, Westinghouse Electrical Corp. Electrical 
Systems Division, NASA CR-165413, WAED— 81-05E 

REMOTE BUS ISOLATOR, Specification No. ED406490-1 150 AMP - 160 

Vdc , Westinghouse Electric Corp., Electrical Systems Division, 

April 25, 1988 

SOLID STATE CIRCUIT BREAKER, Specification No. RV-52157, Lockheed 
Missiles and Space Co., January 9, 1981 

CURRENTS, Westinghouse Electrical Corp. Spring 1986 

SPECIFICATION: POWER CONTROLLER, 270 Vdc, LOAD SWITCHING, SPST, 

NORMALLY OPEN, 1, 2, AND 5 AMPERES, NADC SPECIFICATION No. TS-8001/07 

BACKGROUND AND EXPERIENCE IN THE DEVELOPMENT AND APPLICATION OF 
RPCs - REMOTE POWER CONTROLLERS, Westinghouse Electric Corp. 

TELEDYNE SOLID STATE 1988 DATA BOOK - Teledyne Solid State Division 
of Teledyne Relays 


12-3 


5-15. 

5.16. 

5-17. 

5-18. 

5-19. 

5-20. 

5-21. 

5-22. 

5-23. 

5-24. 

5-25. 

5-26. 

5-27. 

5-28. 

5-29. 


REMOTE POWER CONTROLLER - Product Control Dwgs., for 2, 5, 10, 15, 
20, 50 AMP Devices [CF001102M4A, M5A, M10A, M15A, M20A, M50A - 
Teledyne Solid State 

INTELLIGENT SWITCHING DEVICE (ISD), SSPC (DC-SPST-NO) , Leach Corp. 
and Personal Communication, S. Krauthamer, JPL and S. Atreya of 
Leach Corp., March 30, 1988 

SOLID STATE 270V dc AIRCRAFT POWER CONTROLLERS (N62269-84-R-0249 ) , 
Rockwell International - Autonetics Strategic Systems Division 

POWER CONTROLLER KPC 1; KPC 2 - Kilovac Corp. 

VACUUM REMOTE POWER CONTROLLERS FOR HIGH CURRENT, HVDC, AEROSPACE 
POWER DISTRIBUTION AND LOAD CONTROL APPLICATIONS - H. Jabagchorian, 
Kilovac Corp, Fourth S.A.E. Aerospace Interconnect Systems 
Conference, October 1987 

AEROSPACE TECHNOLOGY DEVELOPMENT OF SOLID STATE REMOTE POWER 
CONTROLLERS FOR 300 VDC WITH CURRENT RATINGS OF ONE AND TWO AMPERES 
ONE TYPE HAVING CURRENT LIMITING, W. Billings, Report No. NASA 
CR-135199 (WAED 77-01E), Westinghouse Electrical Corp., June 1977 

KILOVOLT dc SOLID STATE REMOTE POWER CONTROLLER DEVELOPMENT - Final 
Report - J. Mitchell, Report No. NASA CR 168041, WESDL-82-07E, 
Westinghouse Electric Corp., October 1982 

EXERPT FROM WESD REPORT on NASA Contract NA53-17771 covering 
development of 120 Vdc RPCs , Westinghouse Electric Corp. 

Personal Communication: A. Hilbert of Inland Motors and 
S. Krauthamer, JPL, April 6, 1988, Re; Power Conditioners 

Mid Voltage Power Conditioning, S. Wong, Space Power Inc., Report 
No. F33615-86-C-2674, January 1987 

Personal Communication, S. Krauthamer, JPL and A. Coconi of 
Aeroenvironment Inc., Re; dc/dc Converters and ac Motor Drives, 

May 1988 

THE SECOND ELECTRONICS REVOLUTION: THE IMPACT POTENTIAL OF NEW POWER 
ELECTRONIC TECHNOLOGY ON AIRCRAFT ACTUATION SYSTEMS, J. Lyford, 

GE (ACSD) , NAECON, 1986, IEEE No. 0547-3578/86/0000-0386 

PHOTOVOLTAIC POWER CONDITIONING SUBSYSTEM: STATE-OF-THE-ART AND 

DEVELOPMENT OPPORTUNITIES, S. Krauthamer, et al.. Jet Propulsion 
Laboratory, JPL Pub. 83-81, January 1984 

PERSONAL COMMUNICATION, S. Krauthamer, JPL and N. Britt of Space 
Power Inc., May 1988 

INTEGRATED ACTUATOR SYSTEM, GE Aircraft Control Systems Department, 
General Electric Co., Report No. 525-787 


12-4 



5-30. 

5-31. 

5-32. 

5-33. 

5-34. 

5-35. 

5-36. 

5- 37. 

6 - 1 . 


6 - 2 . 


6-3. 


7-1. 


8 - 1 . 


9-1 


10 - 1 . 


10 - 2 . 


HIGH PERFORMANCE VEHICLE INVERTER REPORT, W. Rippel, D. Edwards, 

JPL, November 1984, JPL D-2199 

MOS CONTROLLED THYRISTORS (MCTs) IN ENERGY SYSTEMS, V. Temple, GE 
Corporate RS<D, IECEC 1987, Paper No. 879252 

SURVEY OF HALL EFFECT LINEAR OUTPUT TRANSDUCERS, S. Krauthamer, JPL, 
Internal Document, November 1987 

FLAT BUS FAULT SENSORS, C. Linder, NADC 77336-60, October 1979, 
Rockwell International Autonetics Strategic Systems Division 

THE LEM MODULE, LEMSA (see Appendix A) 

SPECIFIERS GUIDE FOR SOLID STATE SENSORS, Copy 20 Issue 8, Micro 
Switch Inc. 

A HIGH PERFORMANCE, INTEGRATABLE CURRENT SENSOR, W. Rippel, JPL, 

Power Conversion International, August 1987 

SAFETY RELATED REQUIREMENTS FOR PV ARRAY AND MODULE, A. Levins of 
UL, JPL Report No. 955392-2 

FEASIBILITY STUDY OF A 270 Vdc FLAT CABLE AIRCRAFT ELECTRICAL POWER 
DISTRIBUTION SYSTEM, M. Musga, R. Rinehart, Boeing Aerospace Corp., 
January 1982, Report No. NADC-82023-60 

270 Vdc AIRCRAFT FLAT CABLE DISTRIBUTION SYSTEM INITIAL HARDWARE 
IMPLEMENTATION AND TEST, D. Persons, M. Musge, Boeing Aerospace 
Corp., February 25, 1984, Report No. NADC-82033-60 

TESTING AND EVALUATION OF THE ELECTROMAGNETIC CHARACTERISTICS 
PERFORMED ON THE STACKED FLAT BUS, David T. Brown Inc., April 20, 
1984, Report No. NADC-85114— 60 

CIRCUIT PROTECTION FOR ADVANCED AIRCRAFT - A FUNCTIONAL AND 
HISTORICAL PERSPECTIVE, F. Cannovo, et al. Texas Instruments, SAE 
Paper No. 872502, Aerospace Avionics Equipment and Integration 
Conference, November 2, 1987 

IMPACT OF POWER DISTRIBUTION ON THE SPACE STATION EMI ENVIRONMENT, 

C. Pistole, SAE/P-85 , Paper No. 859218 

ELECTRIC SHOCK SENSATION - THE EFFECT OF FREQUENCY, H. W. Turner and 
C. Turner, Electrical Review 

GENERAL SPECIFICATION FOR CONTROL GROUP, MULTIPLEX ELECTRICAL 
DISTRIBUTION CONTROL, MIL-C-81883, Preliminary December 1986 

MILITARY SPECIFICATION: GENERAL SPECIFICATION FOR CIRCUIT BREAKERS, 

TRIP FREE LOW SIGNAL REMOTE CONTROL, MIL-C-85814 July 1986, 
Preliminary 


12-5 


10-3. 

C-l. 

C-2. 

C-3. 

C-4. 
C-5 . 


SPACE STATION 20 KHz POWER MANAGEMENT AND DISTRIBUTION SYSTEM, 

I.G. Hansen, G.R. Sundberg, NASA, June 23-27, 1986 

DESIGN OF A 2 KW, 100 KHz SWITCHING REGULATOR FOR SPACE SHUTTLE, 

A. Cocconi and S. Cuk, TESLAco— Optimum Power Conversion, Fifth 
International PCI, 1982 

DESIGNING A HYBRID 50 W/in 3 , 500 kHz CUK CONVERTER, Dr. S. Cuk, 
Professor of Electrical Engineering, California Institute of 
Technology and Dr Z. Zhang, Chief Engineer, TESLAco; Part II 
(Magnetics Optimization), Powertechnics Magazine, June 1988 

COUPLED-INDUCTOR ANALYSIS AND DESIGN, S. Cuk and Z. Zhang, Power 
Electronics Group, California Institute of Technology, Power 
Electronics Specialist Conference Record, pp. 655-665 (IEEE 
Publication 86CH2310-1), 1986 

CAPACITORS - INTEGRATED CERAMIC COMPONENTS LTD, December 1987, 

Types 3740 and CC13 Specification Sheet 

PRODUCT ANNOUNCEMENT — QUAD-OUTPUT 100-WATT dc/dc CONVERTERS 
DELIVER 18 WATTS /in 3 Fabricated with hybrid construction 
techniques, these 500— kHz, low-profile converters operate over a 3il 
input voltage range and include emc filtering that meets the 
requirements of MIL-STD-461B, Class Al, Powertechnics Magazine, May 
1988 


12-6 



SECTION 13 


BIBLIOGRAPHY 


ADVANCED ELECTRICAL POWER SYSTEM TECHNOLOGY FOR THE ALL ELECTRIC AIRCRAFT, 

R. Finke, G. Sundberg, NASA, NAECON 1983 

AEROSPACE COUNCIL OF THE SAE TECHNICAL BOARD, MEMBERSHIP LISTS AND ORGANIZATION 
CHART, 1988 

AIRCRAFT ELECTRICAL POWER SYSTEMS, E. Beauchamp, Airesearch Mfg. Co., NAECON, 
1985 

AIRCRAFT POWER SYSTEM TECHNOLOGY - LONG RANGE PLAN, J. Weimer Aero-Propulsion 
Laboratory, SAE -A3 7 , April 20, 1988 

AUTOMATED POWER MANAGEMENT - A TECHNICAL BRIEF, The Leach Corporation, 

April 15, 1988 

CIRCUIT PROTECTION FOR ADVANCED AIRCRAFT - A FUNCTIONAL AND HISTORICAL 
PERSPECTIVE, Texas Instruments Aerospace Avionics Equipment and Integration 
Conf., November 2-4, 1987, Paper No. 87502 

CIRCUIT PROTECTION FOR ADVANCED AIRCRAFT — A Functional and Historical 
Perspective - F. Cannaur, R. Peterson, C. Cobb, Texas Instruments, Aerospace 
Avionics Equipment and Integration, Proceedings of the Second Conference, 

Paper No. 872502 

COMPARATIVE STUDY OF CABLE CONSTRUCTION FOR 20 KHz POWER DISTRIBUTION, W. Putney 
et al., IECEC 1987, Paper No. 879356 

COSTING TOOLS FOR SPACE STATION TRADE STUDIES: POWER DISTRIBUTION SYSTEMS - 
L. Rosenberg, Jet Propulsion Laboratory, Report No. JPL D-3069, October 1986 

DESCRIPTION OF A 20 KILOHERTZ POWER DISTRIBUTION SYSTEM, I. G. Hansen, NASA, 
August 25, 1986, IECEC 1986, NASA Technical Memorandum 87346 

DESIGN CONCEPTS FOR A 5 kW dc TO dc SERIES RESONANT CONVERTER, J. Martin, 

W. Dudley, IECEC, 1987 Paper No. 87315 

DESIGN CONSIDERATIONS FOR LARGE SPACE ELECTRIC POWER SYSTEMS, D. Renz et al., 
April 1983, NASA Technical Memorandum 83064 

DEVELOPMENT AND TEST OF HIGH FREQUENCY dc/dc CONVERTERS FOR COMMUNICATIONS 
SATELLITES - W. Alsbach, N. Tsuya, T. Gohnai, SAE - 1985 Paper 
No. SAE/D-85/ 164-85946 7 

ELECTRIC POWER MANAGEMENT AND DISTRIBUTION FOR AIR AND SPACE APPLICATIONS 
Byron Mehl and Eric Henderson, IECEC 1985 Paper No. 859355 

HIGH VOLTAGE DESIGN GUIDE: SPACECRAFT, W. Dunbar, Boeing Aerospace Company, 

Report No. AFWAL-TR-82-2067 , January 1983 


13-1 



HIGH VOLTAGE dc STABILITY ANALYSIS MODELS VERIFICATION, D. Sommer, T. Liang, 
Boeing Military Airplane Co. Report No. NADC-8 2034-60, February 1984 

HIGH VOLTAGE dc STABILITY ANALYSIS MODELS, D. Sonmer, I. Mendi, Boeing Military 
Airplane Co. Report No. NADC-79048-60 , February 1981 

INCIPIENT FAULT DETECTION AND POWER SYSTEM PROTECTION FOR SPACEBORNE SYSTEMS, 

B. Russell, I. Hackler , IECEC 1987, Paper No. 879349 

INTEGRATED MAGNETICS VERSUS CONVENTIONAL POWER FILTERING, S. Cuk, Power 
Electronics Group, California Institute of Technology, Intelec 87, June 1987 

INTEGRATED MAGNETICS, QUASI-RESONANT CUK DC-TO-DC CONVERTER, S. Cuk, 

Maksimovic , Power Electronics Group, California Institute of Technology, 

High Frequency Power Conversion Conference, May 1988 

MANNED SPACE STATION ATTACHED ACCOMMODATIONS HANDBOOK, PART 6, NASA Goddard Space 
Flight Center, November 20, 1987 

MILITARY SPECIFICATION: GENERAL SPECIFICATION FOR SWITCHES, TRANSDUCER, 

MIL-S-85772(AS ) , Preliminary, December 1986 

RESPONSE AND COMMENTS FOR NASA QUESTIONS ON JEM POWER SYSTEM, June 10, 1986 
NASA of Japan 

SINEWAVE SYNTHESIS FOR HIGH EFFICIENCY DC-AC CONVERSION - Hans K. Asper, IECEC 

1985, Paper No. 859178 

SPACE STATION 20-KHZ POWER MANAGEMENT AND DISTRIBUTION SYSTEM, I. Hansen, 

G. Sundberg, NASA Technical Memorandum 87314, June 23, 1986 

SPACE STATION ELECTRIC POWER SYSTEM REQUIREMENT AND DESIGN, F. TEREN, 

IECEC 1987, Report No. 879003 

SPACE STATION ELECTRICAL POWER DISTRIBUTION SYSTEM DEVELOPMENT - W. E. Murray 
IECEC 1985, Paper No. 859216 ’ 

STS/DBS POWER SUBSYSTEM END-TO-END STABILITY MARGIN, R. Devaux, et al., IECEC 

1986, Paper No. 869398 

20 KILOHERTZ SPACE STATION POWER SYSTEM, I.G. Hansen, F.J. Wolff, NASA, 

September 8, 1986, NASA Technical Memorandum 88801 

270 VDC VARIABLE SPEED GENERATOR AND CONTROL UNIT, AIRCRAFT ELECTRIC POWER 
SYSTEM, Airesearch Mfg. Co., Naval Air Development Center, Technical Report 
(NADC-80014-60) May 1, 1980 


13-2 



APPENDIX A 


ORGANIZATIONS INVOLVED IN HVdc COMPONENTS /SYSTEMS 


Remote Bus Interrupters (RBI) 

Teledyne Kinetics Division 
410 South Cedrus Ave. 

P.O. Box 427 

Solano Beach, CA 92075 

Westinghouse Electric Corp. 
Electrical Systems Division 
P.O. Box 989 
Lima, OH 45802 

Lockheed Advanced Marine Systems 
3929 Calle Fortunada 
San Diego, CA 92123 

Eaton Corp., Cutler Hammer Division 
Hartman Electric Co. 

Milwaukee Center 
Milwaukee, WI 53216 

Flat Cable 

W. L. Gore Associates Inc. 

4747 East Beautiful Lane 
P.O. Box 50699 
Phoenix, AZ 85076-0699 


Flat Cable Connectors 

W.L. Gore Associates Inc. 
4747 East Beautiful Lane 
P.O. Box 50699 
Phoenix, AZ 85076-0699 

G & H Technology 
1647 17th St. 

Santa Monica, CA 90404-3893 


dc/dc Converters 

Inland Motors 
4020 E. Inland Rd. 
Sierra Vista, AZ 95635 


TESLAco Inc. 

490 S. Rosemead #6 
Pasadena, CA 91107 

Space Power Inc. 

621 River Oaks Parkway 
San Jose, CA 95134 

Boeing Electronics Co. 

Power Electronics and Product 
Development 

P.O. Box 24969, MS-9J-12 
Seattle, WA 98124-6269 


Contactors 

Hartman Electric Co. 

175 N. Diamont 
Mansfield, OH 44902 

Eaton Corp., Cutler Hammer Division 
Milwaukee Center 
Milwaukee, WI 53216 

dc Bus Regulator 

Space Power Inc. 

621 River Oaks Parkway 
San Jose, CA 95134 

Aeroenvironment Inc. 

825 South Myrtle Ave. 

Monrovia, CA 

Remote Power Controllers (RPC) 


Kilovac Corp. 

P.O. Box 4422 

Santa Barbara, CA 93140 

Teledyne Solid State 
12424 Daphne 
Hawthorne, CA 90252 

Westinghouse Electric Corp. 
Electrical Systems Division 
P.O. Box 989 
Lima, OH 45802 


A-l 



Rockwell International-Autonetics 
Strategic System Division 
P.0. Box 4192 
3370 Mira Loma Ave. 

Anaheim, CA 92803 

Leach Relay Corp. 

Power Management Group 
6900 Orangethorpe Ave. 

Buena Park, CA 90620 


Motor Controllers 

Aeroenvironment Inc. 

825 South Myrtle 
Monrovia, CA 

General Electric Co. 

Aircraft Control Systems Department 
P.0. Box 5000 
Binghamton, NY 13902 

Jet Propulsion Laboratory 
4800 Oak Grove Drive 
Pasadena, CA 91109 


Current Sensors 

Liaisons Electroniques 

Mechaniques SA Geneve 

Lem SA 140, Chemin du-Centenaire 

CH-1228 Plan-Les-OUATES/Geneva , 

Switzerland 

Micro Switch 

11 West Spring Street 

Freeport, IL 61022 

Jet Propulsion Laboratory 
4800 Oak Grove Drive 
Pasadena, CA 91109 


Syste ms Organizations 

AF Wright Aeronautical Laboratory 
Dept, of the Air Force 
Wright-Patterson AFB, OH 
45433 AFWAL/000C-1 


NASA Lewis Research Center 
Cleveland, OH 44135 

NASA Goddard Flight Center 
Greenbelt, MD 

McDonnell Douglas Corporation 

McDonnell Douglas Astronautics Co. 

St. Louis Division 

P.0. Box 576 

St. Louis, M0 63166 

McDonnell Douglas Corp. 

Douglas Aircraft Corp. 

Long Beach, CA 

General Dynamics 
Land Systems Division 
1902 Northwood 
Troy, NM 48084 

Rockwell International Corp. 
RocketDyne Division 
6633 Canoga Avenue 
Canoga Park, CA 91303 

Boeing Co. 

Boeing Commercial Airplane Co. 
Seattle, WA 

Boeing Co. 

Boeing Military Airplane Co. 
Seattle, WA 

Boeing Co. 

Boeing Aerospace Co. 

Seattle, WA 

Lockheed Corp. 

Burbank, CA 

Newport News Ship Building Co. 

4101 Washington Avenue 
Newport News, VA 23607 

Department of Navy (NAVSEA) 
Electrical Section 
Washington, D.C. 20362-5101 

RCA Astro-Electronics 
Princeton, NJ 08540 


A-2 



Naval Air Development Center 
Warminster, PA 18974 


SAE, Society for Automotive Engineers 
The Engineering Society for 
Advancing Mobility, Land, Sea, Air 
and Space 

400 Commonwealth Drive 
Warrendale, PA 15096-0001 


A- 3 




APPENDIX B 


POWER CONTROLLER EFFICIENCY DATA COMPARISON 

The data presented on developed power conditioners and dc bus 
regulators (Actions 5.3 and 5.4) have indicated t"or»ed 

efficiency. The European Space Agency (ESA) has a: “g^ator efficiencies. This 
=££ £or Ue station. 4 

comparison of these efficiencies is shown in Table B-l. 

The efficiencies presented above is a 

realistic 'representaUon ^ wL ren he expected from available hardware. 


Table B-l. Power Controller Efficiency Data Comparison 


Function 


dc/dc bus regulator (battery 
charger , bus regulator, heater 
and resistor load control) 


dc/dc converter with isolation 
150 Vdc output 

dc/dc converter with isolation 


Efficiency Results 
of this Report 


98. 5% 
Table 5-6 


95% 

Table 5-5 


85% at 5 Vdc output 
up to 90% 
at 15 V-24 Vdc 
Appendix C 


Efficiency 
ESA Study 
[1-3] 


97.3% 


95.5% 


89% at 28 Vdc 


B-l 



APPENDIX C 


STATE-OF-THE-ART TECHNOLOGY IN dc-TO-dc CONVERTERS 

SLOBODAN CUK 
POWER ELECTRONICS GROUP 
CALIFORNIA INSTITUTE OF TECHNOLOGY 
JUNE 17, 1988 

ABSTRACT 


A critical review is made of the state-of-the-art switching power 
supplies operating from the 120 V to 160 Vdc bus. A special emphasis is 
placed on the high power density and high efficiency, light weight performance 
parameters for the power supplies in the broad power range from 25 W to 3 kW, 
in a single output, as well as multiple-output configurations. The cost of 
the high power density converters is also discussed. 


C-l 


INTRODUCTION 


It is clear that by increasing the switching frequency the size 
and mass of magnetic and capacitive components are reduced* leading to light 
weight and compact converters. This has led to the switching frequencies 
being reported in the 10 MHz range* and as high as 24 MHz for a 50 W resonant 
power supply. However, this unchecked increase in switching speed has a very 
undesirable side effect: the reduction of the overall conversion efficiency. 

This is not only the result of the increased switching losses* but also the 
increase of the magnetic core losses* as well as copper losses due to skin 
effect. Therefore, at about 1 MHz switching frequency level, the point of 
diminishing returns is reached, where further increase leads only to a small 
decrease of the overall size* due to primarily magnetic circuit limitations. 
This* however, leaves wide open the region of 100 kHz to 1 MHz in which the 
light weight, high power density, high efficiency converters can be made, 
using the readily available ferrite magnetic core, and standard semiconductor 
and chip capacitor components, and employing the proven PWM (Pulse Width 
Modulated) techniques. The overall power conditioner topology of the CUK 
converter is shown in Figure C-l for a 100 W unit with a 270 Vdc input and a 5 
Vdc output. 


POWER VERSUS OPERATING FREQUENCY AND FIGURE OF MERIT 

The design difficulty in PWM converters can be directly related to 
how much current is switched at what switching rate, which ultimately leads to 
speed power product. Based on that, one may develop a rough figure of merit 
(FM) of design difficulty as per Table C-l. 


Table C-l. FM = Power Switching Speed = 50 x 10^ W/s 


25 W 

50 W 

100 w 

200 W 

400 W 

1 kW 

2 kW 

2 MHz 

1 MHz 

500 kHz 

250 kHz 

125 kHz 

50 kHz 

25 kHz 


This simply outlines that it is roughly as equally challenging to 
design a 50 W converter at a 1 MHz switching rate as it is to design a 20 W at 
a 250 kHz rate, or a 1 kW at a 50 kHz rate. This measure is, of course, only 
to provide the first rough measure of design complexity and is in no way 
representative of the state-of-the-art limits. For example* several years ago 
TESLAco designed a 2 kW, 100 kHz feasibility prototype for the Space Shuttle 
which was successfully demonstrated. With the better semiconductor switching 
devices presently available, as well as better magnetic and capacitive 
components, this design could be even further reduced in weight (from original 
8 lb) by pushing the switching frequency into the 200 kHz or even the 300 kHz 
region. 


This review has utilized several papers describing 50 W and 100 W 
dc-to-dc PWM converters operating at 500 kHz, 50 W resonant switch converters 
operating at 800 kHz, and 2 kW converter operating at 100 kHz [C-l, C-2] . 


C-2 



They all point out that with the presently available components, ultra light 
and high-power density high efficiency converters can be made. Each of the 
specific performance characteristics is addressed separately. 


POWER DENSITY 

Most recently, the state-of-the-art power density of 50 W/in 
has been reported for the converters in the 50 W to 100 W power range. 
However, this does not assume the availability of the cold plate or some o e 
efficient means of the heat removal, such as fan cooled, since the above 
figure does not include the heat sink. However, with the heat sink included, 
the power density of 20 W to 30 W/in 3 is possible with present-day component 

technology. 


At the higher power levels of 500 W to 2 kW, the power density has 
to be reduced due to the increased total power dissipation, and the derating 
of components to a range of 5 W/in 3 . 


EFFICIENCY 


In the 50 W to 200 W range and operating at 500 kHz, the 
efficiency of 80 percent was attainable for a 5 V, 20 A dc/dc converter for 
example. However, if the extremely small size of this unit achieving t e 
ultra high power density of 25 W/in 3 is traded for increase in efficiency, 
primarily through the use of the bigger, more efficient cores and reduced 
switching frequency to 200 kHz or lower, the efficiency in excess of 
85 percent can be obtained, for the 5 V output loads. For higher output 
voltages, such as 15 V or 24 V outputs, efficiency approaching the 90 percent 
could be obtained. This is primarily because the losses are still dominated 
by the output rectifier dc or steady-state losses. For example, in a 5 V 
output with a 0.5 V forward voltage drop on the Schottkey diode, efficiency is 
reduced by 10 percent due to these losses alone. The combined magnetic losses 
typically amount to the range of 1 percent to 2 percent. For example, for a 
50 W converter, 0.8 W total magnetic circuit loss (including transformer and 
all filtering inductors) is typical. Similarly, the typical switching 
transistor losses of 2 W and conduction losses of 1 W further reduce the 
overall efficiency. 


MAGNETIC COMPONENTS 

Presently, the only useable magnetic material at frequencies above 
100 kHz is ferrite. Although saturation flux density of the ferrite material 
is in the range of 0.4 tesla to 0.5 tesla, the flux density at 500 kHz, for 
example, has to be typically derated to the level of 50 mT to 100 mT because 
of the excessive core losses. The derating is accomplished by use of the 
larger size cores. Even with the use of some sophisticated magnetics 
circuits, such as Coupled-Inductor and Integrated Magnetics Circuits 1C 3, 
4-4], the size and weight of the magnetic circuits still dominate the overa 
converter size and weight, and are typically several times that of the 
converter capacitive content. 


C-3 



CAPACITORS 


The switching frequency increase beyond 100 kHz permits, for the 
first time, the replacement and elimination of the unreliable electrolytic 
capacitors by the film capacitors, which offer practically unlimited 
lifetine> They do not suffer from the electrolytic gradual evaporation, over 
time, and the eventual loss of capacitance. The chip capacitor advances made 
in the last several years are particularly dramatic. Not only is the 
packaging density increased by an order of magnitude by use of material with 
increased dielectric constants, packaging some 100 microfarad into a capacitor 
occupying only a 1 cm*" surface and 2 mm high, but they are now offering much 
higher rms ripple current capability and lower ESR. Only several years ago, 

1 A to 2 A capacitor ripple current ratings were available in 3.3 microfarad 
chip capacitor, which has by now been raised to 11 A ripple current ratings 
and 2 to 3 milliohm ESR at 200 kHz. Consequently with a single capacitor such 
as this operating at 200 kHz rates, the output load currents of up to 30 A 
into 5 V load, or 150 W power levels are made possible [C-4]. This clearly 
points out that the magnetics components due to the high core material losses 
are by far the most important factors limiting the size and weight of 
switching converters. 


FREQUENCY RESPONSE 

With the advent of the current mode programing control method, 
the converter loop-gain response may be shaped to look like an ideal 
single-pole response all the way up to a decade below the switching 
frequency. Consequently, the loop-gain can be closed in that region, to 
result in a closed-loop bandwidth approaching one tenth of the switching 
frequency. For example, the 100 W converter operating at 500 kHz may be 
designed to give a 20 kHz bandwidth with an excellent 70 degree phase margin 
(see reference 4—4 for a practical design example). The corresponding 
transient response is equally excellent, leading to a less than 1 percent peak 
voltage overshoot for a half load to full load step current change. The 
corresponding settling time of 100 ps agrees less well with the 20 kHz 
bandwidth. For the switching converters operating at 100 kHz, for example, 
this results in the reasonable bandwidth of 5 kHz to 10 kHz, which is still 
excellent for even the most demanding applications. 


EMI CONSIDERATION AND CONTROL 

At the increased switching frequencies of 100 kHz to 500 kHz, even 
the EMI filters needed for the reduction of both common mode noise and 
differential mode noise are occupying much smaller size and are much lighter. 
For example, for a 100 W converter, a single section common-mode choke 
designed on a toroid with a 0.5 cm diameter and 3 mm height is typically 
sufficient to suppress the noise switching MIL-STD 461 B. These noise 
reduction filters are also accompanied by other noise reduction measures, such 
as Faraday shields in the transformers, which alone can reduce EMI noise by a 
factor of 10 to 20. Other techniques, such as splitting input filtering 
inductor optimally between the top and return leg, might lead to an additional 
order of magnitude of noise reduction (factor of 10 to 30) [C-2] . 


C-4 



multiple output switching power supplies 

At only a slight increase of the overall space, it is now possible 
to take advantage of high frequency switching t0 

multiple output switching power supplies. or e P » packaged in a 

-if 1 a ? v '°f 6 rr ssMS “rf r ly 

ail rai filtering needed to meet military specification requirements .but 

add f ' ^ilTsCll sf ’opfftf ‘ V& kHzfwhicfappears to he the 
best frequency range for the presently available c^c takes 
ebonite tnmpared^th —ch^rmpler .^rngle^utpnt^ ^ lgnal 

about i 25 percent increase in space requirements t shoe d also be 
“tpftfcaff f f I “ ft tf S! Si nff pe «. f »j for 

the worst case changes in the line voltage or load current [C 5J. 


WIDE DYNAMIC RANGE 

The dc-to-dc converters operating over the wide dynamic range of 
1-1 and even 4:1 input voltage variation are now becoming a practica 

thrffgnsffffn cfrefui xTa ^^Iflrfhe 

changes . 


OUTPUT VOLTAGE RIPPLE PERFORMANCE 

While typical commercial power supplies require 1 percent relative 
For example in a 2 kW, 100 kHz design for Space Shuttle, TESLAco ODtainea 
can be made with present day technologies. 


POWER SUPPLY WEIGHT 

The range of power supply weight can be best appreciated using the 
two extreme Simples, ofone hand, the recent *0 « 500 kHz design werghs^O 

i i tiff tn an effective per weight power density ot PS 

fr^her side the Mw! ^0 kHz design [C-l] weighs 7 kg and results » an 

0.5 W/gram power density. 



CONCLUSION 


, . , 1,118 review of the state-of-the-art technology in the switching 

dc-to-dc converters clearly brought out that very high power density 
converters approaching 50 W/in 3 , operating at high efficiency of 85 percent, 
and resulting in a light weight approaching 1 W/gram can be made with the 
present day component technologies. In addition, increased switching 

toTTlelfto « n -H 6 I™** l°° t0 500 182 for P° wer ran « e from 2 kW 

?o vn, W lead i?° a Slde benefit of considerably improving the bandwidth to a 
20 IcHz as well as step-load transient response settling time to less than 
iuu ms • 


C-6 












TECHNICAL REPORT STANDARD TITLE PAGE 


1 . Report No . 


88-26 


2. Government Accession No. I 3. Recipient's Catalog No. 


4. Title and Subtitle 

The State of the Art of dc Power Distribution 
Systems /Components for Space Application 


7. Author (s) 

S. Krauthamer 


9. Performing Organization Name and Address 

JET PROPULSION LABORATORY 
California Institute of Technology 
4800 Oak Grove Drive 
Pasadena, California 91109 


12, Sponsoring Agency Name and Address 

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 
Washington, D.C. 20546 


15. Supplementary Notes 


5e Report Date 

July 1988 


6e Performing Organization Code 


8. Performing Organization Report No. 


10. Work Unit No. 


11. Contract or Grant No. 

NAS 7 -9 18 


13. Type of Report and Period Covered 

JPL Publication 
External Report 


14. Sponsoring Agency Code 

Re241 BH-476-30-40-11-00 


16. Abstract 

This report is a survey of the state of the art of high voltage dc systems and 
components. This information can be used for consideration of an alternative 
secondary distribution (120 Vdc) system for the Space Station. All HVdc 
components have been prototyped or developed for terrestrial, aircraft, and 
space applications, and are applicable for space applications with appropriate 
modification and qualification. HVdc systems offer a safe, reliable, low mass, 
high efficiency and low EMI alternative for Space Station secondary distribution. 


17. Key Words (Selected by Author(s)) 

Spacecraft Propulsion and Power; 
Subsystems; Conversion Techniques. 


18. Distribution Statement 

Unclassified /Unlimited 


19. Security Classif. (of this report) 20. Security Classif. (of this page) 

Unclassified Unclassified 


21. No. of Pages 

22. Price 

89 


















HOW TO FILL OUT THE TECHNICAL REPORT STANDARD TITLE PAGE 


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REVERSE SlbE JPL 0184 R 9/83 




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