NATIONAL AERONAUTICS AND SP AC E ADM I N ISTRATION
Technical Report 32-1086
Surveyor II Mission Report
Mission Description and Performance
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(ACCESSION NUMsVRr-
JET PROPULSION LABORATORY
CALIFORNIA INSTITUTE OF TECHNOLOGY
PASADENA, CALIFORNIA
April 1, 1967
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NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
Technical Refxort'32-1086
Surveyor II Mission Report
il/fission Description and Performance
Approved by:
Howard H. Haglund
Surveyor Project Manager
JET PROPULSION LABORATORY
CALIFORNIA INSTITUTE OF TECHNOLOGY
PASADENA, CALIFORNIA
April 1,1967
TECHNICAL REPORT 32-1086
Copyright© 1967
Jet Propulsion Laboratory
California Institute of Technology
Prepared Under Contract No. NAS 7-100
National Aeronautics & Space Administration
Preface
This document constitutes the Project Mission Report on Surveyor II, the second
in a series of unmanned missions designed to soft-land on the lunar surface.
The report consists of a technical description and an evaluation of engineering
results of the systems utilized in the Surveyor II mission. Contributions from the
major systems which support the Project were used in preparation of the report.
The report for this mission consists of a single volume only. Premature termina-
tion of the Surveyor 11 mission precluded the obtaining of scientific data and results
which would normally be presented as .separate parts of each Mission Report.
JPL TECHNICAL REPORT 32-1086
Contents
I. Introduction 1
A. Surveyor Project Objectives ]
B. Project Description 2
C. Mission Objectives 3
D. Mission Summary 3
II. Space Vehicle Preparations and Launch Operations 5
A. Spacecraft Assembly and Testing 5
B. Launch Vehicle Combined Systems Testing 6
C. Launch Operations at AFETR 7
D. Launch Phase Real-Time Mission Analysis 10
III. Launch Vehicle System I3
A. At/as Stage I3
B. Centaur Stage 14
C. Launch Vehicle/Spacecraft Interface 15
D. Vehicle Flight Sequence of Events 17
E. Performance 19
IV. Surveyor Spacecraft 23
A. Spacecraft System 23
B. Structures and Mechanisms 41
C. Thermal Control 45
D. Electrical Power 47
E. Propulsion 56
F. Flight Control 64
G. Radar 70
H. Telecommunications 75
I. Television 89
V. Tracking and Data Acquisition System 97
A. Air Force Eastern Test Range 97
B. Goddard Space Flight Center 103
C. Deep Space Network 104
VI. Mission Operations System 115
A. Functions and Organization 115
B. Mission-Dependent Equipment 118
C. Mission Operations Chronology 122
JPL TECHNICAL REPORT 32-1086
Contents (contd)
VII. Flight Path and Events ^29
A. Launch Phase ^^9
B. Cruise Phase 129
C. Midcourse Maneuver Phase 133
D. Post-Midcourse and Mission Termination 136
Appendix A. Surveyor // Flight Events 139
Appendix B. Surveyor Spacecraft Configuration 145
Appendix C. Surveyor // Failure Reviev^ Board Recommendations .149
Appendix D. Surveyor // Temperature Histories 156
Glossary I'l
Bibliography 1^-^
Tables
ll-l. Major operations at Cape Kennedy 7
11-2. Surveyor // countdown time summary 10
IV-1. Content of telemetry signals from spacecraft 28
IV-2. Spacecraft instrumentation 30
IV-3. Notable differences between Surveyors I and // 36
IV-4. Surveyor spacecraft reliability 37
lV-5. Spacecraft anomalies 37
IV-6. Surveyor II vibration levels during flight 39
IV-7. Thermal compartment component installation 42
IV-8. Pyrotechnic device 45
IV-9. Comparison of predicted vs actual values 49
IV-10. Flight control modes ^^
IV-ll. Star angles and intensities: indicated vs predicted 67
IV-1 2. Nitrogen gas consumption °°
IV-13. Surveyor II RADVS temperature data 73
IV-1 4. Data from in-flight calibration of spacecraft receiver AGC ... 87
IV-15. Typical signal processing parameter values °°
V-1. AFETR configuration 93
V-2. GSFC Network configuration 104
V-3. Characteristics for S-band and L/S-band tracking systems . . .106
JPL TECHNICAL REPORT 32-1086
Contents (contd)
Tables (contd)
V-4. Operational test schedule 107
V-5. Commands transmitted by DSIF stations 109
VI-1 . CDC mission-dependent equipment support of Surveyor II at
DSIF stations 119
VI-2. Surveyor // command activity 120
VII-2. Injection and uncorrected encounter conditions 133
VII-2. Midcourse maneuver alternatives 137
Figures
ll-l. Surveyor // spacecraft prepared for encapsulation 8
11-2. Atlas/Centaur AC-7 launching Surveyor II 10
11-3. Final Surveyor // launch window design for September 1966 . . .11
lll-l. At/as/Centaur/Surveyor space vehicle configuration 14
III-2. Surveyor/Centaur interface configuration 16
III-3. Launch phase nominal events 18
IV- 1. Surveyor // spacecraft in cruise mode 24
IV-2. Simplified spacecraft functional block diagram 25
IV-3. Spacecraft coordinate system 26
IV-4. Spacecraft coordinates relative to celestial references 27
IV-5. Surveyor // data mode/rate profile 31
IV-6. Terminal descent nominal events 33
IV-7. RADVS beam orientation 34
IV-8. Altitude velocity diagram 34
IV-9. Surveyor II reliability estimates 37
IV-10. Launch-phase accelerometer location 38
IV-11. Surveyor // spin rate profile 40
IV-12. Landing leg assembly 42
IV-13. Antenna/solar panel configuration 43
IV-14. Thermal switch 44
IV-15. Thermal design 46
IV-16. Simplified electrical power functional block diagram 48
IV-17. Regulated output current 50
IV-18. Unregulated output current 50
IV-19. OCR output current 51
IV-20. Solar cell array current 51
JPL TECHNICAL REPORT 32- J 086
Contents (contd)
Figures (contd)
IV-21. Solar cell array voltage 51
lV-22. Main battery manifold pressure 52
IV-23. Main battery voltage 52
IV-24. Main battery discharge current 53
IV-25. Auxiliary battery voltage 53
IV-26. BR preregulator voltage 54
lV-27. 29-v nonessential voltage 54
l\/-28. Unregulated bus voltage 55
IV-29. Actual vs predicted battery energy consumption 56
IV-30. Vernier propulsion system installation 57
IV-31. Vernier propulsion system schematic showing locations of
pressure and temperature sensors 58
IV-32. Vernier engine thrust chamber 59
IV-33. Strain gages and thrust command signals at midcourse .... 60
lV-34. Main retrorocket motor 61
IV-35. Helium-tank and propellont-tank pressures vs time 62
IV-36. High-resolution plot of helium supply pressure during
propellant pressurization 63
IV-37. Simplified flight control functional diagram 64
IV-38. Gas-jet attitude control system block diagram 65
IV-39. Altitude marking radar functional diagram 70
IV-40. Simplified RADVS functional block diagram 72
IV-41. Glystron power supply modulator temperature 73
IV-42. Signal data converter temperature 74
IV-43. Doppler velocity sensor temperature 74
IV-44. Altitude marking radar temperature 75
IV-45. Radio subsystem block diagram 75
IV-46. Total received power, Receiver A 76
IV-47. Receiver A AGC vs GMT 77
lV-48. Total received power. Receiver B 78
IV-49. Receiver B AGC vs GMT 79
IV-50. DSS51 received RF power vs time °^
IV-51. DSS total received power vs GMT ^^
IV-52. Omniantenna A contour map, down-link °2
IV-53. Omniantenna B contour map, up-link °3
IV-54. Omniantenna B contour mop, down-link °^
JPL TECHNICAL REPORT 32-1086
Contents (contd)
V Figures (contd)
IV-55. Omniantenna A, Receiver A signal level vs angular
displacement 85
IV-56. Omniantenna B, Receiver B signal level vs angular
displacement 85
IV-57. Omniantenna B, Transmitter B signal level vs angular
displacement 86
IV-58. Simplified signal processing subsystem block diagram 86
IV-59. Survey TV camera 90
IV-60. Simplified survey TV camera functional block diagram 91
lV-61. Relative tristimulus values of the color filter elements 92
IV-62. TV photometric/colormetric reference chart 92
IV-63. Camera 600-line light transfer characteristic as a function
of brightness (T No.) 93
IV-64. Camera 200-line light transfer characteristic as a function
of lunar brightness 93
IV-65. Camera 600-line light transfer characteristic as a function
of lunar brightness 94
IV-66. Camera 600-line transfer characteristics as a function of
color filter position for the f/4 iris stop 94
IV-67. Camera shading near saturation 95
IV-68. Camera sine-wave response characteristic 95
V-1. Planned launch phase coverage for September 20, 1966 .... 99
V-2. AFETR radar coverage: liftoff through Antigua 100
V-3. AFETR radar coverage: Antigua through Pretoria 100
V-4. AFETR VHP telemetry coverage 101
V-5. AFETR S-band telemetry coverage 102
V-6. DSS 42, Tidbinbilla, Australia 105
V-7. Station tracking periods 108
V-8. DSS received signal level 110
V-9. DSN/GCS communications links Ill
V-10. General configuration of SFOF data processing system 113
VI-1. Organization of MOS 116
VII-1. Earth-moon trajectory and nominal events 130
VII-2. Surveyor // trajectory in earth's equatorial plane 130
VI 1-3. Surveyor II earth track 131
VI 1-4. Surveyor II target, uncorrected impact, and final impact
points 132
VII-5. Surveyor // impact locations 134
JPL TECHNICAL REPORT 32-1086 ix
Contents (contd)
Figures (contd)
VII-6. Midcourse capability contours for September 20 launch .... 135
Vll-7. Effect of noncritical velocity component on terminal
descent parameters 136
D-1. Compartment A transit temperatures 157
D-2. Compartment B transit temperatures 159
D-3. RADVS transit temperatures 160
D-4. Flight control transit temperatures 162
D-5. Survey TV camera transit temperatures 163
D-6. Vernier propulsion transit temperatures 164
D-7. Miscellaneous transit temperatures 168
JPL TECHNICAL REPORT 32-1086
Abstract
Surveyor 11, the second of a series of unmanned missions designed to soft-land
on the moon, was launched from Cape Kennedy, Florida, on September 20, 1966.
After a nominal launch phase and accurate injection into lunar transfer trajectory, a
normal transit phase was achieved until execution of midcourse velocity correction,
when one of the three vernier engines failed to fire, causing unbalanced thrust.
A spacecraft tumbling condition resulted which could not be corrected either by
use of the cold-gas jet system or repeated firings of the vernier engines. The unstable
condition caused premature termination of the mission and prevented attainment
of the flight objectives, for which a soft-landing was prerequisite. A thorough
investigation by a specially appointed Failure Review Board has not disclosed a
specific cause for the faiMre. A technical description of the mission and an evalu-
ation of engineering data obtained are presented herein.
JPL TECHNICAL REPORT 32-1086
I. Introduction
Surveyor II was launched from Cape Kennedy, Florida,
at 12:31:59.824 GMT on September 20, 1966. The launch
vehicle provided a very satisfactory injection into lunar
transfer trajectory, and a nominal mission was achieved
until initiation of midcourse correction thrusting. During
midcourse thrusting, one of the three vernier engines did
not fire and spacecraft moment control was lost. A space-
craft tumbling condition resulted which could not be
corrected either by activation of the cold-gas jet system
or by repeated attempts to fire the vernier engines. In the
unstable mode, spacecraft power was insufficient to com-
plete the transit phase. The mission was terminated when
loss of spacecraft signal occurred at 09:35:00 GMT on
September 22, 1966, about 30 sec after a final command
was sent that fired the retro motor. A thorough investi-
gation by a formally appointed Failure Review Board
has not disclosed a specific failure mode.
A. Surveyor Project Objectives
Surveyor is one of two unmanned lunar exploration
projects currently being conducted by the National Aero-
nautics and Space Administration. The other, Lunar
Orbiter, is providing medium- and high-resolution photo-
graphs over broad areas to aid in site selection for the
Surveyor and Apollo landing programs.
The overall objectives of the Surveyor Project are:
1. To accomplish successful soft landings on the moon
as demonstrated by operations of the spacecraft
subsequent to landing.
2. To provide basic data in support of Apollo.
3. To perform operations on the lunar surface which
will contribute new scientific knowledge about the
moon and provide further information in support
of Apollo.
Prior to the initial Surveyor mission {Surveyor I) it was
planned to utilize the first four Surveyor spacecraft to
satisfy Project Objective 1 above, and to utilize the fol-
lowing three spacecraft to satisfy Objective 2. Preliminary
design was under way for follow-on spacecraft, in addi-
tion to the first seven, which would carry special scientific
instruments to satisfy Objective 3. However, advantage
has been taken of the highly successful Surveyor I mission,
which satisfied Objectives 1 and 2 (for one possible Apollo
landing site) and contributed significantly to the attain-
ment of Objective 3, to accelerate the attainment of
Objective 3 and also shorten the total program.
JPL TECHNICAL REPORT 32-1086
To this end, the follow-on missions have been cancelled
and a plan is being implemented to incorporate on the
remaining missions the most desirable scientific instru-
ments that can be added without major alterations in
Project schedule. The Surveyor II mission was unaffected
by this redirection; the third mission is the earliest upon
which it may be possible to incorporate additional scien-
tific equipment.
On the Surveyor 11 mission, attainment of any of the
Project Objectives was precluded by the spacecraft tum-
bling condition that ensued from the attempted midcourse
correction maneuver.
B. Project Description
The Surveyor Project is managed by the Jet Propulsion
Laboratory for the NASA Office of Space Science and
Applications. The Project is supported by four major ad-
ministrative and functional elements or systems: Launch
Vehicle System, Spacecraft System, Tracking and Data
Acquisition System (T&DA), and Mission Operations
System (MOS). In addition to overall project manage-
ment, JPL has been assigned the management responsi-
bility for the Spacecraft, Tracking and Data Acquisition,
and Mission Operations Systems. NASA/Lewis Research
Center (LeRC) has been assigned responsibility for the
Atlas/Centaur launch vehicle system.
1. Launch Vehicle System
Atlas/Centaur launch vehicle development began as
an Advanced Research Projects Agency program for
synchronous-orbit missions. In 1958, General Dynamics/
Convair was given the contract to modify the Atlas first
stage and develop the Centaur upper stage, and Pratt 6;
Whitney was given the contract to develop the high-
impulse LHj/LO. engines for the Centaur stage.
The Kennedy Space Center, Unmanned Launch Oper-
ations branch, working with LeRC, is assigned the Centaur
launch operations responsibility. The Centaur vehicle
utilizes Launch Complex 36, which consists of two launch
pads (A and R) connected to a common blockhouse. The
blockhouse has separate control consoles for each of the
pads. Pad 36A was utilized for the Surveyor II mission.
The launch of Atlas/Centaur AC-7 on the Surveyor II
mission was the second operational use of an Atlas/
Centaur vehicle, the first having been the successful flight
of AC-10 on the Surveyor I mission. Roth Surveyor mis-
sions utilized the "direct ascent" mode, wherein the
Centaur second stage provided only one continuous bufn
to achieve injection into the desired lunar transfer trajec-
tory. Eight R&D flight tests were conducted in the Centaur
vehicle program prior to the Surveyor II mission. Th'e
final Centaur development flight (AC-9), conducted on
October 26, 1966, subsequent to the Surveyor II mission,
successfully demonstrated capability to launch via "park-
ing orbit" ascent trajectories. The parking orbit ascent
mode, involving a second burn of the Centaur stage (after
a coast in parking orbit), will be utilized on the next
mission and on all but one of the remaining missions.
2. Spacecraft System
The Surveyor spacecraft weight of about 2200 lb and
overall dimensions were established in accordance with
the Atlas/Centaur vehicle capabilities. Three major fea-
tures, first demonstrated on the Ranger missions, were
incorporated in the Surveyor spacecraft system: fully
attitude-stabilized spacecraft, earth-directed high-gain
antenna, and the midcourse maneuver. Demonstration of
TV communication at lunar distances is another Ranger
achievement which has been of value to Surveyor and
the other lunar programs. In addition, the Surveyor space-
craft utilizes several new features associated with the
complex terminal phase of flight and soft-landing: throttle-
able vernier rockets with solid-propellant main retro-
motor; extremely sensitive velocity- and altitude-sensing
radars, and an automatic closed-loop guidance and control
system. The demonstration of these devices on Surveyor
missions is a direct benefit to the Apollo program, which
will employ similar techniques. Design, fabrication, and
test operations of the Surveyor spacecraft are performed
by Hughes Aircraft Company under the technical direc-
tion of JPL.
3. Tracking and Data Acquisition System
The T&DA system provides the tracking and commu-
nications link between the spacecraft and the Mission
Operations System. For Surveyor missions, the T&DA
system uses the facilities of: (1) the Air Force Eastern
Test Range for tracking and telemetry of the spacecraft
and vehicle during the launch phase, (2) the Deep Space
Network for precision tracking, communications, data
transmission and processing, and computing, and (3) the
Manned Space Flight Network and the World-Wide
Communications Network (NASCOM), both of which are
operated by Goddard Space Flight Center.
The critical flight maneuvers and most picture-taking
operations on Surveyor missions are commanded and
JPL TECHNICAL REPORT 32- J 086
recorded by the Deep Space Station at Goldstone, Cali-
fornia (DSS 11), during its view periods. Other stations
which provided prime support for the Surveyor II mission
wfere DSS 42, near Canberra, AustraHa, DSS 51, at
Johannesburg, South Africa, and DSS 72, at Ascension
[ Island; at Cape Kennedy, DSS 71 provided support during
prelaunch and launch operations. In addition, backup
[ support was provided by DSS 61, near Madrid, Spain, and
DSS 12 and 14 (with a 210-ft antenna), Goldstone, Cali-
! fomia.
4. Mission Operations System
The Mission Operations System essentially controls the
spacecraft from launch through termination of the mission.
In carrying out this function, the MOS constantly evalu-
ates the spacecraft performance and prepares and issues
appropriate commands. The MOS is supported in its
activities by the T&DA system as well as special hardware
provided exclusively for the Surveyor Project and referred
to as mission-dependent equipment. Included in this
category are the Command and Data Handling Consoles
installed in the DSS's, the Television Ground Data Han-
dling System, and other special display equipment.
C. Mission Objectives
The specific objectives of each Surveyor mission are de-
noted as "flight objectives." For Surveyor II the flight
objectives were specified in two categories: primary and
secondary.
(1) Primary Flight Objectives:
a. Accomplish a soft landing on the moon at a
site east of the Surveyor I landing point.
b. Demonstrate the capability of the spacecraft to
soft-land on the moon with an oblique approach
angle not greater than approximately 25 deg.
c. Obtain post-landing television pictures and
touchdown dynamics, radar reflectivity, and
thermal data of the lunar surface.
(2) Secondary Flight Objective:
a. Demonstrate the capability of DSS 72 to sup-
port future Surveyor missions.
The early part of the Surveyor II mission was carried
out as planned with the full capability of meeting all of
the above objectives. However, the primary flight objec-
tives could not be achieved owing to the spacecraft
tumbling condition which developed after the midcourse
maneuver was commanded. The secondary objective was
met, with DSS 72 providing useful data during gaps
between the view periods of other DSIF prime stations.
D. Mission Summary
Surveyor II was launched with Atlas/Centaur AC-7
from Pad 36A at Cape Kennedy. Because of holds called
during the countdown to overcome launch vehicle prob-
lems, liftoff was delayed until the end of the first planned
launch window, at 12:31:59.824 GMT on September 20,
1966. Following a nominal direct-ascent boost phase, the
spacecraft was very accurately injected into a lunar trans-
fer trajectory. The uncorrected lunar impact point was
within 150 km of the prelaunch aiming point.
Deep Space Station 72, at Ascension, was the first
DSIF station to achieve one-way lock. Received data
confirmed the satisfactory condition of the spacecraft
and the successful completion of the automatic post-
injection events such as sun acquisition and solar panel
deployment. As planned, DSS 51 was the first station to
establish two-way lock and exercise control of the space-
craft by command.
For approximately the first 16y2 hr of flight, a nominal
mission was achieved, including Canopus star acquisition.
When the three vernier engines were commanded on for
midcourse velocity correction, which lasted 9.8 sec (as
preset by earth-based command). Vernier Engine 3 failed
to provide thrust, causing the spacecraft to tumble at a
rate of about 1.22 rev/sec.
The nitrogen gas-jet system, which is normally enabled
during and after the midcourse velocity correction, oper-
ated for several minutes to stabilize the spacecraft. Al-
though the spin rate was reduced to 0.97 rev/sec, the
gas-jet system was inhibited after about 60% of the gas
had been expended, and it became evident that the
remaining gas supply was insufficient to stop the spinning.
Since the spacecraft was rotating in such a way that
energy could not be obtained from the solar panels, the
only source of electrical power was the spacecraft bat-
teries. Steps were therefore taken to conserve power.
Nevertheless, stored spacecraft power was insufficient to
complete the lunar transit.
JPL TECHNICAL REPORT 32-1086
During the remaining life of the spacecraft, a total of
39 attempts were made to overcome the vernier engine
problem by firing the engines for short periods, ranging
from 0.2 to 2.5 sec and finally, for 21.5 sec. Vernier Engine 3
did not respond to any of these attempts. However,
thrust was delivered by the other two verniers in each
firing, and the spacecraft finally reached a spin rate of
2.3 deg/sec.
About 28.5 hr after the attempted midcourse correc-
tion, when very little battery power remained, a final
sequence was commanded which fired the main retro
motor and Vernier Engines 1 and 2. The spacecraft sig-
nal was lost about 30 sec after main retro ignition, bring-
ing the Surveyor II mission to an end.
The entire Mission Operations System and Deep Space
Network responded well to the unexpected difficulties
which developed in the mission and provided the Project
with tracking and telemetry data as well as the com-
mand function until mission termination.
A Failure Review Board was appointed consisting of
representatives from JPL, NASA Office of Space Science
and Applications, HAC, and Reaction Motors Division
of Thiokol. A thorough investigation conducted by the
Board has not revealed the exact cause of the Surveyor II
spacecraft failure. However, as a result of the detailed
investigation, a number of recommendations have been
made relative to the spacecraft system to assure against a
similar failure as well as to provide better diagnostic data
on future missions.
JPL TECHNICAL REPORT 32-7086
II. Space Vehicle Preparations and
Launch Operations
The Surveyor II spacecraft was assembled and sub-
jected to flight-acceptance testing at the Hughes Aircraft
Corporation facihty, El Segundo, California. After com-
pletion of these tests it was shipped to the Air Force
Eastern Test Range (AFETR), Cape Kennedy, on the
Super Guppy cargo aircraft, arriving on July 19, 1966.
The Atlas/Centaur launch vehicle stages were airlifted
to AFETR after undergoing testing in the Combined
Systems Test Stand (CSTS) at San Diego. Prelaunch
assembly, checkout, and systems tests were accomplished
successfully at AFETR, and the space vehicle was
launched on September 20, 1966 at 12:31:59.824 GMT,
near the end of the first scheduled launch window.
A. Spacecraft Assembly and Testing
Tests and operations on each spacecraft are conducted
by a test team and data analysis team which work with
the spacecraft throughout the period from the beginning
of testing until launch. The test equipment used to con-
trol and monitor the spacecraft system performance at
all test facilities includes (1) a system test equipment
assembly (STEA) containing equipment for testing each
of the spacecraft subsystems, (2) a command and data
handling console (CDC) similar to the units located at
each of the DSIF stations (see Section VI) for receiving
telemetry and TV data and sending commands, and (3)
a computer data system (CDS) for automatic monitoring
of the spacecraft system. Automatic monitoring capability
is necessary because of the large number of telemetered
data points and high sampling frequency of most of the
Surveyor telemetry modes. The CDS provides the fol-
lowing features to aid the data analysis personnel in
evaluating the spacecraft performance:
(1) Digital magnetic tape recording of all input data.
(2) Suppression of nonchanging data. Only data points
which reflect a change are printed on display
devices.
(3) Alarm limit capability. Critical telemetry functions
are monitored for out-of-tolerance indications
which would be damaging to the spacecraft. An
audible alarm sounds if these limits are exceeded.
(4) Request message. In the event that telemetry data
is desired for evaluation, a print of requested data
is provided.
Jn TECHNICAL REPORT 32-1086
The Surveyor II spacecraft (SC-2) was initially assem-
bled December 8, 1964 and then passed through the fol-
lowing test phases:
1. Spacecraft Ambient Testing
The ambient testing phase consists of group tests,
initial system checkout, and mission sequence tests. In
the initial systems checkout, each subsystem is tested
for compatibility and calibration with other subsystems,
and a systems readiness test is performed for initial sys-
tem operational verification. The primary objectives of
the mission sequence tests are to obtain system perform-
ance characteristics under ambient conditions and in the
electromagnetic environment expected on the launch pad
and in flight prior to separation from the Centaur.
After the group tests and initial system checkout,
three mission sequences were completed on the Sur-
veyor U spacecraft. The last of these was a plugs-out
run approaching flight configuration with simulated
electromagnetic environment. A period of extensive re-
work and preparation for solar-thermal-vacuum testing
followed these ambient tests.
2. Solar-Thermal-Vacuum (STV) Testing
The STV sequence of tests is conducted to verify
proper spacecraft performance in simulated missions at
various solar intensities and a vacuum environment. In
these tests, as well as the vibration test phase which fol-
lows, the propellant tanks are loaded with "referee" fluids
to simulate flight weight and thermal characteristics.
The Surveyor 11 spacecraft began the STV sequence
of tests in the solar-vacuum chamber in mid-April 1966.
The first sequence, run at 87% of nominal sun intensity,
was successful except for a major compartment over-
heating problem and failure of the battery case, which
damaged the structure of Compartment A and Trans-
mitter B. The second sequence, run at 112% solar in-
tensity, was aborted when improper operation of two
thermal control switches in Compartment A again caused
compartment overheating. After correction of the thermal
control problem, the test at 112% solar intensity was re-
peated and all systems were normal until the terminal
descent phase, when problems were encountered which
involved a short in the flight control sensor group (FCSG),
improper transmission of telemetry data by the central
signal processor (CSP), and a Canopus tracker malfunc-
tion. Following repair of the FCSG and CSP plus instal-
lation of a new gyro package and Canopus tracker, the
sequence at 112% solar intensity was performed with no
mission critical problems. The third sequence, conducted
in the plugs-out configuration at 100% solar intensity,
was successful up to terminal descent, when the boost-
regulator (BR) failed. The BR failure stopped the
spacecraft telemetry and command systems, preventing
commands from turning off the radar altimeter and dop-
pler velocity sensor (RADVS) and resulted in RADVS
overheating. Repair consisted of replacing the BR and
parts of the RADVS. A final 20-hr STV run was performed
to flight-qualify the system with the new BR and RADVS.
This run was successful, and the spacecraft was removed
from the vacuum chamber and preparations for vibration
testing were initiated.
3. System Vibration Testing
Vibration tests are conducted in the three orthogonal
axes of the spacecraft to verify proper operation after
exposure to a simulated launch-phase vibration environ-
ment. For these tests the spacecraft is placed in the
launch configuration, with legs and omniantennas in the
folded position. In addition, a vernier engine vibration
test is conducted, with vibration input at the vernier
engine mounting points, to simulate the environment
during the midcourse maneuver and terminal descent
phases of flight.
The Surveyor II vibration test phase began on June
29, 1966. The spacecraft was moved to the vibration test
facility and progressed through three axes of vibration
test, with no problems until the final test, during which
the shaker system delivered a severe transient pulse caus-
ing a high shock load of short duration to be received
by the spacecraft. Resulting damage was a partial break-
age of the retro rocket mountings and two antenna and
solar panel positioner (A/SPP) snubbers. The damage
was repaired and the Z-axis vibration was repeated to
complete the test phase on July 10. The vernier engine
vibration and "buzz" tests were successfully completed
on July 13, 1966. An unrelated RADVS crystal detector
failure was found during the vernier engine vibration
phase. This detector was replaced on the spacecraft and
then vibrated by a "sting" drive and functionally tested
to validate the installation. This completed all system
testing at El Segundo and, after preparation. Surveyor II
was shipped to Cape Kennedy via the Super Guppy
cargo aircraft on July 18, 1966.
B. Launch Vehicle Combined Systems Testing
Following successful completion of factory acceptance
testing of each stage, the Atlas was installed in the CSTS
JPt TECHN/CAL REPORT 32-7086
at 'San Diego, California, on April 27, 1966; the Centaur
was installed in the CSTS on May 5, 1966. Test se-
quences in the CSTS culminated in the vehicle Compati-
bility Composite Test on May 25. Test data evaluation
was completed on June 1, 1966. Minor hardware modi-
fications were completed and the Atlas was shipped by
air to AFETR on June 18, followed by the nose fairing
and interstage adapter on June 21 and the Centaur on
June 24, 1966.
C. Launch Operations at AFETR
The major operations performed at AFETR after ar-
rival of the launch vehicle and the spacecraft are listed
in Table II-l.
1. Initial Preparations
The Atlas and Centaur stages of AC-7 were erected on
June 22 and June 29, respectively, and proceeded through
component, system, and ground support equipment (GSE)
compatibility checks with no significant problems. Guid-
ance/autopilot testing, the last major airborne system
test before spacecraft mating, was satisfactorily com-
pleted on August 5.
The spacecraft arrived at AFETR on July 19 and pro-
ceeded through receiving inspection and spacecraft
assembly. Performance Verification Tests (PVT) 1 through
4 and calibration of the TV system were performed with
only minor anomalies occurring.
On August 2, 1966, the spacecraft was transported to
the Explosive Safe Facility (ESF), where it was prepared
for its first trip to the launch pad. Preparations included
(1) installation of a dummy retro rocket and altitude
marking radar (AMR), (2) installation of leg and omni-
antenna squib mufflers, (3) mating of the spacecraft
with the Centaur forward adapter, (4) flight level pres-
surization of the attitude control and propulsion system
tanks, (5) encapsulation of the spacecraft within the nose
fairing (Fig. II-l), and (6) performance of a spacecraft
System Readiness Test (SRT).
Surveyor II was transported to launch pad 36A and
mated with the Centaur on August 9, 1966.
Table 11-1. Major operations at Cape Kennedy
Operation
AC-7* erection
SC-2 inspection, reassembly, and initial testing
SC-2 preparation for on-pod testing; encapsulation and spacecraft
System Readiness Test
SC-2 mate to Cenfour
DSS-7I/SC-2 compatibility test
AC-7 Propellant Tanking Test;
AC-7/SC-2 Joint Flight Acceptance Composite Test (J-FACT)
SC-2 demote
SC-2 decapsulation, depressurizotion, removal of J-FACT items, and
alignment checks
SC-2 final preparation: RADVS ranging test, vernier propulsion phasing,
mission sequence test, and TV photogrammetric calibration
AC-7 Flight Acceptance Composite Test (without SC-2)
SC-2 propellant loading, pressurization, and weight and balance checks
AC-7 Composite Readiness Test (CRT)
SC-2 final encapsulation and spacecraft SRT
SC-2 remote to Centour
Launch
'Atlas/Centaur vehicle designation.
"Survej-or // spacecroft deslgnotlon.
Location
Launch Complex 36A
Building AO
Explosive Safe Facility (ESF)
Launch Complex 36A
Launch Complex 36A
Launch Complex 36A
Launch Complex 36A
Launch Complex 36A
Explosive Safe Facility
Building AO
Launch Complex 36A
Explosive Safe Facility
Launch Complex 36A
Explosive Safe Facility
Lounch Complex 36A
Launch Complex 36A
Dote completed, 1966
June 29
August 3
August 8
August 9
August 1 1
August 1 1
August 1 6
August 16
August 21
August 30
September 8
September 14
September 14
September 15
September 16
September 20
JPL TECHNICAL REPORT 32-1086
Fig. Il-l . Surveyor It spacecraft prepared
for encapsulafion
2. Propellant Tanking Test and Joint Flight Acceptance
Composite Test (J-FACT)
Aftrr tlu- sjiac'ccratt was mated with the Centaur, a
scries of ti'sts were performed to \erify proper operation
with the blockhouse e(jui]5ment and the RF air link be-
tween the lainicli pad and Building A(). All control of
the spacecraft excejit for external jKnver and some moni-
tor functions is at liuilding AO. On August 9 an SRT
and practice countdown were performed, and on August
10 and 11 a compatibility test between the siiacecraft
and DSS 71 at dxpv Kennedy was successfully accom-
plished. During this test a frequenc\- shift problem was
observed when the sjiacecnift transinitter frequency was
transferred from wide-band to narrow-band mode. Tests
on SC-3* demonstrated that a generic drift problem exists
but that the condition was acceptable.
*.S<Tial ih'sinnalUnt tor Si/rr)'i/i>r spaccerafl .selicdnlcd for next
SurvnjDr Diission.
The spacecraft next participated in the launch vehicle
Propellant Tanking Test. During this test the launch
vehicle is completely tanked with propellants, the launch
pad stand is moved back, and the complete vehicle 'is
exi:)osed to off- and on-board RF sources as in the launch
countdown. No RF interference problems were noted.
J-FACT was performed on August 16, with the space-
craft operating in the actual prelaunch environment. This
test culminated in the simulation of launch events
through spacecraft separation. A predicted failure mode
(first seen on SC-3) was ol)served when the auto deploy
logic was not enabled with the Centaur command for
"transmitter to high power." The problem was traced to
diode stray capacitance which turns off the deployment
logic flip-flop. The decision was made to disable the
diode deploy logic and command on the deploy logic
prior to launch.
The spacecraft temperature was maintained constant
during the entire on-pad period to permit accurate pres-
sure decay measurements of the vernier propulsion sys-
tem and attitude control system.
3. Final Flight Preparations
Following demate from the Centaur, the spacecraft
was mo\'ed hack to the ESF, ulicrc if was dccupsuhited
and depressurized. During depressuri/ation a system
pressure ealil:)ration was performed on the helium and
nitrogen pressure transducers. The dummy retro, AMR,
and battery were removed and ahgnment was completed
at the ESF. On August 21, the spacecraft was transferred
back to the Spacecraft Checkout Facility for the per-
formanci' of PVT .5, which consisted of a number of
special tests such as the JMDVS ranging vernier pro-
pulsion system phasing, and the final Mission Sequenci'
Test in the plugs-out configuration. A special photo-
grammetric calibration of the television subsystem was
also performed at this time to improve the accuracy of
view angle calibration. 1 his would permit more accurate
mapping of surface featuri's on the moon.
On August .'11 the sjiacecraft was moved back to the
ESF for final preparations, including propellant loading,
final weight, balance and alignment checks, and per-
formance of the PVT 6 test series. Somi' thermal surfaces
were damaged during i:)ropellant loading and retjuired
rework. Propellant lumes entrap])ed in the protective bag
attacked tlu' tlu'rmal finish, particularly around Landing
1 .eg 3. 1 ndi vidual steps of PVT 6 test sequence were phased
with the final jireparations to verify flight readiness.
JPL TECHNICAL REPORT 32-1086
-The final weight, balance, and alignment checks were
conducted after loading of the vernier system propel-
lants and installation of the main retro motor. A problem
was encountered in alignment when the A/SPP was
stepped. The A/SPP polar axis missed steps in a certain
sector of its travel. Analysis indicated that the unit prob-
ably had not degraded but that the test fixture subjected
the unit to unusual loading which could not occur in the
mission. The unit was accepted for flight on this basis.
After the spacecraft was mated to the Centaur adapter,
the helium and nitrogen tanks were brought to flight
pressures, the spacecraft was encapsulated, and a final
SRT was performed on September 15. On September 16,
the spacecraft was moved to the launch pad and mated
to the Centaur.
An Atlas/Centaur FACT was performed on September
8 with only one anomaly occurring. The FACT count-
down was conducted in a routine manner until T-5 min,
at which time a hold was called because of an indication
of excessive Centaur inverter temperature. The problem
was found to be within the transducer recorder; it was
corrected, and the count was resumed after a hold of
1 min.
The launch vehicle CRT, the last multiple systems test
prior to spacecraft mating and launch countdown, was
I conducted on September 14 and proved the launch readi-
ness of all Atlas and Centaur electrical and RF systems.
No abnormal deviations of test events were noted.
4. Countdown and Launch
Final spacecraft and launch vehicle checks began
immediately after spacecraft mating on September 16.
After a spacecraft SRT was performed, the retro motor
safe-and-arm check was performed. Following installa-
tion of the launch vehicle ordnance devices, the pyro-
technic circuits were checked and launch readiness tests
were started. Atlas fuel tanking and Centaur hydrogen
peroxide loading were accomplished on September 17.
During launch vehicle checkout operations on September
19 (launch day minus 1), Centaur engine feedback traces
did not appear normal. Troubleshooting isolated the
problem to the aerospace ground equipment (AGE) de-
modulator in the instrumentation circuit. The Atlas tele-
metry package and its accessory were replaced because
of improper operation of the commutator.
The final spacecraft SRT began at 19:16 EST on
September 19 at a countdown time of 7-615 min and
was completed at 7-430 min.
At 7-260 min, activation of the Atlas telemetry battery
resulted in a low open-circuit voltage, which reflected a
decaying trend. A spare battery was installed, activated,
and successfully load tested with no impact on the count.
After service tower removal at approximately 7-120
min, a weak signal strength was recorded for spacecraft
Receiver B. First evaluation of this anomaly attributed
the problem to a poor RF link caused by the tower
removal. (Later, during Canopus acquisition maneuvers,
it was confirmed that the signal was indeed low all
across the antenna pattern. Subsequent checks disclosed
that an AGC shift had occurred between final spacecraft
encapsulation and arrival at the launch pad. This prob-
lem is also discussed in Section VI-C, Mission Opera-
tions Chronology.)
The countdown proceeded without interruption to
7-90 min (04:01 EST), when a scheduled 70-min hold
was started. The spacecraft system joined the launch
vehicle countdown during this period. The count was
resumed at 05:11 EST and proceeded as planned to
7-5 min, when a second scheduled hold of 15 min was
started. This hold was extended an additional 7 min to
investigate an apparent low temperature indication
within the Centaur hydrogen peroxide engine system.
The temperature was determined to be correct; the count
resumed at 06:58 EST and proceeded down to 7-115 sec,
when it was necessary to hold and recycle to 7-5 min.
This hold was called because the Atlas liquid oxygen
(LO2) boil-off valve did not close at the start of the
flight pressurization sequence and the proper pressure
could not be reached. During the investigation of the
problem, the Atlas was pressure-cycled several times,
with the LO:; boil-off valve closing properly each time.
It was then decided to resume count as soon as the Atlas
LO. supply, which had fallen below the acceptable
level, could be replenished. At this time it was dis-
covered that the automatic LO,, topping sy.stem was not
operating properly, and manual operation was employed.
After a hold of 26 min, the count was picked up at 07:27
EST, although the Atlas LO, was still below the proper
level, because the launch window for the day closed at
07:32. The flight level was reached at approximately
7-3 min, and the count continued down to liftoff (Fig.
II-2), which occurred at 07:31:59.824 EST (12:31:59.824
GxMT), September 20, 1966. The flight azimuth was
114.361 deg.
The countdown included a total of 85 min of planned,
built-in holds — one of 70-min duration at 7-90 min, and
a second of 15-min duration at 7-5 min. AC-7/SC-2 con-
sumed a total hold time of 118 min. The launch window
JPL TECHNICAL REPORT 32-1086
Table 11-2. Surveyor // countdown time summary
Fig. 11-2. Af/os/Cenfaur AC-7 launching Surveyor //
for September 20 extended from 06:56 to 07:32 EST,
providing a duration of 36 min. A countdown time sum-
mary is shown in Table II-2.
The general performance of GSE was satisfactory
throughout the launch countdown. Commercial power
failed at T I 60 sec, and a Cape-wide loss was experienced
for approximately 5 sec. Except for some loss of optical
coverage, the power loss did not cause any significant
problems, since most facilities were operating from criti-
cal or generator power.
With the exception of an anomaly in the Centaur
telemetry system, all systems performed normally through-
out the launch, and the spacecraft was injected into an
accurate lunar trajectory. The anomaly in the telemetry
system was a failure to receive spacecraft vibration data
from three of five accelerometers. The powered flight
sequence of events and launch vehicle perfonnance are
described in Section III.
Event
Countdown
time
EST
(September 19)
Started spccecra^t SRT
T-615 min
19:16
Completed spocecroft SRT
T-430 min
22:21
(September 20)
Started ]aunch veh'cle countdown
T-300 min
00:31
Started 70-mln built-in hold (BIH)
T-90 min
04:01
Spacecraft joined launch vehicle
r-90 min
04:35
countdown
End BIH; resumed countdown
r-90 min
05:11
Started 15-min BIH
7-5 min
06:36
BIH extended
7-5 min
06:51
Resumed countdown
T-5 min
06:58
Hold; recycled to T-5 min
T-llSsec
07:01
Resumed countdown
T-5 min
07:27
Liftoff
T-0 min
07:31:59.824
The atmospheric conditions on launch day were favor-
able. Surface winds were 6 to 12 knots from 190 deg,
with unrestricted visibility of 10 miles. Surface tempera-
ture was 79° F, with relative humidity of 93% and a
dcwpoint of 77 °F. Sea level atmospheric pressure was
29.900 ill. Cloud cover consisted of 0.10 cumulus at
1,400 ft, 0.10 altocumulus at 11,000 ft, and 0.5 cirrus at
an unknown altitude. Maximum winds aloft were re-
ported to be 63 ft/sec from 22.5 to 245 deg at ^3,000 ft
altitude, with a decrease to .50 ft/sec at 60,000 ft. The
maximum expected shear parameter was 4 ft/see per
thousand feet of altitude.
D. Launch Phase Real-Time Mission Analysis
The launch windows which were finally established
for the September 1966 launch period are shown in Fig.
1 1-3. Launching on days jirior to September 20 was not
acceptable because touchdown at the desired lunar land-
ing site would have occurred in darkness (prior to sun-
rise). The "performance constraint" was based upon a
requirement on the Surveyor II mission that there be a
minimum excess Centaur propellant weight of 235 lb to
cover 3-siu;ma launch vehicle performance dispersions.
This represents an increase from the value of 175 lb used
for the Surveyor I mission to provide additional pro-
tection against the uncertainties as.sociated with the
determination of performance dispersions. The "post-
MECO* tracking constraint" was ba.sed upon the Class I
*Main engine cutoff.
10
JPl TECHNICAL REPORT 32- J 086
LAUNCH DAY
SEPTEMBER 20 SEPTEMBER 21 SEPTEMBER 22 SEPTEMBER 23
PERFORMANCE
CONSTRAINT
POST-MECO
TRACKING
CONSTRAINT
(ANTIGUA
COVERAGE)
GMT
Fig. 11-3. Final Surveyor II launch window design for September 1966
requirement (see Section V) for 60 sec of post-MECO
radar tracking data for which it was considered impor-
tant to have Antigua coverage. Station coverage capa-
bility for the launch phase of the Surveyor II mission is
presented in Section V.
1. Countdown to Launch
During countdown operations, those factors acting to
constrain the launch window or period were continually
evaluated by the Launch Phase Mission Analyst. The
Mission Director was advised of these evaluations for
consideration in the launch or hold decision.
Several temporary problems occurred relative to the
Tracking and Data Acquisition (T&DA) facilities dur-
ing the countdown. These problems included the follow-
ing: (1) interruptions were experienced in the teletype
circuits to DSS 42 and 72 and in the voice circuits to
DSS 51 and 72; (2) a marginal condition developed in
the high-speed data line to DSS 72; (3) Bermuda and
Trinidad radars were inoperative temporarily owing to
elevation encoder and coherent memory filter problems,
respectively; (4) RF propagation fade interfered with the
transmission links with Range Instrumentation Ship (RIS)
Coastal Crusader; and (5) a heavy storm near Canberra
threatened DSS 42. The Mission Director was advised
that none of these problems constituted a hold condition,
and the T&DA system was in a go condition at liftoff.
2. Launch to DSN Acquisition
During the launch-to-DSN-acquisition phase of flight,
the occurrence of "mark" events was reported in real
time by the AFETR and MSFN, followed later with a
report of the times at which they occurred. The small
deviations of the mark times from nominal were judged
to be well within the S-sigma dispersions. The Centaur
bum time was about 2.5 sec longer than expected; how-
ever, similarly longer Centaur bum times had been ex-
perienced on previous flights. Consequently, the powered
flight was considered to be quite normal.
Launch vehicle telemetry was retransmitted to Cape
Kennedy in real-time from all stations down to Antigua
until the spacecraft transmitter was switched to high
power, at which time the spacecraft S-band data was
JPL TECHNICAL REPORT 32-7086
11
sent up the subcable and real-time retransmission of
launch vehicle data was ceased. Reports of the real-time
analysis of launch vehicle data indicated a nominal
powered flight. The normality of powered flight was con-
firmed by reports of nominal acquisition times by each
of the tracking stations.
Four minutes after MECO, the AFETR real-time com-
puter system (RTCS) at Cape Kennedy had computed
the first orbit based on Antigua radar data. This orbit,
considered a "fair" fit, further indicated a nominal
powered flight and injection into a satisfactory lunar
transfer orbit. There were reports, however, which indi-
cated the slight possibility that spacecraft separation had
not been normal. It was reported that Trinidad, which
is capable of tracking more than one object, did not s«e
separation. Also, the intermittent data from Ascension
and Pretoria was thought possibly to have been caused
by a tumbling vehicle. This concern was proved un-
founded when it was learned that unfavorable aspect
angles may have prevented Trinidad from observing
separation, and near-real-time voice reports on the Cen-
taur roll, pitch, and yaw rate gyros confirmed that a
stable vehicle had been observed by Ascension.
12
JPL TECHNICAL REPORT 32-J086
III. Launch Vehicle System
The Surveyor spacecraft was injected into its lunar
transit trajectory by a General Dynamics Atlas/Centaur
vehicle (AC-7). The vehicle was launched on a "direct
ascent" powered flight from Launch Complex 36A of the
AFETR at Cape Kennedy, Florida. This was the second
operational flight of an Atlas/Centaur vehicle, the first
having been the successful flight of AC- 10 on the Sur-
veyor I mission. AC-10 and AC-7 were identical in all
essential respects except for the use of 1000-lb-thrust
vernier engines on the AC-7 Atlas stage. Later-model
vernier engines, rated at 670-lb thrust, were used on
AC-10, which was assembled after AC-7.
The Atlas/Centaur vehicle with the Surveyor space-
craft encapsulated in the nose fairing is 113 ft long and
weighs 303,000 lb at liftoff (2-in. rise). The basic diameter
of the vehicle is a constant 10 ft from the aft end to the
base of the conical section of the nose fairing. The con-
figuration of the completely assembled vehicle is illus-
trated in Fig. III-l. Both the Atlas first stage and Centaur
second stage utilize thin-wall, pressurized, main pro-
pellant tank sections of monocoque construction to pro-
vide primary structural integrity and support for all
vehicle systems. The first and second stages are joined
by an interstage adapter section of conventional sheet
and stringer design. The clamshell nose fairing is con-
structed of laminated fiberglass over a fiberglass honey-
comb core and attaches to the forward end of the
Centaur cylindrical tank section.
A. Atlas Stage
The first stage of the Atlas/Centaur vehicle is a modi-
fied version of the Atlas D used on many previous NASA
and Air Force missions such as Ranger, Mariner, and
OGO. The Atlas propulsion system bums RP-1 kerosene
and liquid oxygen in each of its five engines to provide
a total liftoff thrust of approximately 388,000 lb. The
individual sea-level thrust ratings of the engines are: two
booster engines at 165,000 lb each; one sustainer engine
at 57,000 lb; and two vernier engines at 1000 lb each. The
Atlas can be considered a IVa-stage vehicle because the
"booster section," weighing 6000 lb and consisting of
the two booster engines together with the booster turbo
pumps and other equipment located in the aft section, is
jettisoned after about 2.5 min of flight. The sustainer and
vernier engines continue to bum until propellant deple-
tion. A mercury manometer propellant utflization system
is used to control mixture ratio for the purpose of mini-
mizing propellant residuals at Atlas burnout.
JPL TECHNICAL REPORT 32-1086
13
NOSE FAIRING-
FORWARD EQUIPMENT.
COMPARTMENT
LIQUID HYDROGEN
TANK
INSULATION ■
PANEL (4)
LIQUID OXYGEN -
TANK
CENTAUR MAIN-
ENGINE (2)
INTERSTAGE -
ADAPTER
LIQUID OXYGEN -
TANK
ANTI SLOSH BAFFLE -
ASSEMBLY
LIQUID OXYGEN-
DUCT
FUEL TANK ■
EQUIPMENT POD (2)~
ANTIVORTEX-
BAFFLE
ATLAS RETR0R0CKET(8)'
VERNIER THRUST -
CHAMBER (2)
BOOSTER THRUST'
CHAMBER (2)
^SURVEYOR
"SPACECRAFT
y CENTAUR
r STAGE
^
L^
1^
V ATLAS
r STAGE
^SUSTAINER THRUST
CHAMBER
Fig. III-1. Af/as/Cenfaur/Surveyor space vehicle
configuration
Flight control of the first stage is aecoinplished by the
Mlas autopilot, which contains displacement gyros for
attitude reference, rate gyros for response damping, and
a programmer to control flight sequencing until Atlas-/
Centaur separation. After booster jettison, the Atlas auto-
pilot also is fed steering commands from the all-inertial
guidance set located in the Centaur stage. Vehicle atti-
tude and steering control are achieved by the coordinated
gimballing of the five thrust chambers in response to
autopilot signals.
The Atlas contains a single VHF telemetry system
which transmits data on 108 first-stage measurements un-
til Atlas separation. The system operates on a frequency
of 229.9 mc over two antennas mounted on opposite sides
of the vehicle at the forward ends of the equipment pods.
Redundant range-safety command receivers and a single
destructor unit are employed on the Atlas to provide the
Range Safety Officer with means of terminating the flight
by initiating engine cutoff and destroying the vehicle.
The system is inactive after normal Atlas staging occurs.
The AZUSA tracking system has been deleted from the
Atlas for Surveyor missions, leaving only the C-band
tracking system on the Centaur stage.
B. Centaur Stage
The Centaur second stage is the first vehicle to utilize
liquid hydrogen/liquid oxygen, high-specifie-impulse pro-
pellants. The cryogenic propellants require special in-
sulation to be used for the forward, aft, and intermediate
bulkheads as well as the cylindrical walls of the tanks.
The cylindrical tank section is thermally insulated by
four jettisonable insulation panels having built-in fairings
to accommodate antennas, conduits, and other tank pro-
trusions. The insulation panel hinges were redesigned
for AC-10 to overcome a deployment control problem
which had been suspected on vehicle development flights
AC-6 and AC-8. Most of the Centaur electronic equip-
ment packages are mounted on the forward tank bulk-
head in a compartment which is air-conditioned prior to
liftoff.
The Centaur is powered by two constant-thrust en-
gines rated at 433-sec specific impulse and 15,000 lb
thrust each in vacuum. Each engine can be gimballed to
provide control in pitch, yaw, and roll. Propellant is fed
from each of the tanks to the engines by boost pumps
driven by hydrogen peroxide turbines. In addition, each
engine contains integral "boot-strap" pumps driven by
hydrogen propellant, which is also used for regenerative
cooling of the thrust chambers. A propellant utilization
system is used on the Centaur stage to achieve minimum
residual of one propellant at depletion of the otiier. The
system controls the mixture ratio valves as a continuous
14
JPL TECHNICAL REPORT 32-1086
function of propellant in the tanks by means of tank
probes and an error ratio detector. The nominal oxygen/
hydrogen mixture ratio is 5:1 by weight.
The second-stage all-inertial guidance system contains
an on-board computer which provides vehicle steering
commands after jettison of the Atlas booster section. The
Centaur guidance signals are fed to the Atlas autopilot
until Atlas sustainer engine cutoff and to the Centaur auto-
pilot after Centaur main engine ignition. Platform gyro
drifts are compensated for by the guidance system com-
puter, which is programmed to set the torquing signals to
zero during flight. The Centaur autopilot system provides
the primary control functions required for vehicle stabili-
zation during powered flight, execution of guidance sys-
tem steering commands, and attitude orientation follow-
ing the powered phase of flight. In addition, the autopilot
system employs an electromechanical timer to control the
sequence of programmed events during the Centaur phase
of flight, including a series of commands required to be
sent to the spacecraft prior to spacecraft separation.
The Centaur reaction control system provides thrust to
control the vehicle after powered flight. For small correc-
tions in yaw, pitch, and roll attitude control, the system
utilizes six individually controlled, fixed-axes, constant-
thrust, hydrogen peroxide reaction engines. These engines
are mounted in clusters of three, 180 deg apart on the
periphery of the main propellant tanks at the interstage
adapter separation plane. Each cluster contains one 6-lb-
thrust engine for pitch control and two 3.5-lb-thrust
engines for yaw and roll control. In addition, four 50-lb-
thrust hydrogen peroxide engines are installed on the aft
bulkhead, with thrust axes parallel with the vehicle axis.
These engines are for use during retromaneuver and for
executing larger attitude corrections if necessary.
The Centaur stage utilizes a VHF telemetry system
with a single antenna transmitting through the nose fair-
ing cylindrical section on a frequency of 225.7 mc. The
telemetry system provides data on 149 measurements
from transducers located throughout the second stage
and spacecraft interface area as well as a spacecraft
composite signal from the spacecraft central signal pro-
cessor.
Redundant range safety command receivers are em-
ployed on the Centaur, together with shaped charge
destruct units for the second stage and spacecraft. This
provides the Range Safety Officer with means to terminate
the flight by initiating Centaur main engine cutoff and
destroying the vehicle and spacecraft retrorocket. The
system can be safed by ground command, which is nor-
mally transmitted by the Range Safety Officer when the
vehicle has reached injection energy.
A waiver has been obtained for Surveyor missions to
permit elimination of the inadvertent separation system,
which was designed to provide for the automatic destruc-
tion of the Centaur and spacecraft in the event of pre-
mature spacecraft separation.
A C-band tracking system is contained on the Centaur
which includes a light-weight transponder, circulator,
power divider, and two antennas located under the in-
sulation panels. The C-band radar transponder provides
real-time position and velocity data for the Range Safety
Instantaneous Impact Predictor as well as data for use
in guidance and trajectory analysis.
C. Launch Vehicle/Spacecraft Interface
The general arrangement of the Surveyor/Centaur in-
terface is illustrated in Fig. III-2. The spacecraft is com-
pletely encapsulated within a nose fairing/adapter system
in the final assembly bay of the Explosive Safe Facility
prior to being moved to the launch pad. This encapsula-
tion provides protection for the spacecraft from the en-
vironment before launch as well as from aerodynamic
loads during ascent.
The spacecraft is first attached to the forward section
of a two-piece, conical adapter system of aluminum sheet
and stringer design by means of three latch mechanisms,
each containing a dual-squib pin puller. The following
equipment is located on the forward adapter: three sep-
aration spring assemblies each containing a linear poten-
tiometer for monitoring separation; a 52-pin electrical
connector with a pyrotechnic separation mechanism; three
pedestals for the spacecraft-mounted separation sensing
and arming devices; a shaped-charge destruct assembly
directed toward the spacecraft retromotor; an accelerom-
eter for monitoring lateral vibration at the separation
plane; and a diaphragm to provide a thermal seal and to
prevent contamination from passing to the spacecraft
compartment from the Centaur forward equipment com-
partment.
The low-drag nose fairing is an RF-transparent, clam-
shell configuration consisting of four sections fabricated
of laminated fiberglass cloth faces and honeycomb fiber-
glass core material. Two half-cone forward sections are
brought together over the spacecraft mounted on the
forward adapter. An annular thermal bulkhead between
JPL TECHNICAL REPORT 32-7086
15
NOSE-FAIRING
PARTING LINE
SURVEYOR
SPACECRAFT
SPACECRAFT
SEPARATION LATCHES (3)
ADAPTER
DIAPHRAGM
SPACECRAFT -
SEPARATION
PLANE
VENT HOLES (30)
NOSE-FAIRING
CYLINDRICAL SECTION
CENTAUR FORWARD
EQUIPMENT AREA
NITROGEN JETTISON
BOTTLE (2)
JETTISON COMPARTMENT
BULKHEAD
PYROTECHNIC LATCH
(4 EACH SIDE)
52-PIN ELECTRICAL
DISCONNECT
FORWARD ADAPTER
ADAPTER FIELD
JOINT
THERMAL
BULKHEAD
AFT ADAPTER
CENTAUR FORWARD
BULKHEAD
Fig. III-2. Surveyor/Cenfaur interface configuration
the adapter and base of the conical section completes
encapsulation of the spacecraft.
The encapsulated spacecraft assembly is mated to the
Centaur at a flange field joint requiring 72 bolts between
the forward and aft adapter sections. The remaining two
half-cylindrical sections of the nose fairing are attached
to the forward end of the Centaur tank around the equip-
ment compartment prior to mating of the spacecraft.
Doors in the cylindrical sections provide access to the
adapter field joint. The electrical leads from the forward
adapter are carried through three field connectors and
routed across the aft adapter to the Centaur umbilical
connectors and to the Centaur programmer and telemetry
units.
Special distribution ducts are built into the nose fairing
and forward adapter to provide air conditioning of the
spacecraft cavity after encapsulation and until liftoff.
Seals are provided at the joints to prevent shroud leakage
except out through vent holes in the cylindrical section.
The entire nose fairing is designed to be ejected by
separation of two clamshell pieces, each consisting of a
conical and cylindrical section. Four pyrotechnic pin-
puller latches are used on each side of the nose fairing
to carry the tension loads between the fairing halves. A
bolted connection, with a flexible linear-shaped charge
for separation, transmits loads between the nose fairing
and Centaur tank. A nitrogen bottle is mounted in each
half of the nose fairing near the forward end to supply
gas for cold gas jets to force the panels apart. Hinge
fittings are located at the base of each fairing half to
control ejection, which occurs under vehicle acceleration
of approximately 1 g.
16
JPL TECHNICAL REPORT 32-7086
D. Vehicle Flight Sequence of Events
All vehicle flight events occurred as programmed at
Bear nominal times with no anomalies. Predicted and
actual times for the vehicle flight sequence of events are
included in Table A-1 of Appendix A. Figure III-3 illus-
trates the major nominal events. Following is a brief
description of the vehicle flight sequence of events with all
times referenced to liftoff (2 in. rise) unless otherwise noted.
1. Atlas Booster Phase of Flight
Hypergolic ignition of all five Atlm engines was in-
itiated 2 sec before liftoff. Vehicle liftoff occurred at
12:31:59.824 GMT on September 20, 1966, only 8.5 sec
prior to closing of the launch window. The launcher
mechanism is designed to begin a controlled release of
the vehicle when all engines have reached nearly full
thrust. At 2 sec after liftoff, the vehicle began a 13-sec pro-
grammed roll from the fixed launcher azimuth setting of
105 deg to the desired launch azimuth of 114.36 deg. The
programmed pitchover of the vehicle began 15 sec after
liftoff and lasted until booster engine cutoff (BECO).
The vehicle reached Mach 1 at 58 sec and maximum
aerodynamic loading occurred at 75.7 sec. During the
booster phase of flight the booster engines were gim-
balled for pitch, yaw, and roll control, and the vernier
engines were active in roll control only while the sustainer
engine was centered.
At 142.2 sec, BECO was initiated by a signal from the
Centaur guidance system when vehicle acceleration
equalled 5.76 g (expected value: 5.7 ±0.08 g). At 3.1 sec
after BECO, with the booster and sustainer engines
centered, the booster section was jettisoned by release of
pneumatically operated latches.
2. Atlas Sustainer Phase of Flight
At BECO + 8 sec the Centaur guidance system was
enabled to provide steering commands for the Atlas sus-
tainer phase of flight. During this phase the sustainer
engine was gimballed for pitch and yaw control, while
the verniers were active in roll. The Centaur insulation
panels were jettisoned by firing shaped charges at
176.0 sec at an altitude of approximately 50.5 nm where
the aerodynamic heating rate was rapidly decreasing. At
201.9 sec, squibs were fired to unlatch the clamshell nose
fairing, which was jettisoned 1.0 sec later by means of
nitrogen gas thruster jets activated by pyrotechnic valves.
Other programmed events which occurred during the
sustainer phase of flight were: the unlocking of the
Centaur hydrogen-tank vent valve to permit venting as
required to relieve hydrogen bofloff pressure; starting of
the Centaur boost pumps 43 sec prior to Centaur main
engine ignition (MEIG); and locking of the Centaur
oxygen-tank vent valve followed by oxygen-tank pres-
surization.
Sustainer and vernier engine cutoff (SECO and VECO)
occurred at 235.1 sec as a result of fuel depletion with
oxidizer depletion imminent. Oxidizer depletion with fuel
depletion imminent was the predicted cutoff mode. If
cutoff had been as predicted, shutdown would have be-
gun with an exponential thrust decay phase of about
5-sec duration due to low oxidizer inlet pressure to the
turbopump and resulting loss in turbopump performance.
Then, final fast shutdown by propellant valve closure
would have been initiated by actuation of a switch when
fuel manifold pressure dropped to 625 psi. In the pre-
dicted mode, reduction in fuel manifold pressure is
caused by reduction in speed of the turbopump, which
also utilizes the main fuel and oxidizer propellants. On
this flight, only about 2 sec of exponential decay occurred,
due to decay in oxidizer inlet pressure to the turbopump,
before fast shutdown was initiated by the uncovering of
a fuel depletion sensor located near the bottom of the
fuel tank.
Separation of the Atlas from the Centaur occurred
1.9 sec after SECO by firing of shaped charges at the
forward end of the interstage adapter. This was followed
by ignition of eight retrorockets located at the aft end of
the Atlas tank section to back the Atlas, together with
the interstage adapter, away from the Centaur.
3. Centaur Phase of Flight Through Spacecraft
Separation
The Centaur main engines were ignited 9.6 sec after
Atlas/Centaur separation and burned for 439.7 sec, or
until 686.25 sec. Main engine cutoff (MECO) was com-
manded by the guidance system when the desired injec-
tion conditions were reached. At main engine cutoff, the
hydrogen peroxide engines were enabled for attitude
stabilization.
During the 66.3-sec period between MECO and space-
craft separation, the following signals were transmitted
to the spacecraft from the Centaur programmer: extend
spacecraft landing gear; unlock spacecraft omniantennas;
turn on spacecraft transmitter high power. An arming
JPL TECHNICAL REPORT 32-1086
17
INJECTION INTO LUNAR
TRANSFER ORBIT
SURVEYOR
SEPARATION
/ITLAS
SUSTAINER
PHASE
PROGRAMMED
PITCHOVER
(Z. + l5secT0
BECO)
BOOSTER -
PHASE
PROGRAMMED
ROLL
(A + 2 TO
L + I5sec)
SECO/VECO (SUSTAINER/
^^ VERNIER ENGINE CUTOFF)
<>/ ^JETTISON
NOSE FAIRING
JETTISON
INSULATION PANELS
BOOSTER
JETTISON
BECO (BOOSTER
ENGINE CUTOFF)
CENTAUR I SURVEYOR PRESEPARATION EVENTS
SAFE SURVEYOR DESTRUCT SYSTEM
SURVEYOR PRESEPARATION ARMING
EXTEND SURVEYOR LANDING LEGS
EXTEND SURVEYOR OMNIANTENNAS
TURN ON SURVEYOR TRANSMITTER
HIGH POWER
SEPARATE ELECTRICAL DISCONNECT
-LAUNCH FROM
AFETR COMPLEX 36A
Fig. III-3. Launch phase nominal events
18
JPt TECHN/CAl REPORT 32-7086
signal also was provided by the Centaur during this
period to enable the spacecraft to act on the presepara-
tion commands.
The Centaur commanded separation of the spacecraft
electrical disconnect 5.5 sec before spacecraft separation,
which was initiated at 752.6 sec. The Centaur attitude-
control engines were disabled for 5 sec during spacecraft
separation in order to minimize vehicle turning moments.
4. Centaur Retromaneuver
At 5 sec after spacecraft separation, the Centaur began
a turnaround maneuver using the attitude-control engines
to point the aft end of the stage in the direction of the
flight path. Four sec after the Centaur had reached the
mid-point of the turn and while continuing the turn, two
of the 50-lb-thrust hydrogen peroxide engines were fired
for a period of 20 sec to provide initial lateral separation
of the Centaur from the spacecraft. At 240 sec after sep-
aration, the propellant blowdown phase of the Centaur
retromaneuver was initiated by opening the hydrogen
and oxygen prestart valves. Oxygen was vented through
the engine nozzles while hydrogen discharged through
vent tubes. Propellant blowdown was terminated after
250 sec by closing the prestart valves. At the same time
(1242.9 sec) the Centaur power change-over switch was
energized to turn off all power except telemetry and
C-band beacon.
E. Performance
The Atlas/Centaur AC-7 vehicle performance was
near nominal, providing a very satisfactory powered
flight phase and accurate injection of the Surveyor II
spacecraft into the prescribed lunar transfer trajectory.
1. Guidance and Flight Control
Performance of the guidance system was excellent as
evidenced by the projected miss distance of the injected
spacecraft of only about 142 km from the prelaunch aim-
ing point. (Refer to Section VII for a presentation of
vehicle guidance accuracy results in terms of equivalent
midcourse velocity correction.)
The guidance system discrete commands (BECO, SECO
backup, and MECO) were issued well within system
tolerance. When guidance steering was enabled from
BECO + 8 until SECO and again from main engine igni-
tion (MEIG) + 4 sec until MECO, the initial attitude
errors (maximum 6 deg nose up and 3 deg nose right)
were quickly nulled, after which the vehicle was held
in close alignment with the commanded steering vector.
Autopilot performance was satisfactory throughout the
flight, with proper initiation of programmed events and
control of vehicle stability. Vehicle transients at liftoff
were similar to those occurring on previous Centaur
flights and were quickly damped following autopilot
activation at 42-in. motion. Vehicle disturbances during
the balance of the flight were at or below the expected
levels based on previous flights, except for a brief but
high roll transient occurring about 5 sec after booster
jettison and an unexplained high roll torque during the
first 8 sec of the Centaur turn-around maneuver.
The Centaur reaction control system apparently per-
formed properly, maintaining vehicle control throughout
the entire post-MECO and retromaneuver period, when
the system was active. Hydrogen peroxide engine duty
cycles averaged less than 2%.
2. Propulsion and Propellant Utilization
Both Atlas and Centaur propulsion systems operated
satisfactorily throughout the flight. As has occurred on
previous flights, the Centaur engines burned longer than
expected (about 3 sec), but this was a relatively small
deviation in relation to the allowable dispersion.
All vehicle propellant systems performed properly. The
Atlas propellant utilization (PU) system controlled pro-
pellants to effect nearly simultaneously depletion at
SECO with fuel depletion shutdown. This resulted in
minimum propellant residuals above the pump inlets.
The Centaur PU system also performed well, con-
trolling the calculated unbalance of propellants at MECO
to 33 lb of Hquid hydrogen. Comparing this to the pre-
dicted value of 15 lb residual hydrogen indicates a
Centaur PU system error of 18 lb excess hydrogen. The
Centaur "burnable" residuals were calculated to be
131 lb oxygen and 59 lb hydrogen, which could have
provided an additional burn time of about 2.3 sec at
normal engine flow rates until theoretical oxygen deple-
tion. The predicted values for burnable residuals were
205 lb oxygen and 56 lb hydrogen for a completely
nominal flight.
3. Pneumatics, Hydraulics, and Electrical Power
Pressure stability and regulation were satisfactory in
both the Atlas and Centaur hydraulic and pneumatic
JPL TECHNICAL REPORT 32-1086
19
circuits, and propellant tank pressures were maintained
within the required hmits to assure structural integrity.
Performance of vehicle electrical power systems, in-
cluding range safety power supphes, was normal through-
out the flight except for somewhat higher than expected
current demands on the Centaur electrical system com-
mencing with squib firing for spacecraft separation. Post-
flight investigation and simulation tests indicate that the
abnormal power demands were probably due to a faulty
thermal relay in the squib firing circuit. The thermal
relays are designed to remove power to the squibs after
squib firing because of the high probability that the
squibs will develop short circuits when fired. On this
flight, it is believed the leads to one of the thermal relays
were shorted because of a design defect.
4. Telemetry, Tracking, and Range Safety Command
In general, the Atlas and Centaur instrumentation and
telemetry systems functioned satisfactorily. However,
some instrumentation anomalies occurred, including fail-
ure of three of the five accelerometers located at the
spacecraft/adapter interface to provide vibration data.
Cause for the failure to obtain vibration output from
the three accelerometers is presently believed to be due
to faulty harness connections between the transducers
and amplifiers. Atlas signal dropout (for 138 millisec)
and Centaur signal degradation occurred as expected at
booster jettison. Some radio frequency interference (RFI)
was noted on a few of the Centaur telemetry channels
until nose fairing separation, when the RFI pattern pre-
sumably changed. Cause for the RFI is unknown since
protective shielding had been added to the telemetry
system design following detection of RFI susceptibility
during the AC-6 development flight prelaunch checkout.
Performance of the Atlas and Centaur range safety
command systems was satisfactory. At 9.9 sec after
MECO, a range safety command to disable the destruct
system was sent from Antigua and properly executed.
the five accelerometers installed in the vicinity of the
Centaur/Surveyor interface. The only high-frequency
accelerometer on a continuous telemetry channel was
among the three from which useful telemetry output wa%
not obtained. The only unusual oscillation observed was
a high roll rate of 4.3 deg/sec (peak to peak) which
occurred 5.3 sec after booster jettison (0.4 sec after Centaur
guidance enable). (See Section IV-A for a discussion of
spacecraft launch phase vibration environment.)
Special transducers were installed at the base of the
Atlas stage on this flight and the previous flight of AC-8
to obtain data for improvement of the base pressure
model which is used for trajectory design. Data from this
flight indicated generally lower base pressure levels than
obtained for the AC-8 flight.
The Surveyor compartment pressure dropped in a
normal manner from atmospheric to essentially zero at
L + 107 sec.
6. Separation and Retro Maneuver Systems
All vehicle separation systems fimctioned normally,
although the magnitude of high-frequency transients
associated with shaped charge firing is not well deter-
mined owing to missing accelerometer data.
Booster section jettison occurred as planned under
sustainer engine pitch and yaw attitude control. Rela-
tively high roll and yaw transient rates occurred as a
result of this event.
Proper separation of all four insulation panels was con-
firmed by indications received from four breakwires, one
of which was attached to each panel near a hinge arm.
Only low transient rates were imparted to the vehicle as
a result of this event, providing further evidence of the
improvement afforded by the redesigned panel hinges
first flown on AC-10.
5. Vehicle Loads and Environment
All vehicle loads were within expected ranges. Aero-
dynamic bending loads were well within vehicle capa-
bility and less than the predicted values, based upon
wind sounding data obtained two hours before launch.
In general, the vibration profile of the AC-7 vehicle
appeared to be similar to that of preceding Centaur
flights. However, the AC-7 vibration environment cannot
be as well established because of the failure of three of
Normal separation of the nose-fairing was verified by
indications from disconnect wires which were utilized
for the first time on this flight. These were incorporated
in the pullaway electrical connectors of each fairing half.
As expected, there was no indication of pressure buildup
in the spacecraft compartment at nose-fairing thruster
bottle actuation.
Atlas/Centaur separation occurred as planned. Telem-
etry data indicated all eight Atlas retro rockets fired. The
20
JPL TECHNICAL REPORT 32-?086
pitch rate gyros indicated no vehicle rotation in the
critical pitch plane during this separation event.
* At spacecraft separation, all three pyrotechnic release
latches actuated within 1 millisec of each other. Data
from the extensometers indicated that the three spring
assemblies extended normally, producing a spacecraft
separation rate of about 1 ft/sec. Residual Centaur rota-
tion at separation was 0.19 deg/sec as determined from
Centaur gyro data. No angular motion between the space-
craft and Centaur could be detected from the three
extensometers traces, which show that spring stroke vs
time was identical for the springs.
The Centaur retromaneuver was executed as planned.
Five hours after spacecraft separation, the distance be-
tween the Centaur and spacecraft had increased to
730 km, which is well in excess of the required minimum
separation of 336 km at that time. The Centaur closest
approach to the moon was 5,675 km.
JPL TECHNICAL REPORT 32-1086
21
IV. Surveyor Spacecraft
The Surveyor 11 spacecraft was to have flown a flight
profile quite similar to that flown on Surveyor I. The
Surveyor 11 primary flight objectives were (1) to soft-land
on the lunar surface at a site east of the Surveyor I land-
ing point (0.00 deg latitude and 0.67 deg west longitude
in Sinus Medii), (2) to demonstrate the capability of the
spacecraft to land with an oblique approach angle not
greater than 25 deg (predicted approach angle was 2.'3 deg
from the vertical), (3) to transmit post-landing television
pictures, and (4) to obtain touchdown dynamics, radar
reflectivity, and lunar surface thermal data. To these
ends, the Surveyor 11 spacecraft performed the early
phases of the mission up to midcourse as planned. During
execution of the midcourse maneuver, one vernier engine
failed to fire, resulting in a spacecraft tumbling condition
which prevented attainment of planned mission objectives.
The third vernier engine also failed to ignite during each
of 39 post-midcourse attempts to fire the vernier engines.
The Surveyor 11 spacecraft failure was thoroughly in-
vestigated by a formally appointed Failure Review Board.
Although this investigation did not disclose a specific
cause for the failure, many recommendations have been
made by the Board to provide on future missions a greater
assurance of spacecraft flight readiness and better pre-
flight and in-flight diagnostic data.
A. Spacecraft System
In the Surveyor spacecraft design, the primary
objective was to maximize the probability of successful
spacecraft operation within the basic limitations im-
posed by launch vehicle capabilities, the extent of knowl-
edge of transit and lunar environments, and the current
technological state of the art. In keeping with this pri-
mary objective, design policies were established which
(1) minimized spacecraft complexity by placing responsi-
bility for mission control and decision-making on earth-
based equipment wherever possible, (2) provided the
capability of transmitting a large number of different
data channels from the spacecraft, (3) included provisions
for accommodating a large number of individual com-
mands from the earth, and (4) made all subsystems as
autonomous as practicable.
Figure IV-1 illustrates the Surveyor spacecraft in the
cruise mode and identifies many of the major components.
A simphfied functional block diagram of the spacecraft
system is shown in Figure IV-2. The spacecraft design
is discussed briefly in this section and in greater detail
in the subsystem sections which follow. A detailed con-
figuration drawing of the spacecraft is contained in
Appendix B. The configuration of the Surveyor space-
craft is dictated by the selection of a tripod landing gear
with three foldable landing legs for the soft landing.
1. Spacecraft Coordinate System
The spacecraft coordinate system (Fig. IV-3) is an
orthogonal, right-hand Cartesian system. Figure IV-4
shows the spacecraft motion about its coordinate axes
relative to the celestial references. The cone angle of the
earth is the angle between the sim vector and the earth
vector as seen from the spacecraft. The clock angle of
the earth is measured in a plane perpendicular to the
sun vector from the projection of the star Canopus vector
to the projection of the earth vector in the plane. The
JPL TECHNICAL REPORT 32-1086
23
SOLAR PANEL
OMNIANTENNA
HIGH-GAIN
ANTENNA
STAR CANOPUS
SENSOR
THERMALLY "^
CONTROLLED 4
COMPARTMENT-^'
RADAR ALTITUOE-
DOPPLER VELOCITY
ANTENNA
VERNIER ENGINE
OMNIANTENNA
VERNIER PROPELLANT
PRESSURIZING GAS
(HELIUM) TANK
AUXILIARY BATTERY
ATTITUDE CONTROL GAS
(NITROGEN) TANK
RETRO ROCKET MOTOR
LANDING GEAR
ALTITUDE MARKING
RADAR ANTENNA
Fig. IV-1. Surveyor // spacecraft in cruise mode
spacecraft coordinate system may l)e related to tlie cone
and tlie clock angle coordinate system, provided sun
and Canopus lock-on lias l)een achic\-ed. In tins case the
spacecraft minus Z-axis is directed toxsard the sun, and
the minus X-axis is coincident with the projection of the
Canopus vector in the plane perpendicular to the direc-
tion of the sun.
2. Spacecraft Mass Properties
Purveyor II weighed 2203.67 lb at launch, with a final
predicted touchdown weight of 644.07 lb nominal, (x-nter
of gravity of the vehicle is kept low to obtain stability
over a wide range of landing conditions. C'enfer-of-gravity
limits after Stiivcyor/Ccutaitr separation for mideourse
and retro maneuvers are constrained by the attitude cor-
rection cai:)al)ilities of the flight control and vernier
engine sulisystems dining retrorocket burning. Limits of
travel of the vertical center of gravity in tlie touchdown
configuration are designed to lauding site assumptions
and approach angli> reijuirements so that the spacecraft
will not topple when landing.
3. Structures and Mechanisms
The structures and mechanisms subsystem providi'S
basic structural sujiiiort (including touchdown stabiliza-
tion), mechanical actuation, thermal protection, and
electronic packaging and cabling. A tubular aluminum
spaceframe is utilized for basic structural support. Three
landing leg assemblies and crushable blocks for kmar
landing are attached to the spaceframe. Other mech-
anisms provided are the high-gain antenna and solar
panel positioner (A/SPP), two omniantenna mechanisms,
a separation sensing and arming device, the secondary
sun sensor, and pyrotechnic devices. Two compartments
incorporating special insulation and thermal switches are
provided for thermal protection of critical spacecraft
components.
4. Thermal Control
Thermal control of ecjuipment over the extreme tem-
perature range of the limar surface ( i 260 to 260°F)
is accomplished by a combination of passive, semipas-
sive, and active methods including the use of heaters
controlled by ground command. The design represents
the latest state of the art in the application of structural
and thermal design {principles to lightweight spacecraft.
Units that require critical thermal control are the ap-
jiroach and survey television cameras, altitude marking
radar, Compartments A and B, and vernier engine pro-
pellant tanks and lines.
5. Electrical Power
'I'lie electrical pouer subsystem is designed to gener-
ate, store, convert, and distribute electrical energy. A
single solar panel is utilized which is capable of gen-
erating continuous unregulated power at 90 to 55 w, de-
pending upon environmental temperature and incidence
angle of solar radiation. Peak imregulated power capa-
bility is limited to f 000 w by the two spacecraft batteries
(main and auxiliary). Thi' initial energy storage ol the
subsystem is 4400w-hr. Only one battery, the main battery,
can be recharged, to an energy storage of 3520 w-hr.
The batteries determine the unregulated power voltage
and are designed to sustain a voltage between f7.5 and
27.5 V, with a nominal value of 22 v. The unregulated
power is distributed to the loads via an unregulated bus.
liegulatcd power is provided by a boost regulator at
29 V, controlled to f% for the flight control and "non-
essential" loads and to 2% for the "essential" loads. The
maximum regulated power capabilit)' of the boost regu-
lator is 270 w.
6. Propulsion
The j^rojMilsion subsystem supplies thrust force during
the mideourse correction and terminal descent phases of
24
JPL TECHNICAL REPORT 32-1086
E
a
o
a
c
o
o
a
13
0)
a.
E
«75
I
>
d>
JPL TECHNICAL REPORT 32- J 086
25
Fig. IV-3. Spacecraft coordinate system
26
JPL TECHNICAL REPORT 32-1086
SUN
-PROJECTION OF CANOPUS
STAR VECTOR IN A PLANE
PERPENDICULAR TO
DIRECTION OF SUN
CONE ANGLE
EARTH
0°< CONES 180"
0° < CLOCK < 360°
PROJECTION OF EARTH
VECTOR IN A PLANE
PERPENDICULAR TO
DIRECTION OF SUN
Fig. IV-4. Spacecraft coordinates relative to
celestial references
the mission. The propulsion subsystem, consisting of a
bipropellant vernier engine system and a soHd-propellant
main retrorocket motor, is controlled by the flight control
system through preprogrammed maneuvers, commands
from earth, and maneuvers initiated by flight control
sensor signals.
The three thrust chambers of the vernier engine sub-
system supply the thrust forces for midcourse maneuver
velocity vector correction, attitude control during main
retrorocket burning, and velocity vector and attitude
control during terminal descent. The thrust of each ver-
nier engine can be throttled over a range of 30 to 104 lb.
The main retrorocket, which performs the major por-
tion of the deceleration of the spacecraft during lunar
landing maneuver, is a spherical, solid-propellant motor
with partially-submerged nozzle to minimize overall
length.
7. Flight Control
The purpose of the flight control subsystem is to con-
trol spacecraft flight parameters throughout the transit
portion of the mission. Flight control uses three forms of
reference to perform its function. These are celestial
sensors, inertial sensors, and radar sensors. The outputs
of each of these sensors are utilized by analog electronics
to create thrust commands for operation of attitude gas
jets and the spacecraft vernier and main retro propulsion
systems. Flight control requires ground commands for
initiation of various sequences and performance of
"manual" operations. Flight control programming initi-
ates and controls other sequences.
The celestial sensors allow the spacecraft to be locked
to a specific orientation defined by the vectors to the sun
and the star Canopus and the angle between them. Initial
search and acquisition of the sun is accomplished by the
secondary sun sensor. The primary sun sensor then main-
tains the orientation with the sun line.
Of the inertial sensors, integrating gyros are used to
maintain spacecraft orientation inertially when the celes-
tial references are not available. Accelerometers measure
the thrust levels of the spacecraft propulsion systems dur-
ing midcourse correction and terminal descent phases.
The attitude gas jets are cold gas (nitrogen) reaction
devices for control of the orientation of spacecraft atti-
tude in all three axes during coast phases of the flight.
They are installed in opposing pairs near the ends of the
three landing legs. The three vernier engines provide
thrust, which can be varied over a wide range, for mid-
course correction of the spacecraft velocity vector and
controlled descent to the lunar surface. A roll actuator
tilts the thrust axis of Vernier Engine 1 away from the
spacecraft roll axis for attitude and roll control during
thrust phases of flight when the attitude gas jets are not
effective. The main retro motor is utilized to remove the
major portion of the spacecraft approach velocity during
terminal descent.
8. Radar
Two radar systems are employed by the Surveyor
spacecraft. An altitude marking radar (AMR) provides a
mark signal to initiate the main retro sequence. In addition,
a radar altimeter and doppler velocity sensor (RADVS)
functions in the flight control subsystem to provide three-
axis velocity, range, and altitude mark signals for flight
control during the main retro and vernier phases of ter-
minal descent. The RADVS consists of a doppler velocity
sensor, which computes velocity along each of the space-
craft X, Y, and Z axes, and a radar altimeter, which com-
putes slant range from 50,000 ft to 14 ft and generates
1000-ft mark and 14-ft mark signals.
JPL TECHNICAL REPORT 32-1086
27
Table IV-1 . Content of telemetry signals from spacecraft
Data
mode
Commutator
Mode 1
Commutator
Mode 2
Commutator
Mode 3
Commutator
Mode 4
Cruise phase
commutator
Mode 5
Thrust phase
commutator
Mode 6
TV commutator
Mode 7
Vibration data
Shock absorber
data
Gyro speed
Source
Flight control,
propulsion
Flight control,
propulsion,
approach TV, AMR,
RADVS
Inerttol guidance,
approach TV, AMR,
RADVS, vernier
engines
Temperatures,
power status,
telecommunications
Flight control,
power status,
temperature
Flight control,
power status, AMR,
RADVS, vernier
engine conditions
TV survey camera
Launch phase
occelerometers
Post-DSIF
acquisition phase
occelerometers
Strain gages
Inertial guidance
unit
Significance
Provides data required for
midcourse maneuver and
preretro terminal maneuver
Provides data required for
retro descent
Provides data required for
vernier descent
Provides data required for
miscellaneous transit and
lunar surface operations
Provides data required during
cruise mode to determine
general spacecraft status
Provides data required for
backup of Modes 1, 2, and 3
during thrusting maneuvers
Provides frame identification
while survey TV is operating
Indicates vibration during
launch phase
Designed to indicate vibration
from main retro motor and
vernier engine firings and
mechanical shock during
landing
Measures strain on landing
gear due to landing shock
Indicates angular rate of
gyro spin motors
Number of
points
sampled
100
100
50
100
120
120
16
Form
Digital
Digital
Digital
Digital
Digital
Digital
Digital
Analog
Analog
Comments
Analog
Analog
Modes 1, 2, 3, and 4
used one at a time
on command per
Standard Sequence
of Events (SSE)
Used on command per
SSE
Used on command per
SSE
Frame ID alternates
with analog video
signals
Transmitted over Centaur
telemetry link only
Installed but not used on
Surveyors I and II
Samples are pitch, roll,
and yaw axes on
command per SSE
28
JPL TECHNICAL REPORT 32-1086
9. Telecommunications
The spacecraft telecommunications subsystem provides
■* for (1) receiving and processing commands from earth,
(2) providing angle tracking and one- or two-way doppler
data for orbit determination, and (3) processing and trans-
mitting spacecraft telemetry data.
Continuous command capability is assured by two
identical receivers which remain on throughout the life
of the spacecraft and operate in conjunction with two
omniantennas and two command decoders through switch-
ing logics.
Operation of a receiver in conjunction with a trans-
mitter through a transponder interconnection provides a
phase-coherent system for doppler tracking of the space-
craft during transit and after touchdown. Two identical
transponder interconnections (Receiver/Transponder A
and Receiver/Transponder B) are provided for redun-
dancy. Transmitter B with Receiver/Transponder B is
the transponder system normally operated during transit.
Data signals from transducers located throughout the
spacecraft are received and prepared for telemetry trans-
mission by signal processing equipment which performs
commutation, analog-to-digital conversion, and pulse-
code and amplitude-to-frequency modulation functions.
Most of the data signals are divided into six groups
("commutator modes") for commutation by two commu-
tators located within the telecommunications signal proc-
essor. (An additional commutator is located within the
television auxiliary for processing television frame identi-
fication data.) The content of each commutator mode has
been selected to provide essential data during particular
phases of the mission (Table IV-1). Other signals, such as
strain gage data which is required continuously over brief
intervals, are applied directly to subcarrier oscillators.
Summing amplifiers are used to combine the output of
any one commutator mode with continuous data. The
composite signal from the signal processor, or television
data from the television auxiliary, is sent over one of the
two spacecraft transmitters. The commutators can be
operated at five different rates (4400, 1100, 550, 137.5,
and 17.2 bits/sec) and the transmitters at two different
power levels (10 w or 100 mw). In addition, switching
permits each of the transmitters to be operated with any
one of the three spacecraft antennas (two omniantennas
and a planar array) at either the high or low power level.
Selection of data mode(s), data rate, transmitter power,
and transmitter-antenna combination is made by ground
command. A data rate is selected for each mission phase
which will provide sufficient signal strength at the DSIF
station to maintain the telemetry error rate within satis-
factory limits. The high-gain antenna (planar array) is
utilized for efficient transmission of video data. The
Surveyor II data mode/rate profile is shown in Fig. IV-5.
10. Television
The Surveyor II television subsystem included a down-
ward pointing camera for terminal descent photographs,
a survey camera for photographs from the lunar surface,
and a television auxiliary for final decoding of commands
and processing of video and frame identification data for
transmission by either of the spacecraft transmitters. The
standard sequence of events (SSE) did not call for oper-
ation of the approach camera on the Surveyor II mission
because it was desired to minimize spacecraft operational
requirements during the complex and critical terminal
descent phase. The survey camera is designed for post-
landing operation to provide photographs of the lunar
surface panorama, portions of the spacecraft, and the
lunar sky. Photographs may be obtained in either of two
modes: a 200-line mode for relatively slow transmission
over an omniantenna or a 600-line mode for more efficient
transmission over the planar array.
11. Instrumentation
Transducers are located throughout the spacecraft sys-
tem to provide signals that are relayed to the DSIF
stations by the telecommunication subsystem. These sig-
nals are used primarily to assess the condition and per-
formance of the spacecraft. Some of the measurements
also provide data useful in deriving knowledge of certain
characteristics of the lunar surface.
In most cases the individual subsystems provide the
transducers and basic signal conditioning required for
data related to their equipment. All the instrumentation
signals provided for the Surveyor II spacecraft are sum-
marized by category and responsible subsystem in
Table IV-2.
All of the temperature transducers are resistance-type
units except for three microdiode bridge amplifier assem-
blies used in the television subsystem.
The voltage (signals) and position (electronic switches)
measurements consist largely of signals from the com-
mand and control circuits.
JPt TECHNICAL REPORT 32-1086
29
Table IV-2. Spacecraft instrumentation
Structures,
mechanisms, and
thermal control
Electrical
power
Propulsion
Flight
control
Radar
Telecommunication
Television
Total
Temperature
33
5
16
9
7
2
6
78
Pressure
2
2
4
Position (potentiometers)
7
6
13
Position (mechanical switches)
10
3
13
Position (electrical switches)
31
14
9
6
60
Current
18
2
20
Voltage (power)
6
2
1
9
Voltage (signals)
14
6
8
28
Strain gages
3
3
6
Accelerometers
8
1
9
Inertiol sensors (gyro speed)
3
3
RF power
2
2
Optical
9
9
A strain gage is mounted on each of the vernier engine
brackets to measure thrust and on each of the three land-
ing leg shock absorbers to monitor touchdown dynamics.
The flight control accelerometer is mounted on the
retro motor case to verify motor ignition and provide
gross retro performance data. Of the remaining eight
accelerometers, four are designed to provide data on the
vibration environment during launch phase and four are
designed to provide data on the dynamic response of
spacecraft elements to flight events which occur after
spacecraft separation. Only data from the retro motor
accelerometer and launch phase accelerometers has been
telemetered on Surveyor missions to date.
Additional discussion of instrumentation is included
with the individual subsystem descriptions.
12. Terminal Maneuver and Descent Phase Design
The system design for automatic terminal descent,
which has been developed and used for the first time in
the Surveyor program, is described here to illustrate the
critical functions required to be performed by several of
the subsystems.
a. Terminal descent sequence. The terminal phase be-
gins with the preretro attitude maneuvers (Fig. IV-6).
These maneuvers are commanded from earth to reposi-
tion the attitude of the spacecraft from the coast phase
sun-star reference such that the expected direction of
the retro thrust vector will be aligned with respect to the
spacecraft velocity vector. Following completion of the
attitude maneuvers, the AMR is activated. It has been
preset to generate a mark signal when the slant range to
the lunar surface is 60 miles nominal. A backup mark
signal, delayed a short interval after the AMR mark
should occur, is transmitted to the spacecraft to initiate
the automatic sequence in the event the AMR mark is
not generated. A delay between the altitude mark and
main retro motor ignition has been preset in the flight
control programmer by ground command. Vernier engine
ignition is automatically initiated 1.1 sec prior to main
retro ignition.
During the main retro phase, spacecraft attitude is
maintained in the inertial direction established at the
end of the preretro maneuvers by differential throttle
control of the vernier engines while maintaining the total
vernier thrust at the midthrust level. The main retro
burns at essentially constant thrust for about 40 sec, after
which the thrust starts to decay. This tailoff is detected
by an inertial switch which increases vernier thrust to
the high level and initiates a programmed time delay of
about 12 sec, after which the main retro motor case is
30
iPL TECHNICAL REPORT 32-1086
COMMUTATOR
MODE
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1100-
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137.5-
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JPl TECHNICAL REPORT 32-1086
3/
s
o
s
X
o
K
Q.
q:
I
1-
a.
a.
^ "J^ <r ^ "J- ■«■
Fig. IV-5. Surveyor // data mode/rate profile
St,
CRUISE ATTITUDE
PRERETRO MANEUVER 30 min
BEFORE TOUCHDOWN ALIGNS
MAIN RETRO WITH FLIGHT PATH
NOTE: ALTITUDES, VELOCITIES, AND
TIMES ARE NOMINAL
MAIN RETRO START BY ALTITUDE-
MARKING RADAR WHICH EJECTS
FROM NOZZLE; CRAFT STABILIZED
BY VERNIER ENGINES AT
60-mi ALTITUDE, 6,100 mph
MAIN RETRO BURNOUT AND EJECTION;
VERNIER RETRO SYSTEM TAKEOVER AT
25,000 ft, 240 mph
VERNIER ENGINES SHUTOFF
AT 13 ft, 3-1/2 mph
TOUCHDOWN AT 10 mph
"''y
Fig. IV-6. Terminal descent nominal events
JPL TECHNICAL REPORT 32-1086
33
ejected. The main retro phase removes more than 95%
of the spacecraft velocity and puts the spacecraft posi-
tion, velocity, and attitude relative to the lunar surface
within the capability of the final, vernier phase.
The vernier phase generally begins at altitudes be-
tween 10,000 and 50,000 ft and velocities in the range of
100 to 700 ft/sec. This wide range of vernier-phase initial
conditions exist because of statistical variations in param-
eters which affect main retro burnout. About 2 sec after
separation of the main retro case, vernier thrust is re-
duced and controlled to produce a constant spacecraft
deceleration of 0.9 lunar g, as sensed by an axially
oriented accelerometer. The spacecraft attitude is held
in the preretro position until the doppler velocity sensor
locks onto the lunar surface (Fig. lV-7). The thrust axis
is then aligned and maintained to the spacecraft velocity
vector throughout the remainder of the descent until the
terminal sequence is initiated (when the attitude is again
held inertially fixed). With the thrust axis maintained in
alignment with the velocity vector, the spacecraft makes
BEAM 3
a "gravity turn," wherein gravity tends to force the flight"
path towards the vertical as the spacecraft decelerates.
The vehicle descends at 0.9 lunar g until the radar?
sense that the "descent contour" has been reached
(Fig. IV-8). This contour corresponds, in the vertical case,
to descent at a constant deceleration. The vernier thrust
is commanded such that the vehicle follows the descent
contour until shortly before touchdown, when the ter-
minal sequence is initiated. Nominally, the terminal
sequence consists of a constant-velocity descent from
40 to 1.3 ft at 5 ft/sec, followed by a free fall from 13 ft,
resulting in touchdown at approximately 13 ft/sec.
b. Terminal descent design constraints. Constraints on
the allowable main retro motor burnout conditions are
of major importance in Surveyor terminal descent design.
60
50
40
O
X
Q
ID
[— - --] T I I
NOMINAL BURNOUT LOCUS-
MAXIMUM IMPACT
VELOCITY l^
NO M/C CORRECTION
DOPPLER LIMIT
30
20
10
Fig. IV-7. RADVS beam orientation
100 200 300 400 500 600 700 800 900
VELOCITY, fl/sec
Fig. IV-8. Altitude velocity diagram
34
JPl TECHNICAL REPORT 32-1086
RADVS operational limitations contribute to constraints
on the main retro burnout conditions. Linear operation
of the doppler velocity sensor is expected for slant ranges
"below 50,000 ft and for velocities below 700 ft/sec. The
altimeter limit is between 30,000 and 40,000 ft, depend-
ing on velocity. These constraints are illustrated in the
range-velocity plane of Fig. IV-8.
The allowable main retro burnout region is further
restricted by the maximum thrust capability of the ver-
nier engine system. To accurately control the final de-
scent, the minimum thrust must be less than the least
possible landed weight (lunar gravity) of the vehicle. The
result is a minimum thrust of 90 lb. This in turn con-
strains the maximum vernier thrust to 312 lb because of
the limited range of throttle control which is possible.
Descent at the maximum thrust to touchdown defines
a curve in the range-velocity plane below which main
retro burnout cannot be allowed to occur. Actually, since
the vernier engines are also used for attitude stabilization
by differential thrust control, it is necessary to allow some
margin from the maximum thrust level. Furthermore,
since it is more convenient to sense deceleration than
thrust, the vernier phase of terminal descent is performed
at nearly constant deceleration rather than at constant
thrust. Therefore, maximum thrust will be utilized only
at the start of the vernier phase.
The maximum vernier phase deceleration defines a
parabola in the altitude-velocity plane. For vertical de-
scents at least, this curve defines the minimum altitude at
which main retro burnout is permitted to occur with a
resulting soft landing. This parabola is indicated in
Fig. IV-8. (For ease of spacecraft mechanization, the
parabola is approximated by a descent contour consisting
of straight-line segments.)
Main retro burnout must occur sufficiently above the
descent contour to allow time to align the thrust axis
with the velocity vector before the trajectory intersects
the contour. Thus, a "nominal burnout locus" (also shown
in Fig. IV-8) is established which allows for altitude dis-
persions plus an alignment time which depends on the
maximum angle between the flight path and roll axis at
burnout.
The allowable burnout region having been defined, the
size of the main retro motor and ignition altitude are
determined such that burnout will occur within that
region.
In order to establish the maximum propellant require-
ments for the vernier system, it is necessary to consider
dispersions in main retro burnout conditions as well as
midcourse maneuver fuel expenditures. The principal
sources of main retro burnout velocity dispersion are the
imperfect alignment of the vehicle prior to main retro
ignition and the variability of the total impulse. In the
case of a vertical descent, these variations cause disper-
sions of the type shown in Fig. IV-8, where the ellipse
defines a region within which burnout will occur with
probability 0.99. The design chosen provides enough fuel
so that, given a maximum midcourse correction, the
probability of not running out is at least 0.99.
The spacecraft landing gear is designed to withstand
a horizontal component of the landing velocity. The hori-
zontal component of the landing velocity is nominally
zero. However, dispersions arise primarily because of the
following two factors:
(1) Measurement error in the doppler system resulting
in a velocity error normal to the thrust axis.
(2) Nonvertical attitude due to: (a) termination of the
"gravity turn" at a finite velocity, and (b) attitude
control system noise sources.
Since the attitude at the beginning of the constant-
velocity descent is inertially held until vernier engine
cutoff, these errors give rise to a significant lateral ve-
locity at touchdown.
13. Design Changes
Table IV-3 presents a summary of notable differences
in design between the Surveyors I and //.
14. Spacecraft Reliobih'ty
The prelaunch reliability estimate for the Surveyor II
spacecraft was 0.66 for the flight and landing mission,
assuming successful injection. The estimate was based
on systems test data. Owing to the number of unit
changes on the spacecraft, the reliability estimate is con-
sidered generic to Surveyor II rather than descriptive of
the exact Surveyor II spacecraft configuration. Figure
IV-9 shows the history of reliability estimates for Surveyor
II during its system test phases. The detail reliability esti-
mates for flight and landing are listed in Table IV-4. For
comparative purposes, Surveyor I estimates are also
shown.
JPL TECHNICAL REPORT 32-1086
35
Table IV-3. Notable differences between Surveyors / and II
Hem
Boost regulator (BR) overload trip circuit (OTC)
Auto solar panel deploy logic enable
Filter chokes on input to signal processing
equipment, and filter on A/D Converter No. 2
nulling amplifier in the command signal
processor
Telemetry of flight control (FC) return signal
A/SPP pin pullers
A/SPP drive motors
Omniantenna latch and release mechanism
Command assignments
Boost regulator FC regulator filter
Velocity components V, and V„ gains in flight
control sensor group (FCSG)
Auxiliary battery paint pattern
Solder splash in signal processing equipment
RADVS sidelobe rejection logic
Conopus sensor sun reference filter change
Canopus sensor window
A/SPP pulse duration
Description
In Surveyor I the OTC in the BR was disabled because it would trip during normal operation.
The Surveyor // BR has a redesigned OTC which does not trip during normal operation
In Surveyor / the auto solar panel deploy logic was "enabled" by command prior to launch.
In Surveyor II a diode was added in the harness to "enable" the auto solar panel deploy
logic with the some signal from Centour which causes the transmitter to switch to high power.
(Solar panel deployment in both coses is initiated at separation.)
Both of these design improvements were to eliminate the large variation in temperature
readouts that were present on Surveyor / telemetry
In Surveyor // the FC return signal is telemetered so that the varying harness voltage drops
con be accounted for to provide more accurate data on such signals os range and velocfty
The A/SPP pin puller modules were redesigned to simplify installation at AFETR
All of the Surveyor // drive motors on the A/SPP have roller detents instead of the boll detents
used in oil but the roll axis on Surveyor I. This is a design improvement
The Surveyor II release mechanisms for Omniantennas A and B were redesigned to prevent the
deployment problem which occurred in flight on Surveyor /. The clevis opening was broadened
and o kickout spring was added
The engineering mechanism auxiliary (EMA) on Surveyor II was modified to double up on the
functions of two of the commands so that two commond channels were available for fuel and
oxidizer dump
The Surveyor II boost regulator has a new filter on the FC regulator to eliminate the
oscillations which would sometimes occur and cause on overload on the shunt regulator
The V, and V„ radar attitude loop gains were reduced in Surveyor II to eliminate a potential
instability problem at velocities greater than 535 fps
Surveyor I auxiliary battery experienced low temperature
All units were modified to eliminate a solder splash problem (except the spare central
command decoder)
Two resistors in the signal data converter were removed in order to lower the point at which
the sidelobe signals ore rejected from 28 to 25 db
Surveyor / hod a Canopus sensor sun filter with a reduction of 50% (filter factor of 1.5) to
compensate for any possible fogging of Canopus sensor window, in accordance with recent
measurements of Canopus brightness at Tucson. For Surveyor // the filter factor was reduced
from 1.5 to 1.2 because the fogging problem did not materialize at the Canopus sensor
temperature of 79° F for the Surveyor / flight
The O-rings on the Canopus sensor window were chonged in on effort to prevent possible
fogging of the Canopus sensor filter
The battery chorge-regulotor was changed to reduce the A/SPP stepping current pulse duration
from 65 to 40 millisec. This change reduced the power dissipation in the battery chorge
regulator and in the A/SPP drive motors
36
JPL TECHNICAL REPORT 32-J086
Table IV-5. Spacecraft anomalies
200 400 600 800 1000 1200 1400
SURVEYOR SPACECRAFT SYSTEMS EXPERIENCE, hr
Fig. IV-9. Surve/or // reliability estimates
Table IV-4. %\ttyiByor spacecraft reliability
(flight and landing)
Subsystem
Surveyor f
Surveyor //
Telecommunications
0.922
0.941
Vehicle mechanisms
0.854
0.865
Propulsion
0.991
0.991
Electrical power
0.866
0.938
Flight controls
0.954
0.930
Subsystems net
0.645
0.704
System interoction
0.788
0.930
reliability factor
Spacecraft reliability
(0.645) (0.788) = 0.51
(0.704) (0.930) = 0.66
The primary source of data for reliability estimates is
the time and cycle information experienced by Surveyor II
units during systems tests. Data from Surveyor I test and
flight experience was included where there were no sig-
nificant design differences between the units. In general, a
failure is considered relevant if it could occur during
a mission. Relevance of failures is based on a joint
reliabihty /systems engineering decision.
15. Spacecraft System Performance
The Surveyor II spacecraft system performed well dur-
ing the mission until initiation of vernier engine thrusting
for the midcourse velocity correction. Table IV-5 pro-
vides a summary listing of spacecraft anomalies including
the midcourse correction failure. None of the anomalies
which occurred prior to the midcourse maneuver had a
significant effect on the mission; however, failure of
Vernier Engine 3 to provide midcourse thrust resulted
in failure of the mission. (See Mission Operations Chron-
Anomaly
Effect on mission
1. Two launch phase accelerometer
None. Telemetry data from
channels did not function
third accelerometer, located
properly
on the Centaur side of the
separation plane, was also
abnormal
2. The flight control function
None. The flight control was
reverted to inertial mode from
automatically placed back
rale mode 35 sec prior to
into the rate mode at
separation of the spacecraft
separation (normal
from the Centaur
operation)
3. Vernier Line 2 heater was full on
None
4. During Canopus star mapping,
None. Intentional high-gain
star-lock signal was not
setting of Canopus sensor.
observed when Conopus was
to compensate for possible
in the field of view
window fogging, removed
the capability of automatic
star lock-on. Manual
lock-on was executed
successfully
5. Receiver B signal strength was
None. Subsequent inflight
observed to be lower than
calibration of this telemetry
predicted after service tower
channel indicated that the
removal during countdown
premission calibration of
and during star verification
signal strength vs the AGC
prior to Canopus acquisition
reading was in error
6. Helium transducer pressure
None
indicated approximately a
500-psi zero shift in reading
7. Vernier Engine 3 did not
Caused the spacecraft to
respond properly to the
tumble, preventing
vernier engine ignition
completion of a standard
command for midcourse
mission
correction. Subsequent
attempts to obtain normal
Engine 3 thrusting were all
unsuccessful
ology (Section VI-C) for a description of spacecraft flight
events.)
During the boost phase of flight, the Surveyor space-
craft is subjected to a variable vibration environment
consisting of acoustically induced random vibration and
the transient response to discrete flight events. The
Surveyor II space vehicle was instrumented with five
accelerometers in order to obtain information on this
vibration environment. The location and orientation of
these accelerometers in the launch vehicle/spacecraft
JPL TECHNICAL REPORT 32-1086
37
-X
COMPARTMENT B
RETRO MOTOR
CY780
-Y
COMPARTMENT A
TRANSDUCER
CY520
CY530
CY540
CY770
CY780
LOCATION
SPACECRAFT, NEAR ADAPTER
ATTACH POINT I
SPACECRAFT, NEAR ADAPTER
ATTACH POINT 2
SPACECRAFT, NEAR ADAPTER
ATTACH POINT 3
ADAPTER, NEAR SPACECRAFT
ATTACH POINT I
SPACECRAFT, IN FCSG
RANGE, g
±10
±10
±10
±10
±10
FREQUENCY
RANGE, cps
2-2500
2-1260
2-1260
2-1260
2-1260
REMARKS
CONTINUOUS
COMMUTATED
COM MUTATED
COMMUTATED
COMMUTATED
Fig. IV-10. Launch-phase accelerometer location
interface area are shown in Fig. IV-10. The Z-axis accel-
erometer CY520 output was telemetered continuously;
the other four accelerometers were telemetered on a
commutated channel.
Improper output was received from three (CY520,
CY530, and CY770) of the five transducers (see also
Section III for discussion of this anomaly). At this time
it is believed that failure to obtain vibration output
from the three accelerometers was due to faulty har-
nesses between the transducers and amplifiers. Output
from the two transducers from which normal data was
received was commutated with the output from two of
the anomalous accelerometers. Therefore, valid data was
provided for less than half the flight time and most of
the transients were not recorded. In addition, since the
two accelerometers from which good data was received
were both in the Z-direction, no lateral-axis vibration
data was obtained. Typical data is shown in Table IV-6
for those events which were recorded and may be com-
pared with the Surveyor I flight data. The 95 percentile
(approximately 2a high) estimate of the vibration power
spectral density (over the frequency bandwidth 100-1500
cps) at Surveyor II liftoff is 0.011 gVcps. The correspond-
ing specification value is 0.0145 g-'/cps.
Shortly before spacecraft separation, a minor anomaly
occurred when the flight control subsystem switched from
rate to inertial mode. However, at separation from the
Centaur, the flight control subsystem was automatically
returned to the rate mode.
Star verification and acquisition sequence was nominal,
except that it was necessary to achieve Canopus lock-on
38
JPL TECHNICAL REPORT 32-1086
Table IV-6. Surveyor II vibration levels
during launch phase
Event and
accelerometer
Maximum zero-to-peak
acceleration, g*
Surveyor ff
Surveyor /
Liftoff
CY540
CY780
BECO
CY780
Shroud separation
CY780
1.5
1.5
1.5
1.25
2
1
b
b
■At recorded within the frequency bandwidth of 6 to 600 cpt.
Not meotured on Surveyor 1 becouse of accelerometer data commutation.
by ground command rather than automatically with the
spacecraft in cruise mode. During the star mapping
sequence, some difficulty was experienced in readily
identifying the celestial bodies because of earth and moon
reflections entering the star sensor.
Telemetry data received during the star acquisition
sequence indicated an apparent decrease in sensitivity of
Receiver B of approximately 18 db below the predicted
value. A degradation of 16 db in Receiver B sensitivity
(i.e., a receiver malfunction) would have precluded the
possibility of retaining two-way lock during the mid-
course attitude maneuvers and thrusting sequence. A
weak Receiver B signal level had also been noted during
countdown operations (see Section II). However, a first
evaluation of this anomaly had attributed the problem
to a poor RF link due to service tower removal. Subse-
quent in-flight calibration of this telemetery channel indi-
cated that the premission calibration of signal strength
vs the AGC reading was in error. Midcourse maneuver
was done in two-way lock, with the bit rate at 1100
rather than 550 bit/sec, since sufficient margin was pre-
dicted throughout the midcourse sequence.
When the command to ignite the three vernier engines
was sent to the spacecraft as part of the standard mid-
course velocity correction sequence, Vernier Engine 3
did not respond properly. The thrust provided by Ver-
nier Engines 1 and 2 resulted in a spinning of the space-
craft at approximately 1.22 rev/sec. An initial attempt to
halt the spinning, with the cold gas jets being controlled
by the flight control subsystem operating in the rate
mode, was terminated when approximately 60% of the
available gas supply was required to reduce the spin
rate to approximately 0.97 rev/sec, thereby indicating
that the available gas supply would not be sufficient to
stop the spacecraft rotation. Because the spacecraft was
spinning about an axis such that the sun was not in the
upper hemisphere of the vehicle, the solar panel was not
illuminated, and the main and auxiliary batteries were the
only spacecraft power sources from this point in the mis-
sion. Subsequent attempts (39 in all) to obtain normal
firing of Vernier Engine 3 were unsuccessful and resulted
in the spacecraft rotational rate being increased to a
maximum of 2.43 rev/sec. With the available power
decreasing steadily, it was decided to fire the main retro
motor. Communication with the spacecraft was lost ap-
proximately 30 sec following ignition of the retro motor.
The spacecraft spin rate profile is shown in Fig. IV-11.
16. Surveyor // Failure Reviev^ Board Summary and
Recommendations
A Failure Review Board (FRB) was convened at JPL
to review events surrounding the Surveyor II spacecraft
failure and, if possible, to determine its cause. Merriber-
ship of the FRB consisted of representatives from JPL,
HAC, Reaction Motors Division of Thiokol, and NASA
Offices of Space Science and Applications. As a result
of extensive review and analysis of the available flight
data pertaining to the performance of the vernier pro-
pulsion system during each thrusting period, the FRB has
concluded that:
(1) Engine 3 never ignited.
(2) Engines 1 and 2 operated inconsistently during
some firings following midcourse, if not also during
midcourse itself.
Noted examples of performance inconsistencies ob-
tained from analyses of thrust-command, thermal-
response, and strain-gage data for the three engines are:
(1) There is no evidence of oxidizer flow to Engine 3
during midcourse, but there is evidence of possible
oxidizer flow for 2.0-sec or longer firings foUowing
midcourse — all of insufficient quantity to support
combustion.
(2) Engine I may have failed to ignite on any 0.2-sec
commanded impulses, except for Firing 10. (Vernier
engine firings are numbered consecutively begin-
ning with the first post-midcourse firing.) This
firing yielded a temperature response equivalent
to that for the 2-sec or longer firings for some yet-
unexplained reason. The extrapolated peak tem-
perature for Engine 1 on Firing 26 is approximately
4 times as high as temperature responses observed
JPL TECHNICAL REPORT 32-1086
39
<
Q.
2.50
2.25
-
2.00
-
1.75
-
1,50
-
1.25
-
1.00
-
HELIUM DUMPED
21,5-sec THRUST
0.2-sec THRUST (5 TIMES)
2.5-sec THRUST
0.75 -
0.50
0.25
COLD GAS JETS ON FOR
14 min AFTER MIDCOURSE
2 -sec THRUST (2 TIMES)
■9.8 -sec MIDCOURSE THRUST
MAIN RETRO
IGNITED
LOST
COMMUNICATIONS
(30 sec AFTER
RETRO IGNITION)
0.2-sec THRUST (5
TIMES) FOLLOWED
BY A 2 -sec THRUST
2- sec THRUST
0.2-sec THRUST (5 TIMES)
8 12 16 20
TIME AFTER MIDCOURSE, hr
Fig. IV-1 1 . Surveyor (/ spin rate profile
24
28
32
36
for firings of equivalont cominancUxl duration. Con-
.sistcnt with thi.s excessive temperature rise, En-
gine 1 flight strain gage data appears to indicate a
shutdown delayed on the order of 2.5 sec hcyond
the shutolf command.
Oxidizer pressure data plotted from this could
result in an interpretation of Firing 26 as a "normal"
shutdown, the elevated temperature being a prob-
able consequence of an "oxidizer-rich" propellant
mixture. While this latter interpretation cannot be
totally discounted, a preponderance of judgment
supports the theory of a delayed shutdown.
{'.]) Engine 2 apparently ignited on all of the 0.2-sec
impulses. Between 2-sec Firings 1 and 8, the
extrapolated peak temperature rise underwent a
factor of 3 increase, currently attributable to some
unexplained performance variability.
On Firing 39 (21.5-sec duration). Engine 2 ap-
peared to have exhibited a significantly higher
than expected temperatiire for the commanded
minimum thrust.
The above examples show the range of performance
variability for Engines 1 and 2, as opposed to the sus-
tained restriction to oxidizer flow apparently exhibited
for Engine 3.
In conducting the failure investigation, the FRB has
tried to identify and evaluate all possible failure modes
which could account for the Swvcijor II failure. A num-
ber of failure modes (for T(>A's, tanks, lines, etc., of the
VPS) have been explored and dispositioned as unlikely
in accordance with interpretations of flight data. A serious
handicap to this evahiation and dispositioning process
40
JPL TECHNICAL REPORT 32-1086
•■has been the lack of adequate telemetry indicative of VPS
performance, particularly during thrusting periods. Cer-
tain data critical to the analyses are noisy and/or ambig-
"uous, requiring interpretations and extrapolations which
have not always netted agreements between analysts.
Lack of adequate signatures of several critical data chan-
nels had also contributed to the uncertainty.
One of the problems of the FRB was to determine
from the current and voltage telemetry precisely what
current was drawn by the solenoid-operated valves
(SOV) during vernier thrusting periods. It has been dem-
onstrated through analysis and tests that improper elec-
trical signals to the engine SOV solenoids could result in
malfunctioning of the three vernier engines. The limiting
sampling rate and noise or ripple on the telemetry chan-
nel resulted in uncertainty in the data. A significant
portion of the time and energy of the committee was
devoted to a study of the data, yet the FRB has been
unable to support the hypothesis of an electrical failure.
The FRB has been unable to determine the precise
cause of the failure. The position taken by the FRB is
that it cannot postulate and support a most-probable
single cause of failure — although one may exist. Since no
single-failure mode has been estabhshed, the Board must
leave open the possibility of multiple occurrences, pos-
sibly directly linked to a prime failure.
In recognition of failure possibihties remaining open
in several areas, the FRB has recommended specific cor-
rective actions for each suspect area without prejudice.
These recommendations are contained in Appendix C.
They extend to several elements of the VPS, electrical
power system, and flight control; they cover design
changes, improved test procedures, and reassignments of
telemetry channels to provide better preflight and in-flight
diagnostics.
To the extent that improved diagnostic tests (including
the demonstration of electrical performance margins) can
be incorporated, a higher assurance of spacecraft flight
readiness should result, providing a measure of insurance
against a similar type of failure in future flights.
B. Structures and Mechanisms
The vehicle and mechanisms subsystem provides
support, alignment, thermal protection, electrical inter-
connection, mechanical actuation, and touchdown sta-
bilization for the spacecraft and its components. The
subsystem includes the basic spaceframe, landing gear
mechanism, crushable blocks, omnidirectional antenna
mechanisms, antenna/solar panel positioner (A/SPP),
pyrotechnic devices, electronic packaging and cabling,
thermal compartments, thermal switches, separation sens-
ing and arming device, and secondary sun sensor.
1. Spaceframe and Substructure
The spaceframe, constructed of thin-wall aluminum
tubing, is the basic structure of the spacecraft. The sub-
structure is used to provide attachment between some
subsystems and the spaceframe. The landing legs and
crushable blocks, the retrorocket engine, the Centaur
interconnect structure, the vernier propulsion engines and
tanks, and the A/SPP attach directly to the spaceframe.
The subsh-ucture is used for the thermal compartments,
TV subsystem, auxiliary battery, RADVS antennas, flight
control sensor group, attitude control nitrogen tank, and
the vernier system helium tank.
Gyro data and linear potentiometer data indicated
that the separation of the spacecraft from the Centaur
adapter occurred as predicted without physical contact of
one body with the other after initiation of separation. (Also
see Section III for discussion of .spacecraft separation.)
The spacecraft structure survived the boost, cruise and
midcourse phases as well as numerous attempts to start
Vernier Engine 3. The vibration accelerations were sig-
nificant only during the boost phase.
2. LancJing Gear and Crushable Blocks
The three landing-leg mechanisms are each made up
of a landing leg, an intermediate A-frame, a shock
absorber, a footpad, a locking strut, and a position
potentiometer (Fig. IV-12).
The shock absorber, intermediate A-frame, and tele-
scoping lock strut are interconnected to the spaceframe
for folded stowage in the nose shroud. Torsion springs at
the leg hinge extend the legs when the squib-actuated
pin pullers are operated by Centaur or earth command.
The hydraulic shock absorber compresses with landing
load. The shock absorbers, foot pads, and crushable
blocks are designed to absorb the landing shock. After
landing, the shock absorbers are locked in place by
squib-actuated pin locks.
The legs opened and locked in the landing position
when the Centaur gave the command. This was verified by
the potentiometers and the locking strut microswitches.
JPl TECHNICAL REPORT 32-1086
41
-TELESCOPING
LOCK STRUT
STRAIN GAGE
LANDING LEG
-ALUMINUM HONEYCOMB BLOCK
SHOCK ABSORBER
PHOTOMETRIC
TARGET
ATTITUDE-CONTROL JET-
FOOT PAD-
Fig. IV-12. Landing leg assembly
3. Omnidirectional Antennas
The omnidirectional antennas are mounted on the ends
of folding booms hinged to the spaceframe. Pins retain
the booms in the stowed position. Sqiiib-actiiated pin
pullers release the booms by Centaur or earth command,
and torsion springs deploy the antennas. The booms are
then locked in position.
The omniantenna booms were extended by Centaur
command, and both antennas were locked in the landing
or transit position as indicated by the telemetry.
4. Antenna and Solar Panel Positioner (A/SPP)
The A/SPP supports and positions the high-gain planar
array antenna and solar panel. The planar array antenna
and solar panel have four axes of rotation: roll, polar,
solar, and elevation (Fig. IV-13). Stepping motors rotate
the axes in either direction in response to commands from
earth or during automatic deployment following Centaur/
spacecraft separation. This freedom of movement permits
orienting the planar array antenna toward the earth and
the solar panel toward the sun.
The solar axis is locked in a vertical position for stow-
age in the nose shroud. After launch, the solar panel is
positioned parallel to the spacecraft X axis. The A/SPP
remains locked in this position until after touchdown, at
which time the roll, solar, and elevation axes are re-
leased. Potentiometers on each axis are read to indicate
A/SPP orientation. Each command from earth gives Vs
degree of rotation in the roll, solar, and elevation axes
and Viii degree in the polar axis.
The A/SPP operated as expected during the mission.
After the shroud was ejected, the roll and solar axes
moved to their transit positions.
5. Thermal Compartments
Two thermal compartments (A and B) house thermally
sensitive electronic items. Equipment in the compart-
ments is mounted on thermal trays that distribute heat
Table IV-7. Thermal compartment component
installation
Comparlment A
Compartment B
Receivers (2)
Central commond decoder
Transmitters (2)
Boost regulator
Main battery
Central signal processor
Battery charge regulator
Signal processing auxiliary
Engineering mectianisms
Engineering signal processor
auxiliary
Television auxiliary
Low doto rate auxiliary
Ttiermai control and tieater
Thermal control and heater assembly
assembly
Auxiliary battery control
Auxiliary engineering signal
processor
42
JPL TECHNICAL REPORT 32-7086
SOLAR PANEL
PLANAR ARRAY
ANTENNA
Fig. IV-13. Antenna/solar panel configuration
throughout the compartments. An insulating blanket,
consisting of 75 sheets of 0.25-mil-thick aluminized mylar,
is installed between the inner shell and the outer protec-
tive cover of the compartments. Compartment design
employs thermal switches which are capable of varying
the thermal conductance between the inner compartment
and the external radiating surface. The thermal switches
maintain thermal tray temperature below +125°F. Each
compartment contains a thermal control and heater
assembly to maintain the temperature of the thermal tray
above a specified temperature (above 40° F for Compart-
ment A and above 0°F for Compartment B). The thermal
control and heater assembly is capable of automatic
operation, or may be turned on or off by earth command.
Components located within the compartments are identi-
fied in Table IV-7.
6. Thermal Switch
The thermal switch is a thermal-mechanical device
which varies the conductive path between an external
radiation surface and the top of the compartment (Fig.
IV-14). The switch is made up of two contact surfaces
which are ground to within one wavelength of being op-
tically Hat. One surface is then coated with a conforming
substance to form an intimate contact with the mating
surface. The contact actuation is accomplished by four
bimetallic elements located at the base of the switch.
JPL TECHNICAL REPORT 32-1086
43
4f --^^jjrwwi A' i-rtf"-'
INNER CONTACT RING
Bl- METAL CLAMP
Bl- METAL ACTUATOR
INNER CONTACT PLUG
CONDUCTOR FOIL
RADIATING PLATE
OUTER CONTACT
RING
Fig. IV-14. Thermal switch
Tlu-sr t'lt'iiu'iits arc connected mechanically to the top of
the compartment so that the compartment temperature
controls the switch actuation. The switches are identical,
but arc adjusted to open at three different temperatures:
65, 50 and 40 °F.
The external radiator surface is such that it absorbs
only 12% of the solar energy incident on it and radiates
74% of the heat energy conducted to its surface. When
the switch is closed and the compartment is hot, the
switch loses its heat energy to space. When the compart-
ment gets cold, the switch contacts open about 0.020 in.,
thereby opening the iieat-conductive path to the radiator
and thus reducing the heat loss through the switch to
almost zero.
The thermal switches kept the electronics at or below
the maximum temperature at all times during the flight.
7. Pyrotechnic Devices
The pyrotechnic devices installed on Surveyor 11 are
indicated in Table IV-8. All the scjuibs used in these
devices are electrically initiated, hot-bridgevvire, gas-
generating devices. Qualification tests for flight squibs
included demonstration of reliability at a firing current
44
JPt TECHNICAL REPORT 32-7086
Table IV-8. Pyrotechnic devices
Type
Location and use
Quantify of
devices
Quantity of
squibs
Command source
Pin pullers
Lock and release Omniantennos A and B
2
2
Centaur programmer
Pin pullers
Lock and release landing legs
3
3
Centour programmer
Pin pullers
Lock and release planar antenna and solar panel
7
7
Separation sensing and arming device and
ground station
Pin puller
Lock and release vernier thrust chamber No. 1
1
1
Ground station
Separation nuts
Retro rocket attach and release
3
6
Flight control subsystem
Valve
Helium gas release and dump
I
2
Ground station
Pyro switches
EMA board No. 4, RADVS power on and off
4
4
Ground station and flight control subsystem
Initiator squibs
Safe and arm assembly retrorocket initiators
1
2
Flight control subsystem
Locking plungers
Landing leg, shock absorber locks
3
3
Ground station
25
30
level of 4 or 4.5 amp. "No Fire" tests were conducted at
a 1-amp or 1-w level for 5 min. Electrical power re-
quired to initiate pyrotechnic devices is furnished by the
spacecraft main battery. Power distribution is through
19.0- and 9.5-amp constant-current generators in the engi-
neering mechanism auxiliary (EMA).
All scheduled pyrotechnic devices functioned normally
upon command. Mechanical operation of locks, valves,
switches, and plunger, actuated by squibs, was indicated
on telemetry signals as part of the spacecraft engineering
measurement data.
8. Electronic Packaging and Cabling
The electronic assemblies for Surveyor 11 provided
mechanical support for electronic components in order
to insure proper operation throughout the various envi-
ronmental conditions to which they were exposed during
the mission. The assemblies (or control items) were con-
structed utilizing sheet metal structure, sandwich-type
etched circuit board chassis with two-sided circuitry,
plated through holes, and/or bifurcated terminals. Each
control item, in general, consists of only a single func-
tional subsystem and is located either in or out of the
two thermally controlled compartments, depending on
the temperature sensitivity of the particular subsystem.
Electrical interconnection is accomplished primarily
through the main spacecraft harness. The cabling system
is constructed utilizing a light-weight, minimum-bulk,
and abrasion-resistant wire which is an extruded teflon
having a dip coating of modified polyimide.
C. Thermal Control
The thermal control subsystem is designed to provide
acceptable thermal environments for all components dur-
ing all phases of spacecraft operation. Spacecraft items
with close temperature tolerances were grouped together
in thermally controlled compartments. Those items with
wide temperature tolerances were thermally decoupled
from the compartments. The thermal design fits the
"basic bus" concept in that the design was conceived to
require minimum thermal design changes for future mis-
sions. Monitoring of the performance of the spacecraft
thermal design is done by 74 temperature sensors which
are distributed throughout the spacecraft as follows:
Flight control
6
Mechanisms
3
Radar
6
Electrical power
3
Transmitters
2
Approach TV
1
Survey TV
2
Vehicle structure
25
Propulsion
15
Engineering pay load 11
JPl TECHNICAL REPORT 32-1086
45
SUPER INSULATION
LOCATIONS
VYCOR MIRRORS
LOCATIONS
COMPARTMENT A
LEG 2
FUEL AND OXIDIZER
TANKS (6)
, COMPARTMENT B
RETRO ROCKET
MOTOR
LEG 3
COMPARTMENT A
r#.'
LEG 2
^COMPARTMENT B
.FLIGHT-CONTROL
SENSORS
LEG 3
OMNIANTENNA A
POWER SUPPLY
MODULATOR
ALTIMETER/VELOCITY
SENSING ANTENNA
FWSSIVE AREAS
VERNIER ENGINE I
VELOCITY-SENSING
ANTENNA
FLIGHT CONTROL
SENSORS
OMNIANTENNA B
A/SPP ORIENTATION FOR TRANSIT
COMPARTMENT A
SOLAR PANEL
FUEL TANK 2^
VERNIER ENGINE 2
LINES
TV CAMERA' / /
(APPROACH) / /
ACTIVE HEATER
LOCATIONS
VERNIER ENGINE I LINES
COMPARTMENT B
OXIDIZER TANK 3
1 \„
OXIDIZER TANK i' 'YjuRVEyT '^'''""' ''^'"" '"'^^'''"'
Fig. IV-15. Thermal design
46
JPL TECHNICAL REPORT 32-7086
The spacecraft thermal control subsystem is designed
to function in the space environment, both in transit and
on the lunar surface. Extremes in the environment as well
as mission requirements on various pieces of the space-
craft have led to a variety of methods of thermal control.
The spacecraft thermal control design is based upon the
absorption, generation, conduction, and radiation of heat.
Figure IV-15 shows those areas of the spacecraft serviced
by different thermal designs.
The radiative properties of the external surfaces of
major items are controlled by using paints, by polishing,
and by using various other surface treatments. Reflecting
mirrors are used to direct sunlight to certain components.
In cases where the required radiative isolation cannot
be achieved by surface finishes or treatments, the major
item is covered with an insulating blanket composed of
multiple-sheet aluminized mylar. This type of thermal
control is called "passive" control.
The major items whose survival or operating tempera-
ture requirements cannot be achieved by surface finish-
ing or insulation alone use heaters that are located within
the unit. These heaters can be operated by external com-
mand, thermostatic actuation, or both. The thermal con-
trol design of those units using auxiliary heaters also
includes the use of surface finishing and insulating
blankets to optimize heater effectiveness and to minimize
the electrical energy required. Heaters are considered
"active control."
Items of electronic equipment whose temperature re-
quirements cannot be met by the above techniques are
located in thermally controlled compartments (A and B).
Each compartment is enclosed by a shell covering the
bottom and four sides and contains a structural tray on
which the electronic equipment is mounted. The top of
each compartment is equipped with a number of
temperature-actuated switches (9 in Compartment A and
6 in Compartment B). These switches, which are attached
to the top of the tray, vary the thermal conductance
between the tray and the outer radiator surfaces, thereby
varying the heat-dissipation capability of the compart-
ments. When the tray temperature increases, heat transfer
across the switch increases. During the lunar night, the
switch opens, decreasing the conductance between the
tray and the radiators to a very low value in order to con-
serve heat. When dissipation of heat from the electronic
equipment is not sufficient to maintain the required mini-
mum tray temperature, a heater on the tray supplies the
necessary heat. The switches are considered "semi-active."
Examples of units which are controlled by active, semi-
active, or passive means are shown in Fig. IV-15.
The thermal performance of the spacecraft up to mid-
course was completely as expected with one exception.
Vernier Line 2 heater was apparently on, although the
line temperature continued to drop. It is estimated that
the line temperature would have been close to the mini-
mum operating limit of 0°F at the start of thermal
descent.
A change had been made in the auxiliary battery paint
pattern because of the low temperature experienced on
Surveyor I. Auxiliary battery temperature was as desired
on Surveyor II.
After the midcourse failure, the spacecraft tempera-
tures stabilized at a new equilibrium and did not change
much thereafter. Compartment A stabilized at approxi-
mately 75 °F, Compartment B at 53 °F, and flight control
units at slightly above normal cruise temperatures until
the flight control subsystem was turned off. All other
parts of the spacecraft stabilized at or stayed within their
operational temperature limits, except for the RADVS
SDC and the shock absorbers. Thus, even in a highly
nonstandard orientation, the thermal control subsystem
continued to allow most spacecraft operations.
Appendix D shows graphically the temperature history
of the major spacecraft components.
D. Electrical Power
The electrical power subsystem is designed to generate,
store, convert, and distribute electrical energy. A block
diagram of the subsystem is shown in Fig. IV- 16. The
subsystem derives its energy from the solar panel and
the spacecraft battery system. The solar panel converts
solar radiation energy into electrical energy. Solar panel
power capability is affected by temperature and the inci-
dence of solar radiation and varies from 90 to 55 w.
The spacecraft battery system consists of a main battery
and an auxiliary battery. The main battery is a secondary
or rechargeable battery; the auxiliary battery is a primary
or nonrechargeable battery.
The batteries provide about 4090 w during transit, the
balance of the energy being supplied by the solar panel.
The maximum storage capacity of the main battery is
180 amp-hr; that of the auxiliary battery is 50 amp-hr.
JPL TECHNICAL REPORT 32-1086
47
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JPt TECHNICAL REPORT 32-1086
The selection of battery operation mode is determined
by the auxihary battery control (ABC). There are three
modes of battery operation: main battery mode, auxiliary
battery mode, and high-current mode. In the main bat-
tery mode only the main battery is connected to the
unregulated bus. This is the nominal configuration. In
the auxiliary battery mode the main battery is connected
to the unregulated bus through a series diode while the
auxiliary battery is directly connected. In the high-current
mode both the main and auxiliary batteries are connected
to the unregulated bus without the series diode. The
battery modes are changed by earth commands except
that the ABC automatically switches to auxiliary battery
mode from main battery mode in case of main battery
failure. This automatic function can be disabled by earth
command.
The four modes of solar panel operation are controlled
by the battery charge regulator (BCR). In the on mode
the optimum charge regulator (OCR) tracks the volt-
ampere characteristic curve of the solar panel and hunts
about the maximum power point. In the OCR off mode
the solar panel output is switched off. This mode is in-
tended to prevent overcharging of the main battery by
the solar panel. In the OCR bypass mode the solar panel
is connected directly to the unregulated bus. This mode
is used in case of OCR failure. In the trickle charge
mode the main battery charging current is controlled by
its terminal voltage. Three BCR modes, excluding the
trickle charge mode, are controlled by earth commands.
The OCR off and trickle charge modes are automatically
controlled by the battery charge logic (BCL) circuitry.
When the main battery terminal voltage exceeds 27.5 v
or its manifold pressure exceeds 65 psia, the BCR goes
automatically to the off mode. The trickle charge mode
is automatically enabled when the main battery terminal
voltage reaches 27.3 v. The BCL can be disabled by earth
command.
Current from the BCR and spacecraft batteries is dis-
tributed to the unregulated loads and the boost regulator
(BR) via the unregulated bus. The voltage on the unregu-
lated bus can vary between 17.5 and 27.5 v, with a
nominal value of 22 v. The BR converts the unregulated
bus voltage to 29.0 v ±1% and supplies the regulated
loads. The preregulator supplies a regulated 30.4 v dc to
the preregulated bus. The essential loads are fed by the
preregulated bus through two series diodes. The diodes
drop the preregulated bus voltage of 30.4 v to the essen-
tial bus voltage of 29.0 v. The preregulated bus also feeds
the flight control regulator and the nonessential regu-
lator, which in turn feeds the flight control and nonessen-
tial busses. These regulators can be turned on and off by
earth commands. The nonessential regulator has a bypass
mode of operation which connects the preregulated bus
directly to the nonessential bus. This mode is used if the
nonessential regulator fails.
The power subsystem operated normally throughout
the mission. Table IV-9 verifies that telemetered param-
eters were in close agreement with the predicted values.
During the post-injection coast phase the average regu-
lated load was 2.45 amp (Fig. IV-17), with an average BR
efficiency of 77.4%, and the average unregulated current
was 0.72 amp (Fig. IV-18). Comparable time period pre-
dictions indicate that the regulated load should be
2.29 amp and the unregulated current 0.80 to 0.83 amp.
During low-power transmitter interrogation, the regulated
output was approximately 100 to 150 ma higher than pre-
dicted. During high-power transmitter interrogation, this
current was as predicted.
For the above mission period, the average (OCR) out-
put current was 3.22 amp (Fig. IV-19), which agrees
closely with test data. The average solar panel output
current was 1.83 amp (Fig. IV-20) at an average voltage
of 48.2 (Fig. IV-21). The overall OCR efficiency was
about 81%. The OCR solar panel combination supplied
an average of 68% of the total system electrical loads,
with the battery providing the remaining 32% of the load.
Battery pressure stabilized during this period to 15 psi
(Fig. IV-22) at a steady-state battery temperature of
99 °F, both measurements falling well within the normal
safe operating limits. Main battery terminal voltage and
discharge current are shown in Figs. IV-23 and IV-24.
The auxiliary battery voltage history is shown in Fig.
IV-25. Figures IV-26, IV-27, and IV-28 show the BR
preregulator, 29-v nonessential, and unregulated bus volt-
ages vs mission time.
Following midcourse, the power system operated nor-
mally for the life of the spacecraft. The batteries pro-
vided the total spacecraft power from midcourse to the
end of the mission. Figure IV-29 shows actual vs pre-
dicted battery energy consumption for the entire mission.
The solar panel alignment was such that it did not
receive any radiation from the sun. Under these conditions
the solar panel could not provide power to the spacecraft.
JPL TECHNICAL REPORT 32-1086
49
4
ID O
2
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4)
>
diuD 'iN3ManO
duio 'iNayano
50
JPt TECHNICAL REPORT 32-7086
)
/
3.2
2.4
Fig. IV-19. OCR output current
3
0.8
0.4
^V^v-^fWvVTvH r--^"A.>^^ -[!i|fVV||i|#^^^
DATA NOT
AVAILABLE
J 1 u
12 3 4 5 6
-I 1 1 1 L_U.
12 13 14 15 16 17
TIME FROM LIFTOFF, hr
2.0
e 1-6
Z 1.2
LU
tr
a:
3 0.8
-I 1 1 rV-
DATA NOT
AVAILABLE
' 1 1 I I
1 2 3 4 5
-^
-I 1 I I L
Fig. IV-20. Solar eel
array current
II 12 13 14 15 16 17 18 19
TIME FROM LIFTOFF, hr
Fig. IV-21. Solar cell array voltage
<
d 24
>
"xr^vdWy^vAW" — ^Aj'*'^ — yyui^^y^^
DATA NOT
AVAILABLE
-J 1 1 \ u
2 3 4 5 6 7 '
12 13 14 15 16 17
TIME FROM LIFTOFF, hr
JPL TECHNICAL REPORT 32-1086
51
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53
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41
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JPL TECHNICAL REPORT 32-1086
55
200
150
100
50
y
ACTUAL/
^
^
PREDICTED,.,--
-^
^
20 30
TIME FROM LIFTOFF, hr
40
50
Fig. IV-29. Actual vs predicted battery energy consumption
E. Propulsion
The propulsion subsystem supplies thrust force during
the midcourse correction and terminal descent phases of
the mission. The propulsion subsystem consists of a
vernier engine system and a solid-propulsion main retro-
rocket motor. The propulsion subsystem is controlled by
the flight control system through preprogrammed ma-
neuvers, commands from earth, and maneuvers initiated
by flight control sensor signals.
1. Vernier Propulsion
The vernier propulsion subsystem supplies the thrust
forces for midcourse maneuver velocity vector correc-
tion, attitude control during main retrorocket motor
burning, and velocity vector and attitude control during
terminal descent. The vernier engine system consists of
three thrust chamber assemblies and a propellant feed
system. The feed system is composed of three fuel tanks,
three oxidizer tanks, a high-pressure helium tank, pro-
pellant lines, and valves for system arming, operation,
and deactivation.
Fuel and oxidizer are contained in six tanks of equal
volume with one pair of tanks for each engine. Each tank
contains a Teflon expulsion bladder to permit complete
and positive expulsion and to assure propellant control
under zero-^ conditions. The oxidizer is nitrogen tetroxide
(N^Oj) with 10% by weight nitric oxide (NO) to depress
the freezing point. The fuel is monomethyl hydrazine
monohydrate (72 MMH • 28 H.,0). Fuel and oxidizer
ignite hypcrgolically when mixed in the thrust chamber.
The total usable propellant load is 178.3 lb. The arrange-
ment of the tanks on the spaceframe is illustrated in
Fig. IV-30. Propellant freezing or overheating is prevented
by a combination of active and passive thermal controls,
utilizing surface coatings, multilayered blankets, and
electrical and solar heating. The propellant tanks are
thermally isolated from a spaceframe to insure that the
spacecraft structure will not function as a heat source or
as a heat sink.
Propellant tank pressurization is provided by the helium
tank and valve assembly (Fig. IV-31). The high-pressure
helium is released to the propellant tanks by activating a
squib-actuated helium release valve. A single-stage regu-
lator maintains the propellant tank pressure at 730 psi.
Helium relief valves relieve excess pressure from the
propellant tanks in the event of a helium pressure regu-
lator malfunction.
The thrust chambers (Fig. IV-32) are located near the
hinge points of the three landing legs on the bottom of
the main spaceframe. The moment arm of each engine
is about 38 in. Engine 1 can be rotated ±6 deg about
an axis in the spacecraft X-Y plane for spacecraft roll
control. Engine 1 roll actuator is unlocked after boost.
Engines 2 and 3 are not movable. The thrust of each
engine (which is monitored by strain gages installed on
each engine mounting bracket) can be throttled over a
range of 30 to 104 lb. The specific impulse varies with
engine thrust.
Prior to launch the vernier propulsion system propellant
tanks are loaded with a nominal 109 lb of oxidizer and
75 lb of fuel. The propellant tanks are then pressurized
56
JPt TECHNICAL REPORT 32-1086
HELIUM TANK
LEG 3
VERNIER ENGINE 3
(FIXED)
HELIUM LINES
OXIDIZER TANK (3)
FUEL TANK (3)
VERNIER ENGINE 2
(FIXED)
VERNIER ENGINE I
(GIMBALLED)
Fig. IV-30. Vernier propulsion system installation
LEG I
JPL TECHNICAL REPORT 32-1086
57
PRESSURE
TRANSDUCER
SQUIB-ACTUATED
RELEASE VALVE
PI7 TEMPERATURE SENSOR
SQUIB-ACTUATED
DUMP VALVE
HELIUM LINES
TO FUEL TANKS
HELIUM LINES
TO OXIDIZER
TANKS
P9
TEMPERATURE
PI4
TEMPERATURE
PROPELLANT
LINES
-P7 TEMPERATURE
PIO TEMPERATURE
THRUST CHAMBER
ASSEMBLY (TYP)
TEMPERATURE
Fig. IV-31. Vernier propulsion system schematic showing locations of pressure and temperature sensors
58
JPL TECHNICAL REPORT 32-1086
Fig. IV-32. Vernier engine thrust chamber
JPL TECHNICAL REPORT 32-1086
59
to 300 psi pad pressure of helium. The high-pressure
hehum tank is pressurized to a nominal 5175 psi. The
vernier system remains in this condition through launch
until 7 min before midcourse firing, at which time a
squib in the helium release valve is fired. This allows the
helium regulator to pressurize the propellant tanks to
their nominal working pressure of 730 psi. At midcourse
the vernier system is given a command which turns on
all three vernier engines to a thrust level equal to 0.1 g
for a specified time depending upon the correction de-
sired. After the midcourse maneuver the system remains
in the fully pressurized state until the terminal descent
sequence. For the terminal descent operation, the vernier
engines are ignited 1 sec before the main retro motor
fires. During main retro burn, the vernier engines pro-
vide attitude control. At the end of main retro burn, the
vernier engines are programmed to full thrust to facili-
tate the retro separation. The vernier engines are then
throttled to give an optimum range/velocity profile de-
scent. At approximately 13 ft above the lunar surface, the
engines are shut down and the spacecraft free-falls to
the lunar surface.
In addition to the pressure and temperature sensors
shown in Fig. IV-31, the vernier propulsion system was
instrumented with strain gages on each of the engine
mounting brackets. During midcourse firing, the strain
gages at Engines 1 and 2 indicated thrust levels of ap-
proximately the same magnitude as commanded thrust,
while the strain gage at Engine 3 indicated essentially
no thrust (Fig. IV-33). In fact, the resulting tumbling of
the spacecraft verified that Engine 3 did not fire.
Out of several equally probable malfunctions which
could be postulated on the basis of real-time data obser-
vation, the only type which offered the possibility of
correction during the remaining flight time was that
involving a "sticking" component in the vernier engine
assembly. Therefore, the remainder of the flight was
devoted primarily to pulsing and firing the vernier en-
gines and obtaining diagnostic telemetry data.
In support of the Surveyor II Failure Review Board
(FRB) investigation, a thorough analysis of all available
data and a program of simulation tests of various pos-
sible conditions have been conducted to determine the
most probable causes of the failure. The exact cause of
failure has not been determined. (The FRB summary and
recommendations are discussed in Section IV-A.)
During the period from launch to midcourse, all of
the telemetered propulsion system data indicated normal
a:
X
t-
80
40
1/^
-^
r^-
VERNIER ENGINE 1
1
80
40
\f
^^
::;::;;;
I
VERNIER
ENGINE 2
120
80
40
-40
r'
■ STRAIN GAGE READING
THRUST COMMAND
VERNIER ENGINE 3
TIME, sec
(REF. TIME: 05:00 03.6 GMT SEPTEMBER 21)
Fig. IV-33. Strain gages and thrust command
signals at midcourse
operation, with two minor exceptions — a low temperature
on the oxidizer line to Engine 2, and an unexpectedly
large drop in helium supply pressure when the release
valve was actuated to pressurize the propellant feed
system. The two observed anomalies, however, were not
related to the subsequent failure to obtain ignition at
Engine 3 during the midcourse firing.
Temperature histories of the propulsion subsystem
components are shown in Appendix D (Figs. D-6 and
D-7). Locations of temperature sensors on the vernier
system are shown in Fig. IV-31, and on the main retro-
rocket in Fig. IV-34. During the period from launch to
60
JPL TECHNICAL REPORT 32-1086
THRUST SKIRT
STRUT
Fig. IV-34. Main retrorocket motor
midcourse, the temperatures of all propulsion components
were within their prediction bands except the tempera-
ture of Oxidizer Line 2. Post-flight analysis of this
anomaly indicates that, although the line heater thermo-
stat was closed, the heater was apparently defective.
Figure IV-35 shows the history of the gas pressure
measured at the helium tank and the history of the
oxidizer pressure measured at Tank 3. Since there are
check valves between the fuel and oxidizer helium pres-
surization manifolds (see Fig. IV-31), the fuel pressure
may differ from the oxidizer pressure. In the present
spacecraft design, no measurement of fuel pressure is
provided.
sure to the propellant feed system. The oxidizer pressure
rose to its expected level within 2 sec, and during that
2-sec period the indicated helium tank pressure dropped
735 psi. Pressurization of the helium lines and normal
ullage volume of the tanks should have resulted in a
supply pressure drop of only about 225 psi.
When the midcourse failure occurred, it was hypothe-
sized that Fuel Tank 3 might have been empty, which
would explain the abnormal drop of helium tank pres-
sure. This hypothesis, however, was refuted by the tem-
perature measurements on the fuel tank, which indicated
a thermal history characteristic of a fully loaded tank.
The squib-actuated helium release valve was actuated
7 min before midcourse to provide full operating pres-
Post-flight analysis of point-by-point telemetry data
(Fig. IV-36) has revealed that the indicated helium tank
JPL TECHNICAL REPORT 32-1086
61
UJ
q:
ID
CO
(/)
UJ
IT
CL
6000
5000
4000
3000
2000
1000
1
HELIUM TANK,
SENSOR PI
^
1
PROPELLAN
PRESSURIZ
T
ATION
HE
:lium dump
/
TRANSDUCER i
NULL SHIFT ^
O
^— —
' —
/
Oo
MIDCOUF
FIRING
(9.8 se
SE
c)
O
o 6
O
° 8ci
FIRING
(21.5
No. 39
sec)
1
hi/-vf-t>a
^L MISSION
)ICTION
T DATA
T DATA COR
NULL SHIFT
O FLIGH
FLIGH
FOR
RECTED
S
800
600
400
200
OXIDIZER
SENSOR
TANK 3.
P2
C
r^
o
O C
)
■o
o f>^
o
<>-CXOlC0_
K n r\r)
\
!_V
15
20 25 30
TIME FROM LIFTOFF, hr
35
40
45
50
Fig. IV-35. Helium-tank and propellant-tank pressures vs time
pressure fell 530 psi in less than 0.25 see after firing of
the release valve squib. During the next IM; sec, an addi-
tional drop of 205 psi occurred; this rate of supply
pressure reduction is characteristic for filling the normal
ullage volume and helium lines. It has been concluded
that the extreme shock load imposed by the release valve
squib caused a shift of -500 to -530 psi in the null
reading of the helium tank pressure transducer. Trans-
ducer null shifts associated with activation of squib-
actuated helium valves had been observed twice during
62
JPL TECHNICAL REPORT 32-1086
5200
5100
5000
o 4900
UJ
a:
3
<g 4800
UJ
q:
a.
<I 4700
4600
4500
4400
4300
O
OS
-0.5 0.5 1.0 1.5
TIME FROM SQUIB ACTUATION, sec
2.0
2.5
Fig. IV-36. High-resolution plot of helium supply
pressure during propellant pressurization
qualification testing of the vernier propulsion system, but
in those cases the magnitude of shift was smaller:
— 180 psi in the T-2 drop test program and +150 psi in
the S-7 type-approval program.
The reduction in helium tank pressure which was
observed at midcourse (Fig. IV-35) is consistent with
9.8 sec of propellant usage from at least four tanks but
not consistent with usage from six tanks.
The observed helium tank pressure history subsequent
to midcourse is consistent with the corresponding tem-
perature history of the helium tank (Appendix D). The
departure of the helium tank temperature history from
the normal-mission prediction after midcourse suggests
that spacecraft tumbling caused the helium tank to be
shaded more often than it would have been with nominal
transit orientation.
The oxidizer pressure history indicates that the helium
regulator was operating within specification tolerances*
through the mission.
2. Main Retrorocket Motor
The main retrorocket, which performs the major por-
tion of the deceleration of the spacecraft during terminal
descent, is a spherical, solid-propellant unit with a par-
tially submerged nozzle to minimize overall length
(Fig. IV-34). The motor utilizes a carboxyl/terminated
polyhydrocarbon composite-type propellant and conven-
tional grain geometry.
The motor case is attached at three points on the main
spaceframe near the landing leg hinges, with explosive
nut disconnects for post-burnout ejection. Friction clips
around the nozzle flange provide attachment points for
the altitude marking radar (AMR). The retrorocket, in-
cluding the thermal insulating blankets, weighs approxi-
mately 1395 lb. This total includes about 1250 lb of
propellant. The thermal control design of the retrorocket
motor is completely passive, depending on its own
thermal capacity and an insulating blanket (21 layers of
aluminized mylar plus a cover of aluminized Teflon).
The prelaunch temperature of the unit is 70 ±5°F. At
terminal maneuver, when the motor is ignited, the pro-
pellant will have cooled to a thermal gradient with a bulk
average temperature of about 50 to 55° F.
The AMR normally triggers the terminal maneuver
sequence. When the retro firing sequence is initiated, the
retrorocket gas pressure ejects the AMR. The motor oper-
ates at a thrust level of 8,000 to 10,000 lb for approxi-
mately 39 sec at an average propellant temperature of
.50 °F.
The thermal sensors on the main retro motor case
closely followed the predicted values from launch until
the attempted midcourse correction at launch plus 16 hr.
Following the midcourse attempt, temperature deviations
due to the tumbling motion of the spacecraft were noted.
The upper case temperature continued to decrease at a
slightly higher than predicted rate. The lower case tem-
perature increased from 60°F at midcourse to about
72 °F when the retro was ignited at launch plus 45 hr.
Bulk temperature of the retro motor was estimated to be
about 70° F at retro ignition.
*Specification requirements are: operating pressure, 700 to 755 psi;
regulator lockup pressure, 795 psi maximum.
JPL TECHNICAL REPORT 32-1086
63
The retro motor was ignited while the spacecraft was
tumbling at approximately 2.3 rev/sec. Retro ignition
and burning were verified for approximately 32 sec after
which time all data was lost. Evaluation of spacecraft
accelerometer data indicates that the retro motor ignited
and burned normally until all data was lost. Because of
the tumbling nature of the spacecraft and termination of
data, no information was obtained on retro impulse,
thrust alignment, center of gravity, tailoff, or separation
characteristics.
F. Flight Control
The flight control subsystem is designed to (1) acquire
and maintain spacecraft orientation with respect to the
sun and the star Canopus, (2) orient the spacecraft for
a mid-mission trajectory error correction, (3) execute an
incremental velocity change and maintain spacecraft
stability during midcourse correction and terminal de-
scent, (4) execute a lunar terminal orientation, and (5) in
conjunction with a radar system (RADVS), a solid-
propellant retro motor, and three liquid-propellant (ver-
nier) engines, soft-land (at a nominal touchdown velocity
of 14 ft/sec) the spacecraft on the lunar surface.
The flight control subsystem consists of the appropriate
electronic equipment associated with a Canopus star
sensor (roll), primary sun sensors (pitch and yaw) for
spacecraft attitude in pitch, yaw, and roll during cruise,
an acquisition sensor for initial sun acquisition, three
gyros (pitch, yaw, and roll) for rate stabilization and
inertial control, and a d=0.75-^' precision accelerometer
for midcourse velocity control and acceleration control
during the terminal descent.
The control electronics process the reference sensor
outputs, earth-based commands, and the flight control
programmer and decoder outputs to generate the neces-
sary control signals for use by the vehicle control ele-
ments. A simplified flight control functional diagram
appears in Fig. IV-37. The vehicle control elements con-
sist of the attitude-control gas-jet activation valves, the
vernier engine thrust level control valves and gimbal
SENSORS
RANGE MARK
SIGNAL
EARTH -BASED
COMMANDS
r~
ACTUATORS
PROPULSION
--^
AMR
COMMAND
AND DATA
PROCESSING
1
RANGE AND
VELOCITY
IGNITION
EJECTION
MAIN
RETRO
ENGINE
1 1 »
MODE
SWITCHING
AND
CONTROL
ELECTRONICS
RADVS
THROTTLE
CONTROL
VERNIER
ENGINES
— r
ATTITUDE
ERROR
ROLL
CONTROL
1
1
1
1
1
SUN AND
STAR
ATTITUDE
ERROR
ACTUATOR
FEEDBACK |
INERTIAL
1 MANEUVERS
PULSE
CONTROL
1
* 1
a
ACCELERATION
GAS JETS
L
Fig. IV-37. Simplified flight control functional diagram
64
JPl TECHN/CAL REPORT 32-1086
actuator, and the main retro motor ignitor and separation
pyrotechnics.
The gas-jet attitude control system is a cold gas system
using nitrogen as a propellant. This system consists of a
gas supply system and three pairs of solenoid-valve-
operated gas jets interconnected with tubing (see Fig.
IV-38). The nitrogen supply tank is initially charged to
a nominal pressure of 4600 psia. Pressure to the gas jets
is controlled to 40 ±2 psia by a regulator.
Spacecraft attitude, acceleration, and velocity are con-
trolled as required by various "control loops" throughout
the coast and thrust phases of flight, as shown in Table
IV-10. Stabilization of the spacecraft tipoff rates after
Centaur separation is achieved through the use of rate
feedback gyro control (rate mode). After rate capture, an
inertial mode is achieved by switching to position feed-
back gyro control.
Because of the long duration of the transit phase and
the small unavoidable drift error of the gyros, a celestial
reference is used to continuously update the inertially
controlled attitude of the spacecraft.
PRESSURE
TRANSDUCER
TUBING
rC3=
CHARGING
VALVE
REGULATOR
>[ZD=:
TEMPERATURE
TRANSDUCER
PRESSURE
RELIEF VALVE
-GAS
JETS
-GAS-JET CONTROL VALVES
Fig. IV-38. Gas-jet attitude control system
block diagram
Table IV-10. Flight control modes
Control loop
Flight phase
Modes
Remarks
Attitude control loop
Pitch and yaw
Coast
Rate
Inertial
Celestial
Gas jet matrix signals
Thrust
Inertial
Lunar radar
Vernier engine matrix signals
Roll
Coast
Rote
Inertial
Celestial
Leg 1 gas jet signals
Thrust
Inertial
Vernier Engine 1 gimbal command
Acceleration control loop
Thrust axis
Thrust (midcourse)
Thrust (terminal descent)
Inertial (with accelerometer)
Inertial (with accelerometer)
Nominal 3.22 ft/sec=
Minimum 4.77 ft/sec"
Maximum 12.56 ft/sec'
Velocity control loop
Thrust axis
Thrust
Lunar radar
Command segment signals to 43 ft altitude
Constant 5 ft/sec velocity signals to 1 4 ft altiti
de
Lateral axis
Thrust
Lunar radar
lateral/angular conversion signals
JPL TECHNICAL REPORT 32-1086
65
The celestial references (Fig. IV-4), the sun and the
star Canopus, are acquired and maintained after the
spacecraft separates from the Centaur stage and after
automatic deployment of the solar panel. The sun is first
acquired by the acquisition sun sensor during a space-
craft roll maneuver which is automatically initiated at
completion of solar panel deployment. The 10-deg wide
by 196-deg fan-shaped field of view of the acqusition sun
sensor includes the Z-axis and is centered about the
X-axis. The roll command is terminated after initial sun
acquisition, and a yaw command is initiated which
allows the narrow-view primary sun sensor to acquire
and lock on the sun. Automatic Canopus acquisition and
lock-on are normally achieved after initiation of a roll
command from earth. This occurs because the Canopus
sensor angle is preset with respect to the primary sun
sensor prior to launch for each mission. Star mapping for
Canopus verification is achieved by commanding the
spacecraft to roll while maintaining sun lock. A second-
ary sun sensor, mounted on the solar panel, provides a
backup for manual acquisition of the sun if the auto-
matic sequence fails.
The transit phase is performed with the spacecraft in
the celestial-referenced mode except for the initial rate-
stabilization, midcourse, and terminal descent maneuvers.
The midcourse and main retro orientation maneuvers are
achieved in the inertial mode. Acceleration control is
used for controlling the magnitude of the midcourse
velocity increment. During the interval from retro case
separation to initiation of radar velocity control, accelera-
tion control is used to control the descent along the
spacecraft thrust axis, and velocity control is used for
pitch and yaw control to align the spacecraft thrust axis
with the velocity vector.
The lunar reference is first established by a signal from
an AMR subsystem when the spacecraft is nominally
60 miles above the lunar surface. (Refer to Section IV-G
for discussion of radar control during the standard
terminal descent phase.)
1. Launch Phase
At launch, with the spacecraft in the 550-bit/sec mode,
the gyro temperatures were as follows; roll, 172. 3°F;
pitch, 170.2°F; and yaw, 172.0°F. (In the 1100-bit/sec
mode, the gyro temperatures would read about 8°F
higher.) The gyro temperatures stabilized about 1 hr
35 min after liftoff and remained stabilized until flight
control power was first turned off about 22 hr after
launch for post-midcourse vernier engine firings.
The gas-jet attitude control system contained a charge
of 4.5 lb of nitrogen gas based upon a prelaunch pressure
of 4586 psig at 79.74 °F.
An anomaly occurred 35 sec prior to spacecraft separa-
tion from the Centaur (simultaneously with the legs
extend signal) when the flight control reverted to inertial
mode from rate mode. The flight control was automati-
cally returned to rate mode at separation.
Separation of the spacecraft from the Centaur appeared
normal. Direct analysis of the spacecraft gyro outputs was
hindered because of an initial position offset of the pitch
and yaw gyros of approximately 4 deg (due to the 35-sec
abnormal inertial mode operation) at the time that rate
mode was initiated by electrical disconnect (5.5 sec prior
to spacecraft separation).
An analysis of spacecraft separation is continuing
which uses a 6-degrees-of-freedom analog simulation.
Initial conditions (Centaur residual rates) are programmed
in, and body tipoff rates are simulated by applying short-
duration external torques. Analog gyro outputs for vari-
ous body tipolf rates are correlated with the telemetered
spacecraft gyro outputs. Preliminary results of this analy-
sis indicate the pitch and yaw tipoff rates to have been
less than 0.5 deg/sec and well within the capability of
the attitude control system (also refer to Section III).
2. Sun Acquisition
Telemetered data indicated that the automatic sun
acquisition occurred properly. Upon completion of solar
panel deployment, the spacecraft performed a negative
roll for 72 deg until the acquisition sun sensor became
illuminated. The spacecraft then stopped and started a
positive yaw turn. The spacecraft continued to yaw for
16 deg before the primary sun sensor indicated sun
lock-on. The sun acquisition and lock-on sequence took
about 3 min.
3. Star Acquisition
Six hours after launch, cruise mode, manual delay
mode, and positive angle maneuver commands were sent
to the spacecraft to initiate the star mapping sequence.
These commands were followed by the sun and roll
command to start the actual roll maneuver in a positive
direction at a rate of 0.5 deg/sec.
Earthshine had been indicated by a high-intensity
signal from the star sensor after the sun was acquired
66
JPL TECHNICAL REPORT 32-1086
(deflected light entering within dz35 deg of the star
sensor's Hne of sight will yield a star intensity indication).
Based on that indication, it was surmised that the roll
angle of Canopus, with respect to the star sensor's line of
sight before starting the star mapping roll, was either
— 60 or — 120 deg. The analog traces of the star angle
and star intensity signals recorded during the roll indi-
cated only three distinguishable stars plus a 20-deg-wide
low-intensity signal (identified as the moon) and a 48-deg-
wide variable-high-intensity signal (identified as the
earth). The angular spacing of the signals was compared
with the previously calculated star, earth, and moon
angles, thus permitting positive identification of Canopus
and other stars. Subsequent analysis of bulk printer data
identified a fourth star that was not distinguishable on
the analog recording because of analog dropouts and
noise. Because of the small number of identifiable stars
and the variable-intensity indications from the moon and
earth, it was decided to perform a second complete
revolution for star mapping prior to acquiring Canopus.
The first 360-deg map was made using Omniantenna B,
while the second map and acquisition were made using
Omniantenna A.
Both the star angle and star intensity signals appeared
normal when the sensor rolled past the stars Ras Alhague,
Shaula, Canopus, and Zeta C Majoris. When rolling past
the moon and earth the star angle signal was very erratic
about a zero value. The star intensity signal increased
uniformly when rolling past the moon, but was high and
varying when rolling past the earth with its varying
surface brightness.
Prior to launch, a sun filter having 20% increased
filtering action was installed. This provided the star
sensor with a dimmer sun signal, which was expected to
permit 20% increased star intensity values. A comparison
of actual indicated intensities vs predicted intensities is
given in Table IV-11. This comparison indicates either
that the sensitivity of the intensity signal was increased
more than the planned 20% or that star brightness in
space is only known to within about 20%.
No Canopus lock-on signal was received when the
sensor rolled past Canopus. Therefore, it was necessary to
use the manual lock-on command to lock on Canopus.
A cruise mode command was sent at the time Canopus
was in the field of view during the third revolution.
Spacecraft roll was stopped over 2 deg past the center of
the field of view. The subsequent manual lock-on com-
mand caused the roll error angle to null to zero in approx-
Table IV-1 1 . Sfar angles and intensities:
indicated vs predicted
Roll
angle,
deg
Source
Angle from Canopus,
deg
Relativt
intensity
Indicated
Predicted
Indicated
Predicted
1 00.0
(Start of roll)
Ras Alhague
First star mapping roll
-140.0
-139.5
0.971
0.86
135.6
Moon
-104.4
-102.7
1.142
—
150.5
Shaula
-89.5
-89.7
1.430
1.56
240.0
Canopus
4.995
5.00
262.5
ZeloC
Majoris
-f 22.5
4-22.7
0.854
0.72
325.2
Earth
4-85.2
4-85.0
4.916
-
460.0
Ras Alhague
Second star mapping roll
4-220.0
4-220.5
0.976
0.86
496.2
Moon
4-256.2
4-257.3
1.127
—
510.5
Shaula
4-270.5
4-270.3
1.372
1.56
600.0
Canopus
4.995
5.00
622.5
ZetoC
Majoris
4-22.5
4-22.7
0.844
0.72
684.6
Earth
4-84.6
4-85.0
4.936
-
820.0
Ras Alhague
Canopus acquisition roll
4-220.0
4-220.5
0.937
0.86
857.4
Moon
4-257.4
4-257.3
1.137
—
870.5
Shaula
4-270.5
4-270.3
1.352
1.56
960.0
Canopus
4.995
5.00
—
No Star
-
-
0.77
0.66
imately 40 sec. Since Canopus did not yield a usable
lock-on signal, credence is given to the suggestion that
the sensitivity of the intensity signal was indeed in-
creased more than the planned 20% . However, it is also
possible that Canopus appears brighter in space than
predicted.
The Canopus sensor was modified to solve a window
fogging problem which occurred during early solar-
thermal-vacuum (STV) tesing. The thermal paint pattern
on the sun shade was changed to increase the sensor
temperature, and silicone grease was removed from
gaskets to reduce the possibility of contamination. A com-
parison of the Canopus sensor temperatures of Surveyors I
and 77 just prior to the midcourse velocity correction
indicates that the Canopus sensor temperature was 84.7° F
JPL TECHNICAL REPORT 32-1086
67
for Surveyor II compared to a temperature of 78°F for
Surveyor I. There was no evidence of any window fog-
ging of Surveyor II.
4. Gyro Drift and Limit Cycle Dead Bands — Cruise Mode
A single gyro drift check was made on September 20
between 19:26:24 (about 7 hr after launch) and 21:35:22
GMT. Drift checks are necessary to ascertain whether a
correction factor needs to be included in the premid-
course and terminal maneuver computations to correct for
excessive gyro drift. The drift rates were nominal as
follows (in deg/hr): pitch, +0.25; yaw, +1.0; roll, -0.79.
The peak-to-peak single-axis optical deadband mea-
surements based upon the data were as follows (in deg) :
pitch, 0.25; yaw, 0.35; and roll, 0.5.
5. Premidcoorse Maneuvers
The premidcourse attitude maneuvers ( + 75.4 deg roll
and +110.6 deg yaw) were accomplished satisfactorily.
The alternate or backup maneuvers in the event the
spacecraft would not roll were a pitch of —111.1 deg and
a yaw of +13.7 deg. In the event that the spacecraft
would not yaw, the maneuvers- were an additional roll
of -89.9 deg followed by a pitch of -110.5 deg.
6. Midcourse Velocity Correction
When midcourse thrusting was commanded at 05:00:02
GMT on September 21, Vernier Engine 3 failed to ignite
and the spacecraft tumbled, saturating the gyro error
signals, which indicated a minus pitch, plus yaw, and
minus roll. After termination of midcourse thrusting, the
pitch and yaw gyro error signals varied from plus to
minus with a period of about 13 sec until rate mode was
commanded on, at which time the signals returned to
their original saturated positions of minus pitch, plus yaw,
and minus roll. The gas-jet amplifiers were not inhibited
immediately after thrust termination because it was
believed that the tumble rate might be small enough for
the gas-jet attitude control system to dampen it out with-
out using an excessive amount of nitrogen. The tele-
communications analyst later estimated the maximum
tumble rate to be approximately 1.22 rev/sec (a period
less than 1 sec) based upon DSIF AGC data. When it
became obvious that the gas system would not be able to
remove the angular rates, the gas jet amplifiers were
inhibited at 05:14:29, with an estimated 2.16 lb of nitro-
gen remaining and a spacecraft tumble rate of approxi-
mately 0.97 rev/sec. Table IV-12 shows the expected and
actual nitrogen consumption during the mission as
derived from pressure and temperature data.
If it is assumed that the acceleration amplifier was
saturated ( - 14.3-dc diflFerence based on Surveyor II test
Table IV-12. Nitrogen gas consumption
Time,
GMT
Event
pressure,
psig
temperature,
°F
N,
remaining,
lb
Actual N
- usage
Expected
N2
usage
For
event
Cumulative
total
September 20
10:35
Prelaunch
4586
80.0
4.50
13:02:58
Rote stabilization
and sun acqui-
sition
4480
77.4
4.45
0.05
0.05
0.100
19:24:58
Conopus acquisition
4100
47.3
4.44
0.01
0.06
0.159
September 21
04:00:00
Cruise
3970
40.2
4,42
0.02
0.08
0.182
05:14:29
Gas lets
inhibited
06:09:47
Attitude maneuver
and midcourse
thrusting
1630
6.8
2.16
2.26
2.34
0.288
10:13:37
Post-midcourse
1760
36.7
2.16
2.34
68
JPL TECHN/CAt REPORT 32-1086
data) and the pitch and yaw shaping ampUfier was also
saturated (11.04-dc difference) almost immediately after
the start of vernier ignition, then the vernier engine
thrust level commands may be calculated as follows
using nominal vernier amplifier gains:
Engine 1
+ 150 ma due to acceleration error
— 175 ma due to pitch gyro error
+ 64 ma due to yaw gyro error
+ 39 ma
Engine 2
+ 150 ma due to acceleration error
+ 36 ma due to pitch gyro error
— 192 ma due to yaw gyro error
— 6 ma
Engine 3
+ 150 ma due to acceleration error
+ 148 ma due to pitch gyro error
+ 127 ma due to yaw gyro error
+ 80 ma (maximum capability of vernier amplifier)
These values correspond to approximately 85 lb of
thrust on Engine 1 and 65 lb on Engine 2. The low out-
put of the acceleration amplifier during midcourse was
expected in view of the fact that not enough engine thrust
was being generated to achieve the 0.1 g being com-
manded. During all subsequent firings, the output of the
acceleration amplifier always indicated high, presumably
because of the tumbling effect on the accelerometer.
7. Post-Midcourse Attempts to Ignite Vernier Engine 3
Two attempts were made to ignite Engine 3, beginning
at 07:28:27 while the spacecraft was in the normal mid-
course velocity correction mode. The engine burn time
was 2.0 sec in each case. Vernier Engines 1 and 2 ignited
normally, but no indication that Engine 3 ignited was
received from the strain-gage telemetry signal, the tem-
perature sensor, or the pitch and yaw gyro error signals.
Next, the vernier engines were commanded on five
times in succession, for 0.2 sec every 5 min, followed by
a 2.0-sec thrust period. This sequence was performed five
times. After the fifth sequence, a 2.5-sec burst was made at
the high thrust levels normally used to assist in separating
the retro engine after burnout. In order to obtain the
desired thrust levels, which occur only between retro
engine burnout and delayed burnout, without causing
retro ejection, it was necessary to set the retro burnout
and retro eject latches high, with thrust phase power off.
Following another sequence of five 0.2-sec firings start-
ing at 07:44:56, a final attempt was made to open the fuel
pressure regulator valve by commanding high thrust for
20 sec. The gas jets were enabled as part of this command
sequence. The calculated vernier engine thrust commands
for this test were as follows:
Engine 1
+ 48 ma due to high thrust command
— 175 ma due to pitch gyro
+ 64 ma due to yaw gyro
— 63 ma
Engine 2
+ 48 ma due to high thrust command
+ 36 ma due to pitch gyro
— 192 ma due to yaw gyro
— 80 ma (maximum capability of vernier amplifier)
Engine 3
+ 48 ma due to thrust command
+ 148 ma due to pitch gyro
-\r 127 ma due to yaw gyro
+ 80 ma (maximum capability of vernier amplifier)
The difference between the calculated command for
Vernier Engine 1 and the actual can be accounted for by
the expected variation in shaping amplifier saturation
voltages. The ranges of differential saturation voltages
for Surveyor 11 are shown below:
Acceleration amplifier, + 15.43 and — 14.27 v
Pitch shaping amplifier, + 10.63 to + 16.88 and
-11.04 to -16.77 V
Yaw shaping amplifier, + 10.84 to + 16.72 and
-11.04 to -17.10 V
8. Retro Ignition
At retro ignition (09:34:28.65 on September 22), the
indicated thrust level rose from 7.14 to 10.27 g and
jn TECHNICAL REPORT 32-1086
69
remained nominally at that level until 09:34:48, at which
time the indicated acceleration began increasing until
09:35:00, at which time the acceleration was 11.72 g.
Shortly thereafter all spacecraft contact was lost. The
estimated spin rate at time of data loss was 1.85 rev/sec.
G. Radar
Two radar devices, the altitude marking radar (AMR)
and the radar altimeter and doppler velocity sensor
(RADVS), are employed on the Surveyor spacecraft for
use during the terminal descent phase.
1. Altitude Marking Radar
The AMR (Fig. IV-39) is a pulse-type fixed- range
measuring radar which provides a mark signal at a slant
range from the lunar surface that can be preset between
52 and 70 miles. The mark signal is used by the flight
control subsystem to initiate the automatic operations for
spacecraft terminal descent.
The AMR operates at a frequency of 9.3 gc. The mark
range is obtained by use of dual-channel video gating
(early and late gate signals). The early and late gates are
adjacent at the preset range (60 miles for Surveyor II)
so that, as the spacecraft approaches the lunar surface, the
video return becomes equally distributed between these
two gates. When the main lobe return is of equal magni-
tude in both gates and of such an amplitude to overcome
a preset bias, the mark signal is generated and initiates
the automatic operations for spacecraft terminal descent.
The AMR mounts in the retro rocket nozzle and is
retained by friction clasps around the nozzle flange with
spring washers between the AMR and the flange. When
the retro rocket is ignited, the gas generated by the
ignitor develops sufficient pressure to eject the AMR from
the nozzle. The AMR draws power from 22 vdc through
a breakaway plug that also carries input commands, the
output mark signal, and telemetry information.
The AMR was not turned on during the Surveyor 11
mission. Temperatures were normal during the flight
SYNCHRONIZER
POWER ON
(COMMAND)'
18-29 V
RANGE
GATE
PULSES
MODULATOR
TRIGGER,
350 pps
TIMING
PULSE,
350 pps
MODULATOR
350 pps
3.2 /isec
30.0 V dc
»•
RF ASSEMBLY
MAGNETRON
CIRCULATOR
TR TUBE
3.5 /isec
350 pps
1.5 kw PEAK
9.3 kmc
ENABLE SIGNAL.
(COMMAND)
ELECTRICAL
CONVERSION
UNIT
LOCAL
OSCILLATOR,
72.89 mc
34 -db GAIN
3.6° BEAMWIDTH
30- in. DIAM
ANTENNA
MAIN BANK
BLANKING SIGNAL
VIDEO
PROCESSOR
VIDEO
9.33 kmc
9.3 kmc
IF AMPLIFIER
VIDEO DETECTOR
l07-db GAIN
30 mc
MIXER
5-mc PASS BAND
5.5 -db NOISE FIGURE
60-mi.*M/?A' SIGNAL TO
FLIGHT CONTROL
70
Fig. IV-39. Altitude marking radar functional diagram
JPt TECHN/CAl REPORT 32-7086
except after tumbling. The antenna then received in-
creased thermal radiation from the sun and the tempera-
ture at the edge of the antenna increased to +155°F
(normal temperature is -185°F).
2. Radar Altimeter and Doppler Velocity Sensor
The RADVS (Fig. IV-40) functions in the flight control
subsystem to provide three-axis velocity, range, and
altitude mark signals for flight control during the main
retro and vernier phases of terminal descent. The RADVS
consists of a doppler velocity sensor (DVS), which com-
putes velocity along the spacecraft X, Y, and Z axes, and
a radar altimeter (RA), which computes slant range from
40,000 to 14 ft and generates 1000-ft and 14-ft mark
signals. The RADVS comprises five assemblies: (1) kly-
stron power supply /modulator (KPSM), which contains
the RA and DVS klystrons, klystron power supplies, and
altimeter modulator, (2) altimeter/velocity sensor an-
tenna, which contains beams 1 and 4 transmitting and
receiving antennas and preamplifiers, (3) velocity sensing
antenna, which contains beams 2 and 3 transmitting
antennas and preamplifiers, (4) RADVS signal data con-
verter, which consists of the electronics to convert dop-
pler shift signals into dc analog signals, and (5) inter-
connecting waveguide. The RADVS is turned on at about
50 miles above the lunar surface and is turned off at
about 13 ft.
a. Doppler velocity sensor. The doppler velocity sen-
sor (DVS) operates on the principle that a reflected
signal has a doppler frequency shift proportional to the
approaching velocity. The reflected signal frequency is
higher than the transmitted frequency for the closing
condition. Three beams directed toward the lunar surface
enable velocities in an orthogonal coordinate system to
be determined.
The KPSM provides an unmodulated DVS klystron
output at a frequency of 13.3 kmc. This output is fed
equally to the DVSl, DVS2, and DVS3 antennas. The
RADVS velocity sensor antenna unit and the altimeter
velocity sensor antenna unit provide both transmitting
and receiving antennas for all three beams. The reflected
signals are mixed with a small portion of the transmitted
frequency at two points % wavelength apart for phase
determination, detected, and amplified by variable-gain
amplifiers providing 40, 65, or 90 db of amplification,
depending on received signal strength. The preamp out-
put signals consist of two doppler frequencies, shifted
by -H transmitted wavelength, and preamp gain-state
signals for each beam. The signals are routed to the
trackers in the RADVS signal data converter.
The Dl through D3 trackers in the signal data con-
verter are similar in their operation. Each provides an
output which is 600 kc plus the doppler frequency for
approaching doppler shifts. If no doppler signal is pres-
ent, the tracker will operate in search mode, scanning
frequencies between 82 kc and 800 cps before retro bum-
out, or between 22 kc and 800 cps after retro burnout.
When a doppler shift is obtained, the tracker will operate
as described above and initiate a lock-on signal. The
tracker also determines amplitude of the reflected signal
and routes this information to the signal processing
electronics for telemetry.
The velocity converter combines tracker output signals
Dl through D, to obtain dc analog signals corresponding
to the spacecraft X, Y, and Z velocities; D, + D,, is also
sent to the altimeter converter to compute range.
Range mark, reliability, and reference circuits produce
a reliable operate signal if D, through D, lock-on signals
are present, or if any of these signals are present 3 sec
after retro burnout. The reliable operate DVS signal is
routed to the flight control electronics and to the signal
processing electronics telemetry.
b. Radar altimeter. Slant range is determined by
measuring the reflection time delay between the trans-
mitted and received signals. The transmitted signal is
frequency-modulated at a changing rate so that return
signals can be identified.
The RF signal is radiated, and the reflected signal is
received by the altimeter/velocity sensor antenna. The
received signal is mixed with two samples of transmitted
energy % wavelength apart, detected, and amplified by
40, 60, or 80 db in the altimeter preamp, depending on
signal strength. The signals produced are difference fre-
quencies resulting from the time lag between transmitted
and received signals of a known shift rate, coupled with
an additional doppler frequency shift because of the
spacecraft velocity.
The altimeter tracker in the signal data converter
accepts doppler shift signals and gain-state signals from
the altimeter/velocity sensor antenna and converts these
into a signal which is 600 kc plus the range frequency
plus the doppler frequency. This signal is routed to the
altimeter converter for range dc analog signal generation.
The range mark, reliability, and reference circuits
produce the l(X)0-ft mark signal and the 14-ft mark signal
JPL TECHNICAL REPORT 32-1086
71
VELOCITY SENSOR ANTENNA
'>
RECEIVER
TRANS-
NfllTTER
^>
^8)-
SIGNAL DATA CONVERTER
PREAMPLIFIERS
^H^
CRYSTAL
MIXER
SAMPLER
KLYSTRON POWER
SUPPLY AND
MODULATOR
MODULATOR
TRANSMITTER
'>
RECEIVER
'>
TRANSMITTER
^>
DOPPLER
VELOCITY
TRANSMITTER
13.3 gc
RADAR
ALTITUDE
TRANSMITTER
w [~| 12.9 gc
<8h->
RECEIVER
CRYSTAL
MIXER
PREAMPLIFIERS
^ »->
RADAR ALTITUDE VELOCITY
SENSOR ANTENNA
DOPPLER
TRACKER No. 3
DOPPLER
TRACKER No. 2
CONVERTER
■ V, 50 mv/tps
RELIABLE
' OPERATE
CONVERTER
■Vy 50 mv/fps
RELIABLE
OPERATE
DOPPLER
TRACKER No. I
RELIABLE
OPERATE
(D, D2 Dj)
Do —
Dj —
D4 —
H^ RELIABLE OPERATE
•) DOPPLER VELOCITY
PROGRAMMED
-^
SENSOR
(D, D3 D4)
RELIABLE OPERATE
RADAR ALTITUDE
rN(D, D2)+(D| D5)+(D2D3)
\y CONDITIONAL RELIABLE
OPERATE DOPPLER
VELOCITY SENSOR
CONVERTER
RELIABLE
OPERATE
DOPPLER
TRACKER No. 4
•V, 50 mv/fps
RANGE MARK
" 1000 ft
•14 ft
RANGE
CONVERTER
•Itnv/ft (R>IK)
>• RANGE
•20mv/ft(R<IK)J
RELIABLE
■ OPERATE
Fig. IV-40. Simplified RADVS functional block diagram
from the rati^'c signal generated by the altimeter converter
where the doppler velocity V.- is subtracted giving the
true range.
The ran^e mark and rclkible signals are routed to flight
control electronics. The signals are used to rescale the
ran<ie signal, for vernier engine shutoff and to indicate
whether or not the ratine signal is reliable. The reliable
operate signal is also routed to signal processing for
transmission to DSIF.
c. RADVS performance. The HADVS was turned on
prior to the main retro firing (near the end of the mission)
for a battery load test. RADVS power was on for a total
of 10 min 12 sec. Unfortunately, the telemetry mode was
72
JPL TECHNICAL REPORT 32-J086
•not proper for RADVS operational data. The only data
received were temperatures. Table IV-13 presents a
listing of some of the RADVS temperatures during the
•mission. Figures IV-41 through IV-44 present the plots of
the temperature rises during power on. The antennas and
klystron power supply modulator (KPSM) starting tem-
peratures were within the expected ranges. However, the
signal data converter (SDC) temperature (-84°F) was
much too low for reliable operation. Normally, the lowest
expected starting temperature is -21°F. Since data other
than temperatures is not available, an analysis of SDC
operation cannot be made. The temperature rise does indi-
cate that at least the power supply in the SDC operated.
Surveyors I and U unit temperature data was used
to predict the maximum temperature attained on the
Surveyor I KPSM as shown in Fig. IV-41. This prediction
was based on a parallel temperature rise for both units.
Table IV-13. Surveyor // RADVS temperature data
Time, GMT
(September 22)
AMR
electronics,
°F
AMR
antenna,
°F
KPSM,
"F
SDC,
"F
DVS,
"F
RADVS,
"F
Comments
03:57:39
97.58
149.5
40.11
-81.25
63.0
18.0
Last Goldstone pass
09:13:00
36.30
137.3
33.50
-84.00
36.1
98.1
09:19:57
RADVS on
09:26:00
96.30
138.7
103.00
-33.28
48.0
95.7
09:30:09
96.30
138.7
128.30
1.00
50.1
101.0
RADVS off
09:31:19
96.30
138.7
123.80
2.70
51.9
102.7
130
120
100
§ 80
lU
I 70
liJ
60
50
30
rA
M
yAj^
>
A
r
-y
H
)
/
C
/
^s
/
/
sua
yero/f
/ /
'RVEYOl
/
/ 1 SL
/
/
?7
7-RAD\
S ON
//
\,
/ /PREDICTED (DOTTED PORTION OF CURVE)
\
y
/4
\ ,-
i-l
V RADVS OFF
\
? \J
\
\
1
MEASURED
\
1
i
S '
t
5
5 1
r (
J '
> 1
1
1 1
2
3 1
4 1
5 1
6 1
7 18
TIME, min
Fig. IV-41. Klystron power supply modulator temperature
JPL TECHNICAL REPORT 32-1086
73
15 16 17 18
TIME, min.
Fig. IV-42. Signal data converter temperature
53
51
^
>
-
^
^ <
A
(-
47
{
S-
UJ
q: 45
=)
c
3 /
41
/
/
RADVS
ON— 1
-^^^
f
39
/
RADVS OFF— \
37
3 1
J
y
V
[^
\
35
1
2
3
4
i
»
'
<
>
1
2
3
4
5
6
7 18
TIME, min
Fig. IV-43. Doppler velocity sensor temperature
74
JPt TECHNICAL REPORT 32-1086
102
100
^
3
^ <»«
^
)
UI
a:
u
t 96
;
V
q:
UI
a.
z
UI 94
/
1-
92
RADV
S ON-
A^
RADVS OFF-
-^
90
3 \
-^
\
k
Prediction variations of dzlO deg are expected since the
klystron-to-heat-sink transfer characteristics differ for
each unit. An attempt to predict the temperature rise in
S 6 7 8 9 10 II 12 13
TIME, min
Fig. IV-44. Altitude marking radar temperature
H. Telecommunications
14
17 18
The Surveyor telecommunications subsystem contains
the SDC of Surveyor I was not made, owing to the radio, signal processing, and command decoding equip-
vastly different starting temperatures. ment, to provide (1) a method of telemetering information
OMNIANTENNA A
RF IN
RF in/out
HIGH-GAIN
ANTENNA
(PLANAR
ARRAY)
RECEIVER
A
WB INPUT
NB INPUT
COMMAND
OUT
RECEIVER/
DECODER
SELECTOR
TRANSPONDER
INTERCONNECTION
WB INPUT
NB INPUT
TRANSPONDER
INTERCONNECTION
COMMAND OUT
.TO
DECODER
Fig. IV-45. Radio subsystem block diagram
JPL TECHNICAL REPORT 32-1086
75
to the earth, (2) the capability of receiving and processing
commands to the spacecraft, and (3) angle-tracking one-
or two-way doppler for orbit determination.
1. Radio Subsystem
The radio subsystem utilized on the Surveyor space-
craft is as shown in Fig. IV-45. Dual receivers, transmit-
ters, and antennas were originally meant to provide
redundancy for added reliability, although as arranged
this is not completely true because of switching limita-
tions. Each receiver is permanently connected to its
corresponding antenna and transmitter.
Both receivers are identical crystal-controlled double-
conversion units which operate continuously (cannot be
commanded off). Each unit is capable of operation in an
automatic frequency control (AFC) mode or an automatic
phase control (APC) mode. The receivers provide two
necessary spacecraft functions: the detection and pro-.,
cessing of commands from the ground stations for space-
craft control (AFC and APC modes), and the phase-
coherent spacecraft-to-earth signal required for doppler .
tracking (APC mode).
Transmitters A and B are identical units which provide
the spacecraft-to-earth link for telemetry and doppler
tracking information. The transmitters are commanded
on (one at a time) from the ground stations. Each unit
contains two crystal-controlled oscillators (wideband for
TV data; narrowband for engineering data), which can
be commanded on at will. The transmitters may also be
commanded to operate at either 100 mw or 10 w of output
power.
Two identical transponder interconnections permit
each transmitter to be operated, on command, in a trans-
ponder mode. In the transponder mode, a transmitter
-70
-74
-78
o:
UJ
o
0.
o
>
u
o
o
1-
FIRST ACQUISITION
-86
-90
-94
LAST NflEASUREMENT 34 min
PRIOR TO LAUNCH
NO CALIBRATED AGC CURVES
TO SHOW SIGNAL LEVEL
GREATER THAN THIS
NOTE:
COMMAND THRESHOLD -114 dbm.
PHASE LOCK THRESHOLD -131.5 dbm.
ALTHOUGH THE CALIBRATED AGC CURVES
INDICATED NO SIGNAL LEVEL ABOVE -76.6 dbm
(AT + SOT), THERE WERE BCD CHANGES
FROM 100 TO 1 13 DURING THE I hr 40 mIn
(APPROX) PERIOD SHOWN ABOVE AS A
STRAIGHT LINE.
TYPICALLY A I3-8CD CHANGE IS APPROXIMATELY
EQUIVALENT TO A 3-db CHANGE IN SIGNAL LEVEL.
-20
20
40 60 80
TIME FROM LIFTOFF, min
100
120
140
160
Fig. IV-46. Total received power, Receiver A
76
JPL TECHNICAL REPORT 32-7086
is operated with the corresponding receiver voltage-
controlled oscillator to provide coherent signals when
two-way doppler tracking data is required.
Three antennas are utilized on the Surveyor spacecraft.
Two antennas are omnidirectional units which provide
receive-transmit capability for the spacecraft. The third
antenna is a high-gain (27-db) directional unit which is
used primarily for transmission of wideband information.
Either transmitter may be commanded to operate through
any one of the spacecraft antennas as desired.
no evidence of malfunction in the receiver command or
doppler performance during the flight.
The amplitudes of RF signals received at both space-
craft receivers and the DSIF stations were examined over
the period of standard flight. The results are presented in
Figs. IV-46 through IV-51 along with the predicted
signal strengths. No attempts were made to correct for
Receiver B AGC errors in that preflight calibration data
was used. The AGC error becomes very evident in
Fig. IV-49.
The radio subsystem on Surveyor II performed well
during flight, except for Receiver B AGC telemetry.
Receiver B was used as the prime spacecraft command
receiver, and in conjunction with Transmitter B provided
the necessary spacecraft doppler information. There was
Three hours prior to launch, the receiver AGC problem
was first questioned but not identified. Spacecraft telem-
etry indicated that Receiver A was receiving 25-db more
signal than Receiver B. Previous data recorded 9 hr prior
to launch indicated only a 9-db difference. The fact that
-60
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COMMAND THRESHOLD -114 dbm.
PHASE LOCK THRESHOLD -131.5 dbm.
PREDICTED AGC
ENVELOPE
PREDICTED AGC
(NOMINAL)
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SEPTEMBER 21
08:00 10:00
Fig. IV-47. Receiver A AGC vs GMT
JPL TECHNICAL REPORT 32-1086
77
60 80 100
TIME FROM LIFTOFF, min
Fig. IV-48. Total received power, Receiver B
78
JPL TECHNICAL REPORT 32-1086
-70
-80
-90 -
T
T
PREDICTED AGC
ENVELOPE
COMMAND THRESHOLD -114 dbm.
PHASE LOCK THRESHOLD -130 dbm.
SIGNAL LEVELS WERE DERIVED
FROM TELEMETERED AGC.
RECEIVER "b"AGC WAS NOT
ACCURATE.
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22:00
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02:00
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SEPTEMBER 21
Fig. IV-49. Receiver B AGC vs GMT
JPL TECHNICAL REPORT 32-1086
79
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1100
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TIME FROM LIFTOFF, min
Fig. IV-50. DSS 51 received RF power vs time
the gantry was removed between the time.s of the two
measurements leads to the conclusion that a null in the
Receiver B antenna pattern produced the 25-db difference.
After launch, the lack of information on spacecraft
orientation concealed the problem until Canopus was
acquired (some 6 or 7 hr after liftoff). During the Canopus
search, Receiver B never indicated signal levels as high
as expected. After Canopus acquisition, the signal level to
I\eceiver B was 11 db less than the signal to Receiver A.
Omniantenna patterns were reviewed, and they indicated
that during Canopus lock the signal to Receiver B .should
be stronger than the signal to Receiver A by greater than
10 db. Antenna contour maps are presented in Figs. IV-52
through IV-54. The results of the Canopus search are
plotted in Figs. IV-55 through IV-57.
About 13 hr after launch, a special test was performed
on the spacecraft receivers by lowering the DSIF station
up-link power in 2-db steps while monitoring spacecraft
telemetry. The results are presented in Table IV-14 and
indicate the inadequacy of calibration data on Receiver B
AGC. The first eight 2-db steps resulted in Receiver B-
telemetry changes of about 3 db. Telemetry information
was further invalidated by the fact that receiver indexing
did not occur until Receiver B telemetry indicated a
signal level of — 138 dbm.
(Preflight data indicated that Receiver B should index
with a signal of -124 dbm.) It was not clear whether
Receiver A or Receiver B actually produced the indexing,
since the signal to Receiver A was close to its indexing
level, and the telemetry data for Receiver B AGC was
in error.
Preliminary investigations indicated that the signal
processing equipment functioned normally and that the
problem was in the receiver itself. If the spacecraft failure
had not occurred at midcourse, further tests on Receiver B
would have been performed and could have determined
if its performance was actually degraded or only mis-
represented by the telemetry.
The spacecraft radio system performed well during the
nonstandard portion of the flight. When midcourse motor
firing was initiated, the tumbling of the spacecraft was
immediately apparent from the station receiver AGC.
At the termination of firing the midcourse motors, the
spacecraft rate of tumble was observed to be about
1.22 rev/sec, as indicated by receiver AGC. The gas
stabilization jets reduced the tumbling rate to about
0.97 rev/sec in what appeared to be a fairly linear man-
ner. The down-link signal amplitude initially varied as
much as 17 dbm to as little as 3 dbm because the space-
craft rolled and tumbled in a periodic manner. The gas-
jet firing stabilized the spacecraft antenna motion relative
to the eaith-spacecraft vector, thus producing fairly
constant signal amplitude variations as the spacecraft
tumbled. Typical signal amplitude variations during the
early phase of the flight were 10 dbm with Omni-
antenna B and 15 dbm with Omniantenna A. As the
tumbling rate increased with attempts to fire Engine 3,
the station receiver AGC variations appeared to be less;
the slow time constant of the AGC circuit would have
damped out some of the variations.
Each time an attempt was made to fire Engine 3, an
increased rate of AGC variation was noted. The lowest
rate notc-d after midcourse was about 0.8 rev/sec and was
recorded when the gas jets were inhibited prior to
attempts at firing the malfunctioning engine. The many
80
JPL TECHNICAL REPORT 32-1086
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-170
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TRANSMITTER B
HIGH POWER
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Fig. IV-51 . DSS total received power vs GMT
attempts at short engine burns increased the AGC vari-
ations to a rate of about 1.5 rev/sec just prior to the long
bum of 20 sec. The 20-sec burn increased the rate to
about 2.3 rev/sec, which was shghtly decreased when
the hehum was dumped. Firing of the retro engine
reduced the AGC rate change to below 2 rev/sec just
prior to loss of the down-link RF signal.
The ground stations experienced difficulty in maintain-
ing decommutator lock until their receiver bandwidths
were opened up. The bandwidths were first increased
from about 76 cps to the desired 152 cps when in the
152-cps loop noise bandwidth position. The increased
bandwidth was accomplished by physically modifying
one of the receiver modules at the DSIF stations. Later in
the flight, as the spacecraft tumbling rate increased (due
to repeated engine firings) decommutator lock again
became erratic. Further receiver modifications which
increased the loop bandwidth to 300 cps reestablished
firm decommutator lock.
2. Signal Processing Subsystem
The Surveyor signal processing subsystem accepts,
encodes, and prepares for transmission the voltages,
currents, and resistance changes corresponding to various
spacecraft parameters such as events, voltages, tempera-
tures, accelerations, etc.
in TECHNICAL REPORT 32-1086
81
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JPL TECHN/CAl REPORT 32-J086
-30 -60 -90
ANGULAR DISPLACEMENT 4>, deg
Fig. IV-55. Omniantenna A, Receiver A signal level vs angular displacement
-30 -60
ANGULAR DISPLACEMENT <(,, deg
<
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150 120
Fig. IV-56. Omniantenna B, Receiver B signal level vs angular displacement
JPL TECHNICAL REPORT 32-1086
85
<
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-30 -60 -90
ANGULAR DISPLACEMENT <f>, deg
Fig. IV-57. Omniantenna B, Transmitter B signal level vs angular displacement
SPACECRAFT
TELEMETRY
SENSORS
SIX
ENGINEERING
COMMUTATORS
DIGITAL INPUT DATA
ANALOG-
TO-
DIGITAL (A/D)
CONVERTER
SUBCARRIER
OSCILLATOR
ACCELEROMETERS
(NOT USED) AND
STRAIN GAGES
SUBCARRIER
OSCILLATORS
FM
SUMMING
AMPLIFIER
TO TRANSMITTER
*-*
PM
SUMMING
AMPLIFIER
Fig. IV-58. Simplified signal processing subsystem block diagram
86
JPL TECHNICAL REPORT 32-7086
Table IV- 14. Data from in-flight calibration of spacecraft receiver AGC
Time, GMT
September 21
Change in
DSIF signal
level, dbm
Receiver A
Receiver B
BCD"
dbm
Receiver AGC
change, dbm
BCD'
dbm
Receiver AGC change, dbm
Total
Last step
Total
Last step
01:36:39
207
-94.0
215
-104.9
01:37:15
- 2
223
-95.6
1.6
1.6
234
-108.3
3.4
3.4
01:39:48
— 4
242
-97.5
3.4
1.8
255
-112
7.1
3.7
01:42:04
- 6
258
-99.1
5.1
1.7
279
-115.9
11.0
3.9
01:44:21
- 8
278
-101.1
7.1
2.0
301
-119.1
14.2
3.2
01:47:15
-10
299
-103.3
9.3
2.2
321
-122.1
17.2
3.0
01:50:00
-12
318
-105.3
11.3
2.0
338
-124.9
20.0
2.8
01:51:28
-14
336
-107.4
13.4
2.1
353
-128.2
23.3
3.3
01:53:30
-16
356
-110.0
16.0
2.6
364
-131.3
26.4
3.1
01:56:24
-18
374
-113.0
19.0
3.0
371
-133.7
28.8
2.4
01:58:23
-20
388
-115.9
21.9
2.9
376
-135.7
30.8
2.0
02:00:12
-22
399
-118.7
24.7
2.8
379
-137
32.1
1.3
02:04:22
-24"
410
-122.2
28.2
3.5
381
-138
33.1
1.0
* Binary conversion data.
*> Indexing occurred when DSIF signal >
¥as decreased
by 24 dbm.
The signal processor (Fig. IV-58) employs both pulse
code modulation and amplitude-to-frequency-modulation
telemetry techniques to encode spacecraft signals for
frequency- or phase-modulating the spacecraft trans-
mitters.
The input signals to the signal processor are derived
from various voltage or current pickoff points within the
other subsystems as well as from standard telemetry
transducing devices such as strain gages, accelerometers,
temperature transducers, and pressure transducers. These
signals generally are conditioned to standard ranges by
the originating subsystem so that a minimum amount of
signal conditioning is required by the signal processor.
As illustrated in Fig. IV-58, some of the signal inputs
are commutated to the input of the analog-digital
converter while others are applied directly to subcarrier
oscillators. The measurements applied directly are accel-
erometer (not used) and strain gage measurements which
require continuous monitoring over the short intervals
in which they are active.
The commutators apply the majority of telemetry input
signals to the analog-to-digital converter, where they are
converted to a digital word. Binary measurements such as
switch closures or contents of a digital register already
exist in digital form and are therefore routed around the
analog-to-digital converter. In these cases, the commu-
tator supplies an inhibit signal to the analog-to-digital
converter and, by sampling, assembles the digital input
information into 10-bit digital words. The commutators
are comprised of transistor switches and logic circuits
which select the sequence and number of switch closures.
There are six commutator configurations (or modes) used
to satisfy the telemetry requirements other than tele-
vision data for different phases of the mission.
The analog-to-digital converter generates an 11-bit
digital word for each input signal applied to it. Ten bits
of this word describe the voltage level of the input signal
and one bit position is used to introduce a bit for parity
checking by the ground telemetry equipment. When
processing digital input information, the first 10 bits are
not generated; however, the parity bit is still supplied.
JPL TECHNICAL REPORT 32-1086
87
The analog-to-digital converter also supplies commu-
tator advance signals to the commutators at one of six
different rates. These rates enable the signal processor to
supply telemetry information at 4400, HOG, 550, 137.5 and
17.2 bps. The bit rates and commutator modes are
changed by ground commands.
The subcarrier oscillators are voltage controlled oscil-
lators used to provide frequency multiplexing of the
telemetry information. This technique is used to greatly
increase the amount of information transmitted on the
spacecraft carrier frequency.
The summing amplifiers sum the outputs of the sub-
carrier oscillators and apply the composite signal to the
spacecraft transmitters. Two types of summing amplifiers
are employed because of the transmitter's ability to trans-
mit either a phase-modulated or a frequency-modulated
signal.
The signal processing subsystem employs a high degree
of redundancy to insure against loss of vital spacecraft
data. Two analog-to-digital converters, two independent
commutators— the engineering signal processor (ESP) and
auxiliary engineering signal processor (AESP) — and a
wide selection of bit rate (each with the analog-to-digital
converter (ADC) driving a different subcarrier oscillator)
provide a high reliability of the signal processing sub-
system in performing its function.
During the mission, 445 commands effecting changes
in telemetry were received and properly executed by the
signal processor. Of these, 106 were commutator mode
changes and 37 were bit rate changes. This number of
commands is far in excess of what would occur in a
nominal transit sequence. They were executed to obtain
additional data in support of the vernier propulsion
failure. There is no evidence of any abnormal signal
processing behavior due to this increased command den-
sity. In Table IV-15, a comparison of prelaunch and
typical flight data is presented for the signal processing
telemetered functions. These subsystem parameters were
very stable throughout flight. From a review of the
time plots of these channels there is no indication that
anything abnonnal occmred in the subsystem. The tem-
perature in Compartment B stayed at all times within
the 0-125" F range recpiired for proper signal processing
performance.
Approximately 80% of the subsystem was exercised
during the abbreviated mission without the occurrence
of any problems. The (Mily major elements not cheeked
Table IV-15. Typical signal processing
parameter values
Parameter
Prelaunch
Flight •
ESP reference volts, v
4.877
4.882
ESP reference return, v
0.0
0.0
ESP unbalance current, /xa
-1.55
-1.35
ESP full-scale current calibration,
0.4
0.3
% error from nominal
ESP mid-scale current calibration.
0.4
0.4
% error from nominal
ESP zero-scale current calibration,
0.8
0.5
% error from nominal
AESP full-scale current calibration.
0.7
0.8
% error from nominal
AESP mid-scale current calibration.
0.5
0.6
% error from nominal
AESP zero-scale current calibration.
0.5
0.6
% error from nominal
AESP unbalance current, ^o
-1.82
-1.82
out were ADC No. 2, Telemetry Commutator Mode 3,
and acceleroineter and touchdown strain gage channels.
3. Command Decoding Subsystem
Commands are received, detected, and decoded by one
of the four receiver/central command decoder (CCD)
combinations possible in the Purveyor command sub-
system. The selection of the combination is accompli.shed
by stopping the command information from modulating
the up-link radio carrier for V2 sec. Once the selection is
made, the link must be kept locked up continuously by
either sending serial command words or unaddressable
command words (referred to as fill-ins) at the maximum
command rate of 2 words per sec.
The command information is formed into a 24-bit
Manchester-coded digital train and is transmitted in a
PCM/FM/PM modulation scheme to the spacecraft.
When picked up by the spacecraft omniantennas, thc
radio carrier wave is stripped of the command PCM
information by two series FM discriminators and a
Schmitt digitizer. This digital output is then decoded by
the CCD for word sync, bit sync, the 5-bit address and
its complement, and the 5-bit command and its comple-
ment (this latter only for direct commands since the DCs
contain 10 bits of information rather than 5 command bits
and their complements). The CCD then compares the
address with its complement and the command with its
88
JPL TECHNICAL REPORT 32-J086
•complement on a bit-by-bit basis. If the comparisons are
satisfactory, the CCD then selects that one of -the eight
subsystem decoders (SSD's) having the decoded address
•bits as its address, applies power to its command matrix,
and then selects that one of the 32 matrix inputs having
the decoded command bits as its address to issue a 20-
msec pulse which initiates the desired single action.
Those DC commands that are irreversible or extremely
critical are interlocked with a unique command word.
Ten of the DCs and all of the quantitative commands are
in this special category. None of these commands can be
initiated if the interlock command word is not received
immediately prior to the critical command.
The QC's, besides being interlocked, are also treated
somewhat differently by the command subsystem. The
only differences between the DC and QC are: (1) a
unique address is assigned the QC words; (2) the QC
word contains 10 bits of quantitative information in place
of the 5 command and 5 command complement bits.
Therefore, when this unique QC address is recognized,
the CCD selects the flight control sensor group (FCSG)
SSD and shifts the 10 bits of quantitative information
into the FCSG magnitude register. Hence, the QC quan-
titative bits are loaded as they are decoded.
The command subsystem processed approximately 1330
commands during the mission. There were a few cases
where commands had to be repeated, but this was due to
having an improper up-link RF lock at the time. All
commands were executed properly.
I. Television
The television subsystem is designed to obtain photo-
graphs of the lunar surface, lunar sky, and portions of
the landed spacecraft. For the Surveyor 11 mission, the
subsystem consisted of a downward-looking approach
camera, a survey camera capable of panoramic viewing,
and a television auxiliary for processing commands and
telemetry data.
1. Approach Camera
The approach camera was designed to be turned on at
a nominal distance of 1000 km above the lunar surface
to provide overlapping photographs of the surface during
the terminal descent phase. Although an approach cam-
era was installed on the Surveyor 11 spacecraft, it was
planned not to operate the approach camera on the
mission because it was desired to minimize spacecraft
operational requirements during the complex and critical
descent phase.
2. Survey Camera
The survey television camera is shown in Fig. IV-59.
The camera provides images over a 360-deg panorama
and from + 40 deg above the plane normal to the camera
Z-axis to — 60 deg below this same plane. The camera Z-
axis is inclined 16 deg from the spacecraft Z-axis. Each
picture, or frame, is imaged through an optical system
onto a vidicon image sensor whose electron beam scans
a photoconductive surface, thus producing an electrical
output proportional to the conductivity changes resulting
from the varying receipt of photons from the subject.
The camera is designed to accommodate scene lumi-
nance levels from approximately 0.008 ft-lamberts to
2600 ft-lamberts, employing both electromechanical mode
changes and iris control. Camera operation is totally
dependent upon receipt of the proper commands from
earth. Commandable operation allows each frame to be
generated by shutter sequencing preceded by appropriate
lens settings and mirror azimuth and elevation position-
ing to obtain adjacent views of the subject. Functionally,
the camera provides a resolution capability of approxi-
mately 1mm at 4 meters and can focus from 1.23 meters
to infinity.
Figure IV-60 depicts a functional block diagram of
the survey camera and television auxiliary. Commands
for the camera are processed by the telecommunications
command decoder, with further processing by the tele-
vision auxiliary decoder. Identification signals, in analog
form, from the camera are commutated by the television
auxiliary, with analog-to-digital conversion being per-
formed within the signal processing equipment of the
telecommunications subsystem. The ID data in PCM form
is mixed in proper time relationship with the video signal
in the TV auxiliary and subsequently sent to the tele-
communications system for transmission to earth.
The survey camera contains a mirror, filters, lens,
shutter, vidicon, and the attendant electronic circuitry.
The mirror assembly is comprised of a 10.5 X 15 cm
elliptical mirror supported at its minor axis by trunnions.
This mirror is formed by vacuum-depositing a Kanogen
surface on a beryllium blank, followed by a deposition
of aluminum with an overcoat of silicon monoxide. The
mirrored surface is flat over the entire surface to less
than V\ wavelength at A = 550 m/^ and exhibits an average
JPt TECHNICAL REPORT 32-1086
89
HOOD
MIRROR AZIMUTH
DRIVE MOTOR
VARIABLE
FOCAL LENGTH
LENS ASSEMBLY
FOCUS
POTENTIOMETER
IRIS
POTENTIOMETER
SHUTTER
ASSEMBLY
ELECTRONIC
CONVERSION
UNIT
MIRROR
MIRROR
ELEVATION
DRIVE ASSEMBLY
FILTER WHEEL
ASSEMBLY
VIDICON TUBE
VIDICON
RADIATOR
Fig. IV-59. Survey TV camera
90
JPL TECHNICAL REPORT 32-1086
SURVEY
CAMERA
TELECOMMUNI-
CATIONS
1
^ ^.,„...^. >
SIGNAL
PROCESSOR
J
FRAME ID EN/
V
\BLE GATE
ENABLE
GATE
VIDEO A AND
3 OUTPUT
SUMMING
AMPLIFIERS
^COMPOSITE VIDEO
FRAME ID
AND FRAME ID J
16 CHANNEL
COMMUTATOR
FOR FRAME ID
ELECTRONIC
CONVERSION
UNIT
COMMANDS
SUBSYSTEM
DECODER
ELECTRONIC
HEATER BLANKET
SWITCH
1
I
TV AUXILIARY
TELEVISION
COMMAND
SIGNALS
TELECOMMUNI-
CATIONS
GROUND COMMANDS
COMMAND
DECODER
>T0 TRANSMITTERS
FROM RECEIVERS
Pig. IV-60. Simplified survey TV camera functional block diagram
specular reflectivity in excess of 86%. The mirror is
positioned by means of two drive mechanisms, one for
azimuth and the other for elevation.
The mirror assembly contains three filters (red, green,
and blue), in addition to a fourth section containing a
clear element for nonmonochromatic observations. The
filter characteristics are tailored such that the camera
responses, including the spectral response of the image
sensor, the lens, and the mirror match as nearly as
possible the standard CIE tristimulus value curves (Fig.
IV-61). Color photographs of any given lunar scene can
be reproduced on earth by combining three video photo-
graphs, each made with a different monochromatic filter
element in the field of view.
The optical formation of the image is performed by
means of a variable-focal-length lens assembly placed
between the vidicon image sensor and the mirror assem-
bly. Each lens is capable of providing a focal length of
either 100 or 25 mm, which results in an optical field of
view of approximately 6.43 and 25.3 deg, respectively.
Additionally, the lens assembly may vary its focus by
means of a rotating focus cell from near 1.23 meters to
infinity, while an adjustable iris provides effective aper-
ture changes of from //4 to //22, in increments which
result in an aperture area change of 0.5. While the most
effective iris control is accomplished by means of com-
mand operation, a servo-type automatic iris is available
to control the aperture area in proportion to the average
scene luminance. As in the mirror assembly, potentiom-
eters are geared to the iris, focal length, and focus ele-
ments to allow ground determination of these functions.
A beam splitter integral to the lens assembly provides
the necessary light sample (10% of incident light) for
operation of the automatic iris.
Two modes of operation are afforded the camera by
means of a mechanical focal plane shutter located be-
tween the lens assembly and the vidicon image sensor.
JPL TECHNICAL REPORT 32-1086
91
UJ
_J
>
UJ
a:
0.4
400 440 480 520 560 600 640 680 720
WAVELENGTH, m^
Fig. IV-61 . Relative tristimulus values of the color
filter elements
Upon earth command, the shutter bhides are sequen-
tially driven by solenoids across an aperture in the
shutter base plate, thereby allowing light energy to reach
the image sensor. The time interval between the initia-
tion of each blade determines the exposure interval,
nominally 150 millisec. An additional shutter mode
allows the blades to be positioned to leave the aperture
20-30% GRAY
0.03l-in.-diam HOLE
15% GRAY
ORANGE-BLACK
9% GRAY
5% GRAY
5% GRAY
BLACK
BLUE-BLACK
Fig. IV-62. TV photometric/colormetric reference chart
open, thereby providing continuous light energy to the
image sensor. This mode of operation is useful in the
imaging of scenes exhibiting extremely low luminance
levels, including star patterns.
The transducing process of converting light energy
from the object space to an equivalent electrical signal
in the image plane is accomplished by the vidicon tube.
A reference mark is included in each corner of the
scanned format, which provides, in the video signal, an
electronic level of the scanned image. In the normal, or
600-line mode of operation, the camera provides one
600-TV-line frame every 3.6 sec. Each frame requires
nominally 1 sec to be read from the vidicon. A second
mode of operation provides one 200-line frame every
61.8 sec. Each frame requires 20 sec to complete the
video transmission and utilizes a bandwidth of 1.2 kc
in contrast to the 220 kc used for the 600-line mode. This
200-line mode is used for omnidirectional antenna trans-
mission from the spacecraft.
A third operational mode, used for stellar observations
and lunar surface observation under earthshine illumi-
nation conditions, is referred to as an integrate mode.
This mode may be applied, by earth command, to either
the 200- or 600-line scan mode. Scene luminances on the
order of 0.008 ft-lamberts are reproduced in this mode
of operation, thereby permitting photographs under
earthshine conditions.
Integral to the spacecraft and within the viewing ca-
pability of the camera are two photometric/colorimetric
reference charts (Fig. IV-62). These charts, one on
Omniantenna B and the other on a spacecraft leg adja-
cent to Footpad 3, are located such that the line of sight
of the camera when viewing the chart is normal to the
plane of the chart. Each chart is identical and contains
a series of 13 gray wedges arranged circumferentially
around the chart. In addition, three color wedges, whose
CIE chromaticity coordinates are known, are located
radially from the chart center. A series of radial lines is
incorporated to provide a gross estimate of camera reso-
lution. Finally, the chart contains a centerpost which
aids in determining the solar angles after lunar landing
by means of the shadow information. Each chart, prior
to launch, is calibrated goniophotometrically to allow an
estimation of po.st-landing camera dynamic range and to
aid photometric and colorimetric data reduction.
The survey camera incorporates a total of four heaters
to maintain proper thermal control and to provide a
thermal environment in which the camera components
92
JPL TECHNICAL REPORT 32-J086
Operate. The elements are designed to provide a sus-
taining operating temperature during the lunar night
<
o
m
<
o
>
UJ
>
t-
<
I.O
0.8
0.6
0.4
0.2
?
^
^a
J
n
:
¥\
f
f,A
f
o
n
3 OPE
N SHUTTER
X-
o»
?
«J
□
o
X
1
1 1
Mi-
Mi
M'.
» OPE
3 OPE
>2 OP
N SHUTTER
N SHUTTER
EN SHUTTER
1 i i. 1 „
10"
6 10° 2 4 6 10' 2
BRIGHTNESS/{T NUMBER)^
6 lOZ
Fig. IV-63. Camera 600-line light transfer characteristic
as a function of brightness (T No.)
if energized. These consume 36 w of power when initi-
ated. A temperature of — 20°F must be achieved prior
to camera turn-on.
3. Performance
A premission calibration was performed on the survey
camera with the camera mounted on the spacecraft. Each
calibration utilized the entire telecommunication system
of the spacecraft, thereby including those factors of the
modulator, transmitter, etc., which influence the overall
image transfer characteristics. The calibration data was
FM-recorded on magnetic tape for playback through the
ground support equipment (GSE) at Goldstone and Pasa-
dena. Thus the final calibration data recorded on the
real-time mission film and tape represents a complete
system calibration.
The calibration results, at the point of initial FM
recording (i.e., not including the GSE), are shown in
Figs. IV-63 through IV-68. Figure IV-63 represents a
65.8
66.8
67.8
68.8
69.8
70.8
71.8
q:
O
I-
<
_l o
=> -*:
Q -
^.9=
h- U.
a.
2 4 6 10^ 2 4 6
LUNAR BRIGHTNESS, ft-lambert
72.8 BACK
4 6 PORCH
Fig. IV-64. Camera 200-line light transfer characteristic as a function of lunar brightness
JPL TECHNICAL REPORT 32-1086
93
u
<
o
V)
<
llJ
>
UJ
2 4 6 10' 2 4 6
LUNAR BRIGHTNESS, ft-lambert
4.79 BACK
PORCH
Fig. IV-65. Camera 600-line light transfer characteristic as a function of lunar brightness
2.29
10"
OH
2.79
O
H- ,.
^^
3 ^
o in
O CM
3.29
2«
O "
1- Q.
>- 1-
O
Z o
UJ 2
3.79
3 >
a«
ffo:
o
h- u.
3 -"-^
4.29
a.
z
4.79
BACK
4 6
4 6 PORCH
Fig. IV-66. Camera 600-line transfer characteristic as a function of color filter position for the f/4 iris stop
lO' 2 4 6 10* 2 4 6 10°
LUNAR BRIGHTNESS , ft-lambert
94
JPL TECHNICAL REPORT 32-1086
,
1 i r-SYNC IIP
'^
I 1 ! LEVEL
<
8
~' 1 1 ' BACK PORCH
LU
•5? 0-2
1 HORIZONTAL SHADING IN I '-^^^'-
ft
^ 0.4
1 EACH SCAN LINE j^ 1
1 ^..^ttttTTT"
^^ PCM DATA
'H
^0.6
1 yfltrfflU
1 INTERVAL
h-
1/ .ii*^^^ «
u-
_10.8
1/ .W^ VERTICAL SHADING FOR
UJ
!^
7 [ pilJJiUi*-^ ENTIRE FRAME
LoJ
lliiy^
+
>
— ^ TIME
Fig. IV-67. Camera shading near saturation
composite of the 600-line mode light transfer character-
istic data for various f/stops (T number), thus illustrating
the data scatter. Figures IV-64 through IV-66 show the
individual curves that were obtained for various f/stops
and color filters for the 200- and 600-line scan modes.
The curves depict the sensitivity of the camera system
at the central portion of the frame to scenes of constant
or static light level. The camera system, however, did
not respond the same over the entire frame. This non-
uniform response, called "shading," is depicted in
Fig. IV-67.
The response of the spacecraft camera system to sinu-
soidally varying brightness scenes is shown by Fig. IV-68.
Here, the sinusoidal nature of the test scene is given by
<
UJ
80 160 240 320 400 480 560 640
FREQUENCY (TV LINE/PICTURE HEIGHT)
Fig. IV-68. Camera sine-v\^ave response characteristic
the abscissa in terms of frequency, and the relative
attenuation of the sine wave amplitude by the camera
system is given at the ordinate. The "peaking" of the
response curve of the 600-line modes is due to high
peaking electronics and would be compensated by the
GSE frequency response characteristic. From this curve
it is seen that the spacecraft system has a 22% horizontal
response at about 600 TV lines in the 600-line scan mode.
JPL TECHNICAL REPORT 32-1086
95
V. Tracking and Data Acquisition System
The Tracking and Data Acquisition (T&DA) System
for the Surveyor Project consists of facilities of the Air
Force Eastern Test Range (AFETR), Goddard Space
FHght Center (GSFC), and the Deep Space Network
(DSN). This section summarizes the mission preparation,
flight support, and performance evaluation of each fa-
cility within the T&DA System.
The T&DA System support for the Surveyor II mission
was considered excellent: some minor problems were
experienced during operations but had no effect on re-
quired performance coverage. All requirements were
met and in most cases exceeded. When this mission
became nonstandard following the attempted midcourse
correction, the DSIF performed very well, providing
unanticipated support under difficult conditions.
A. Air Force Eastern Test Range
The AFETR performs T&DA supporting functions for
Suroeyor missions during the countdown and launch phase
of the flight.
The Surveyor Mission requirements for launch phase
tracking and telemetry coverage are classified as follows
in accordance with their relative importance to successful
mission accomplishment:
Class I requirements consist of the minimum essen-
tial needs to ensure accomplishment of first-priority
flight test objectives. These are mandatory require-
ments which, if not met, may result in a decision
not to launch.
Class II requirements define the needs to accomplish
all stated flight test objectives.
Class III requirements define the ultimate in desired
support, and would enable the range user to
achieve the flight test objectives earlier in the test
program.
The AFETR configuration for the Surveyor II Mission
is listed in Table V-1. The configuration is similar to the
Surveyor I configuration except for the deletion of
the Range Instrumentation Ship (RIS) General Arnold.
JPL TECHNICAL REPORT 32-1086
97
Table V-1. AFETR configuration
station
Radar
VHP telemetry
S-band
telemetry
Merrilt Island
X
Cape Kennedy
X
X
X
Patrick AFB
X
Grand Bahama Island
X
X
X
Grand Turk
X
Antigua
X
X
X
Coastal Crusader (RIS 1)
X
X
Sword Knot (RIS 2)
X
X
Ascension
X
X
X
Pretoria
X
X
X
Figure V-1 illustrates the disposition of the range instru-
mentation ships and planned coverage for Surveyor U
launch day. Except in the case of S-band telemetry
facilities, AFETR preparations for Surveyor U consisted
of routine testing of individual facilities, followed by
several Operational Readiness Tests. All requirements
were met by AFETR for the Surveyor U mission.
1. Tracking (Metric) Data
The AFETR tracks the C-band beacon of the Centaur
stage to provide metric data. This data is required dur-
ing intervals of time before and after separation of the
spacecraft for use in calculating the Centaur orbit, which
can be used as a close approximation of the post-
separation spacecraft orbit. The Centaur orbit calcula-
tions were used to provide DSN acquisition information
(in-flight predicts).
The significance of the lack of a metric RIS to augment
the downrange land-based radars was recognized by
both the Surveyor Project and AFETR. The resulting
restriction in the launch window was such that a scrub
would have resulted if any additional hold had been
required over that actually experienced. The fact that
the Project elected to launch without metric RIS support
is not to be considered a precedent for future launches.
Estimated and actual radar coverages are shown in
Figs. V-2 and V-3. The combined coverage of all stations
is represented by the top set of bars in each figure. The
Class I requirements were met and exceeded with AFETR
stations downrange to Antigua (including Trinidad radar)
providing continuous coverage to L -I- 938 sec. The Trini-
dad radar operated in the expected skin track mode.
Since spacecraft separation occurred near the end of
Trinidad track, and in view of the small separation rates
between the spacecraft and Centaur, separation distance
was not great enough to be observed with radar in tht*
1-mc fine resolution tracking (FRT) mode. This mode
improves range resolution through pulse compression to
about 470 ft. Rough track was experienced by Grand
Turk as radar approached the loss-of-signal (LOS) point.
At this time AFETR had no indication of occurrence of
balance point shift. Later evaluation of data tended to
confirm that the balance point shift did not occur owing to
the roll attitude of the Centaur, which is not roll-attitude-
stabilized. Farther downrange, Ascension and Pretoria
experienced intermittent tracking conditions due to the
lobing of the C-band beacon caused by vehicle roll.
2. Atfos/Centour Telemetry (VHF)
To meet the Class I telemetry requirements, the AFETR
must continuously receive and record Atlas telemetry
(229.9-mc link) until shortly after Atlas/Centaur separa-
tion and Centaur telemetry (225.7-mc link) until shortly
after spacecraft separation. Thereafter, Centaur telemetry
is to be recorded, as station coverage permits, until
completion of the Centaur retro maneuver. In addition
to the land stations, the AFETR provided RIS Sword
Knot and RIS Coastal Crusader and one range telemetry
aircraft to cover the gap between Antigua and Ascension.
Estimated and actual VHF telemetry coverage is shown
in Fig. V-4. All Class I, II, and III requirements were
met since continuous and substantially redundant VHF
telemetry data was received beginning with the count-
down and through Pretoria LOS at L +3805 sec. Cover-
age was more than predicted. However, since the Centaur
stage is not designed for roll stabilization, the expected
coverage was based on a specified minimum db level at
the antenna null to allow for uncertainty in the antenna
gain.
3. Surveyor Telemetry (S-band)
The AFETR also is required to receive, record and
retransmit Surveyor S-band (2295-mc) telemetry in real-
time after the spacecraft transmitter high power is turned
on until 15 min after DSIF rise.
The S-band telemetry resources assigned to meet this
requirement were the two Range Instrumentation Ships,
the 85-ft antenna system at Grand Bahama Island, and
the 30-ft S-band (TAA-3A) antenna systems located at
Antigua and Ascension Islands. All primary S-band sys-
tems were used on a limited commitment basis since the
98
JPL TECHNICAL REPORT 32-1086
*.*
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JPt TECHNICAL REPORT 32-1086
99
«X>»«XXXX»$55«X 5«««»«»«5«4«
COMBINED
COVERAGE
CAPE
KENNEDY
PATRICK AFB
KENNEDY SPACE
CENTER
GRAND
BAHAMA ISLAND
GRAND
TURK
ANTIGUA
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800
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REQUIREMENT
ESTIMATED COVERAGE
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ACTUAL COVERAGE
INTERMITTENT TRACK
iiiiiiniiii
WM^Mi
MISSION TIME, sec
Fig. V-2. AFETR radar coverage: liftoff through Antigua
COMBINED
COVERAGE
ANTIGUA
TRINIDAD
ASCENSION
PRETORIA
i
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ANY CONTINUOUS GO SECOND INTERVAL
BEST OBTAINABLE
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POSSIBLE ROUGH TRACK !■■!
ACTUAL COVERAGE ^^
INTERMITTENT TRACK »m
1000
1500 2000
MISSION TIME, sec
2500
3500
Fig. V-3. AFETR radar coverage: Antigua through Pretoria
100
JPt TECHN/CAL REPORT 32-1086
■^s^^xxxxxxxxxxxxxxxxxxxxx^^
COMBINED
COVERAGE
CAPE KENNEDY
GRAND BAHAMA ISLAND
ANTIGUA
AIRCRAFT
COASTAL CRUSADER
llllllllllllllllllllillll
SWORD KNOT
ASCENSION
PRETORIA
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illlllllllilllllllllliillil
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500
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MISSION TIME, sec
2000
2500
Fig. V-4. AFETR VHP telemetry coverage
JPL TECHNICAL REPORT 32-1086
101
Centaur vehicle is not roll-attitude-stabilized. Confidence
in these systems was fairly high because of the successful
S-band coverage provided by RIS Sword Knot for the two
previous Centaur launches and the S-band support pro-
vided by RIS Coastal Crusader for Pioneer and Surveyor I.
A three-ft S-band antenna, with its associated down-
converter, receiver, and communications equipment, was
in place at Ascension. This system provided S-band data
for the two previous Centaur launches and was used to
back up the TAA-3A antenna system. A similar backup
system was provided at Antigua.
With the exception of a 10-sec gap at L -i-380 sec,-
AFETR land stations and ships obtained continuous
S-band telemetry coverage from liftoff through Ascension
LOS at L +2675 sec. Although Ascension experienced ^i-
dropout between L + 1742 and L + 1790 sec, the interval
was adequately covered by Pretoria. Estimated and actual
S-band telemetry coverages are shown in Fig. V-5. All
Class I, II, and III requirements were met.
4. Surveyor Real-Time Data
The AFETR retransmits Surveyor data (VHF or S-band)
to Building AO, Cape Kennedy, for display and for re-
COMBINED
COVERAGE
lllllllllllllllllllllllllllllllllllllllllllllllll
CAPE KENNEDY
GRAND BAHAMA ISLAND
ANTIGUA
COASTAL CRUSADER
SWORD KNOT
ASCENSION
PRETORIA
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500
REQUIREMENT ?$$$$$$5
ESTIMATED COVERAGE llllllllllll
ACTUAL COVERAGE ^^^M
1000 1500
MISSION TIME, sec
2000
2500
Fig. V-5. AFETR S-band telemetry coverage
102
JPL TECHNICAL REPORT 32-1086
■transmission to the SFOF. In addition, downrange sta-
tions monitor specific channels and report events via
voice communication.
For the Surveyor U mission, existing hardware and
software facihties were utihzed to meet the real-time data
requirements.
All requirements were met. VHF telemetry data, in-
cluding spacecraft data, were transmitted in real-time to
the user from liftoff to spacecraft high power on. At high
power on, AFETR switched as planned to real-time
transmission of spacecraft S-band telemetry data to
Building AO. Real-time data flow was very good. In
addition, all Mark Events were read out and reported.
5. Real-Time Computer System (RTCS)
For the launch phase of the mission, the RTCS pro-
vides trajectory computations based on tracking data and
vehicle guidance data. The RTCS output includes:
(1) The interrange vector (IRV), the standard orbital
parameter message (SOPM), and orbital elements.
(2) Predicts, look angles, and frequencies for acquisition
use by downrange stations.
(3) I-matrix and moon map for mapping injection con-
ditions and estimating trajectory accuracy. Provides
for early orbit evaluation prior to orbital data gen-
erated by Flight Path Analysis and Command
(FPAC).
The RTCS computed a total of six orbits (five transfer
orbits and one Centaur post-retro orbit) using different
data sources. The first orbit, based on Antigua data, and
the related acquisition look angles and frequencies for
downrange stations were generated and transmitted on
schedule. The RTCS computed the second orbit based
on real-time Centaur guidance data, but some delay
occurred because of errors in the initial guidance data
provided to the RTCS by Kennedy Space Center (KSC).
The third orbit was a recursive solution using Antigua
data which improved the first orbit. The fourth orbit
involved a multistation solution using Antigua C-band
data and Trinidad skin- track data. The third and fourth
orbits were computed on the RTCS Computer B while
Computer A was attempting to process the Ascension
and Pretoria post-retro data to compute the fifth orbit.
However, the Ascension and Pretoria data was difficult
to process because many points were "off-track." (The
intermittent tracking by Ascension and Pretoria was
attributed to a weak C-band beacon.) After completing
the fifth orbit, the RTCS computed a final orbit based
upon data from Antigua and DSS 51.
B. Goddard Space Flight Center
The Manned Space Flight Network (MSFN), managed
by GSFC, supports Surveyor missions by performing the
following functions:
(1) Tracking of the Centaur beacon (C-band) for
approximately 3.5 hr.
(2) Receiving and recording Centour-link telemetry
from Bermuda acquisition until loss of signal at
Kano.
(3) Providing real-time confirmation of certain Mark
Events (see Appendix A).
(4) Providing real-time reformatting of Carnarvon
radar data from the hexidecimal system to the
38-character octal format and retransmitting these
data to the RTCF at AFETR.
(5) Providing NASCOM support to all NASA elements
for simulations and launch and extending this com-
munications support as necessary to interface with
the combined worldwide network.
The GSFC supported the Surveyor II mission with the
tracking facilities and equipment listed in Table V-2.
However, GSFC did not support the Operational Readi-
ness Test (ORT) prior to launch.
1. Acquisition Aids
Stations at Bermuda, Canary Island, and Kano are
equipped with acquisition aids to track the vehicle and
provide RF inputs to the telemetry receivers. Perform-
ance recorders are used to record AGC and angle errors
for post-mission analysis. The acquisition aids provide
telemetry RF inputs from Bermuda acquisition through
loss of signal at Kano. All MSFN acquisition aid systems
performed their required functions during the Surveyor U
mission.
2. Telemetry Data
Bermuda, Canary Island, and Kano were also equipped
to decommutate, receive, and record telemetry. Capa-
bility for coverage was provided from Bermuda acquisi-
tion through loss of signal at Kano. Mark Event readouts
JPl TECHNICAL REPORT 32-1086
103
Table V-2. GSFC Network configuration
Location
Acquisition
aid
VHF
Telemetry
C-band radar
SCAMA
Radar
high-speed data
Real-time
readouts
Bermuda
X
X
X
X
X
X
Canary Island
X
X
X
X
Kano
X
X
X
X
Carnarvon
X
GSFC
X
Cape Kennedy
X
X
were required from all stations in real-time or as near
real-time as possible when the vehicle was in view of a
station.
MSFN telemetry support was good. There were no
equipment failures or discrepancies during the operation.
Only Mark Events 6 and 7 were confirmed by Bermuda.
3. Tracking Data (C-Band)
Bermuda provided radar beacon tracking, magnetic
tape recording (at a minimum of 10 points/sec), and
real-time data transmission to GSFC and AFETR.
1. The DSIF
The following Deep Space Stations were committed
as prime stations for the support of the Surveyor U
mission:
DSS 11 Pioneer, Goldstone Deep Space Communi-
cations Complex (DSCC), Barstow, Cali-
fornia
DSS 42 Tidbinbilla, Australia, near Canberra (Fig.
V-6)
DSS 51 Johannesburg, South Africa
DSS 72 Ascension Island (first pass only)
4. Computer Support, Data Handling, and Ground
Communications
The GSFC Data Operations Branch provided computer
support during the prelaunch, launch, and orbital phases
of the mission. Data was provided by MSFN stations at
Bermuda, Canary Island, Kano, and Carnarvon in accord-
ance with the requirements of the Network Operations
Plan. Existing NASCOM and DOD Network facility
voice and teletype circuits provided ground communica-
tions to all participating stations.
C. Deep Space Network
The DSN supports Surveijor missions with the inte-
grated facilities of the Deep Space Instrumentation
Facility (DSIF), the Ground Communication System
(GCS), and the DSN facilities in the Space Flight Oper-
ations Facility (SFOF).
DSS 11, 42, and 51 are equipped with 85-ft-diameter,
polar-mount antennas. DSS 72 is equipped with a 30-ft
azimuth-elevation antenna. These prime stations were
committed to provide tracking coverage on a 24-hr/day
basis, from launch to lunar landing, and for the first
lunar day and night. For succeeding lunar days and
nights, the commitment was for 24-hr/day coverage dur-
ing the first three and last two earth days and for
lO-hr/earth day in between.
In addition to the basic support provided by prime
stations, the following facilities support was provided for
the Surveyor 11 mission:
(1) DSS 71, Cape Kennedy, provided facilities for
spacecraft/DSIF compatibility testing, and also
received telemetry after liftoff for engineering
evaluation of its new Command and Data Han-
dling Console (CDC) equipment.
104
JPt TECHN/CAl REPORT 32-1086
Fig. V-6. DSS 42, Tidbinbilla, Australia
(2) DSS 61, Robledo, Madrid DSCC, was designated
a training station during the Surveyor 11 mission
and was committed to provide tracking capability
within a 1- to 1.5-hr callup. During Pass 2, DSS 61
provided emergency telemetry and command cov-
erage when communications problems with DSS 51
were encountered.
(3) DSS 12, Echo, Goldstone DSCC, provided a backup
transmitter capability during midcourse maneuver.
This support, however, was not required.
(4) DSS 14, Mars, Goldstone DSCC, provided backup
telemetry coverage using the 210-ft antenna during
both Goldstone passes. The Mars station assisted
with accurate measurement of spacecraft tumbling
rates during Pass 2.
Data is handled by the prime DSIF stations as follows:
(1) Tracking data, consisting of antenna pointing angles
and doppler (radial velocity) data, is supplied in
near-real-time via teletype to the SFOF and post-
flight in the form of punched paper tape. Two-
and three-way doppler data is supplied full-time
during the lunar flight, and also during lunar oper-
ations when requested by the Surveyor Project
OfBce. The two-way doppler function implies a
transmit capability at the prime stations.
(2) Spacecraft telemetry data is received and recorded
on magnetic tape. Baseband telemetry data is sup-
plied to the CDC for decommutation and real-time
readout. The DSIF also performs precommunica-
tion processing of the decommutated data, using
an on-site data processing (OSDP) computer. The
data is then transmitted to the SFOF in near real-
time over high-speed data lines (HSDL).
(3) Video data is received and recorded on magnetic
tape. This data is sent to the CDC and, at DSS 11
only, to the TV Ground Data Handling System
JPL TECHNICAL REPORT 32-1086
105
(TV-GDHS, TV-11) for photographic recording. In
addition, video data from DSS 11 is sent in real-
time to the SFOF for magnetic and photographic
recording by the TV-GDHS (TV-1). Since a soft
landing could not be achieved on the Surveyor 11
mission, no video data was received. After lunar
landing on a standard mission, DSS 11 performs a
special function. Two receivers are used for dif-
ferent functions. One provides a signal to the CDC,
the other to the TV-GDHS. (Signals for the latter
system are the prime Surveyor Project requirement
during this phase of a mission.)
(4) Command transmission is another function pro- -
vided by the DSIF. Approximately 280 commands
are sent to the spacecraft during the nominal
sequence from launch to touchdown. Confirmation- •
of the commands sent is processed by the OSDP
computer and transmitted by teletype to the SFOF.
The characteristics for the S-band and L/S-band track-
ing systems are given in Table V-3. The L/S-band con-
version is located at DSS 51, the Johannesburg station, and
consists of a hybrid, interim arrangement of L-band
Table V-3. Characteristics for S-band and L/S-band tracking systems
Antenna, tracking
Type
Mount
Beamwidth ±3 db
Gain, receiving
Gain, transmitting
Feed
Polarization
Max. angle troclting rote"
Max. angular acceleration
Tracking accuracy (Iff)
Antenna, acquisition
Type
Gain, receiving
Gain, transmitting
Beamwidth ±3 db
Polarization
Receiver
Typical system temperature
With poromp
With maser
Loop noise bandwidth
threshold (2B, „)
Strong signal (2B/.ii)
Frequency (nominol)
Frequency channel
85-ft parabolic
Polar (HA-Dec)
— 0.4 deg
53.0 db, 4 1.0, -0.5
51.0 db, + 1.0, 0.5
Cossegrain
LH'' or RH circular
51 deg/min 0.85 deg/sec
5.0 deg/sec/sec
0.1 4 deg
2 X 2-ft horn
21.0db±1.0
20.0 db ±2.0
— 16 deg
RH circular
S-band
270±50''K
55 ±10''K
12, 48 or 152 cps
10, -10%
120, 255, or 550 cps
t 0, -10%
2295 mc
14a
Transmitter
Frequency (nominal)
Frequency channel
Power
Tuning range
Modulator
Phase input impedance
Input voltage
Frequency response (3 db)
Sensitivity at carrier output frequency
Peak deviation
Modulation deviation stability
Frequency, standard
Stability, short-term (Iff)
Stability, long-term (Iff)
Doppler accuracy at F, , (Iff)
Data transmission
2113 mc
14b
1 kw, mox
±100 kc
> 50 U
< 2.5 V peak
DC to 100 kc
1 .0 rod peok/v peak
2.5 rod peak
±5%
Rubidium
1X10"
5X10"
0.2 cps " 0.03 m/$ec
TTY and HSDL
" Both oxes.
''Goldstone only.
106
JPL TECHNICAL REPORT 32-1086
and S-band components configured to operate at the
S-band frequency. Differences are in the hardware and
operational techniques rather than in performance char-
acteristics. The essential differences are as follows: The
L/S-band receiver subsystem consists of standard L-band
receiver elements with modification equipment added to
permit acceptance of the nominal S-band receive fre-
quency of 2295 mc. Where the S-band stations are
equipped with two standard receivers, the L/S-band
conversion consists of only one receiver suitable for
tracking functions (two angles and radial velocity) and a
second, "suitcase" receiver which is used for telemetry
reception. Telemetry bandwidths and loop noise band-
widths are restrictive compared to the S-band system.
The angle-tracking parameters for stations equipped
with 8.5- ft antennas are as follows:
(1) Maximum angle tracking rate (both axes): 51 deg/
min = 0.85 deg/sec.
(2) Maximum angular acceleration: 5.0 deg/sec/sec.
(3) Tracking accuracy (one standard deviation):
a = 0.14 deg.
(4) The system doppler tracking accuracy at the re-
ceiver carrier frequency for one standard deviation
is 0.2 cps = 0.03 m/sec.
The maximum doppler tracking rate depends on the
loop noise bandwidth. For phase error of less than
30 deg and strong signal (-100 dbm), tracking rates are
as follows:
Loop noise
bandwidths, cps
12
48
152
Maximum
tracking rate, cps/sec
100
920
5000
The angle tracking parameters for the DSS 72 30-ft
antenna are as follows:
(1) Maximum azimuth tracking rate: 6 deg/sec.
(2) Maximum elevation angle tracking: 3 deg/sec.
(3) Tracking accuracy: 0.01 deg.
(4) The system doppler tracking accuracy and doppler
tracking rates are the same as for 85-ft antennas.
The receiver characteristics for S-band and L/S-band
stations are as follows:
(1) Noise temperature. The total effective system noise
temperature including circuit losses when looking
at or near the galactic pole is:
Traveling- wave maser, 55 ±10°K
Parametric amplifier, 270 ±50°K
(2) Loop noise bandwidth. The closed-loop noise band-
width for various signal conditions is:
Threshold 26^,,, 12, 48, or 152 cps +0, -20%
Strong signals 2B„ 132, 274 or 518 cps +0, -20%
(3) Thre.shold. Carrier lock will be maintained with an
rms phase error due to noise of less than 30 deg
when the ratio of carrier power to noise power in
the closed-loop noise bandwidth is 6 db or greater.
Owing to the nature of the operation of the phase-
lock loop, this condition requires the carrier power
at the receiver input to be 9 db greater than the
value at threshold, which is defined as a carrier-
to-noise power ratio of zero db in the threshold
loop noise bandwidth 2Bui.
a. DSIF preparation testing. Operational Tests (C-
Tests) are conducted for each mission to verify that all
prime stations, communication lines, and the SFOF are
fully prepared to meet Mission responsibility. Selected
portions of the Sequence of Events are followed rigidly,
using both standard and nonstandard procedures.
The operational test schedule is presented in Table V-4.
DSS 11, 42, 51, and 72 participated in the Operational
Readiness Test (ORT) C-5.0 Phase 1, which was con-
ducted a week prior to launch. An evaluation of station
and Net Control support during the ORT indicated the
Table V-4. Operational test schedule
Test
Stations
Date, 1966
C-1.5
DSS 72
8/21
C-1.6
DSS 72 and AFETR
8/25
C-3.0 Phase 1
DSS 11, 42,51, 72 ond AFETR
8/30
C-3.0 Phase 2
DSS 11
9/1
C-5.0 Phase 1
DSS 11,42, 51, and 72
9/13
C-5.0 Phase 2
DSS 11
9/15
JPL TECHNICAL REPORT 32-1086
107
readiness of the T&DA System. Although the operational
tests were minimal, each station was adequately manned
and trained to properly support the Surveyor 11 mission.
Surveyor on-site computer program (SOCP) integration
tests are conducted to check out the SOCP and to verify
that data can be transmitted from a DSIF station to the
SFOF and then processed. Such tests were run on a
regular basis with each prime station (DSS 11, 42, and
51). These tests were concluded with a checkout of the
final SOCP program in April 1966. Operational tests con-
tinued up until three weeks prior to launch to provide
additional training for operational personnel.
b. DSIF flight support. All of the DSIF prime and
engineering practice stations reported "go" status during
the countdown. All measured station parameters were
within nominal performance specifications, and com-.-
munications circuits were up.
Figure V-7 is a profile of the DSIF mission activity
from launch until mission termination. This figure con-
tains the periods each station tracked the spacecraft
plotted against mission time. Table V-5 is a tabulation
of all commands sent during the Surveyor 11 Mission.
The DSIF stations operated very well, providing con-
tinuous tracking and telemetry coverage from L + 00:17
72
51
<
£
in
a.
u
Il/I2»
14
00.17
13:09
0023
0048
42
I0:I9
8|
ii
ill
^1^
17:28
22:37
34:37
27:17
32:16
34:24
34:37
42:18
34:33
42:15
I I
»DSS 12 ON ACTIVE STANDBY WITH TRANSMITTER POWER DURING OSS II FIRST PASS
I I I I I
38:04
45:24
16 20 24 28
TIME FROM LAUNCH, hr
Fig. V-7. Station tracking periods
32
108
JPt TECHN/CAl REPORT 32-1086
Table V-5. Commands transmitted by DSIF stations
Pass No.
Date,
1966 GMT
Event
Commands transmitted
DSS71
DSS72
DSS51
DSS 11
DSS 42
DSS 61
Lounch
9/20
launch
9/20
Launch/01
9/20
85
8
01
9/21
Midcourse correction commanded
102
01
9/21
86
02
9/21
102
02
9/22
744
3
02
9/22
Retro ignition commanded
Station Total
Grand Total
449
187
854
535
3
1579
(DSS 72 rise) to end of mission at L +45:24. This was
the first use of DSS 72 as a committed Surveyor station,
and the two-way tracking data suppHed was quite useful
in the initial orbit calculations. This station was also used
to fill a gap in station coverage between first pass set at
DSS 51 and first pass rise at DSS 11. The gap had a
duration of 31 min and was a result of launching late in
the window on a more southerly azimuth. Because of the
lower antenna gain (30-ft antenna) and higher system
noise temperature (paramp instead of maser), it was nec-
essary to reduce the spacecraft data rate to 17.2 bit/sec
in order to obtain usable telemetry data. This was done,
and DSS 72 was able to obtain telemetry data during
the gap period.
DSS 42 experienced some difficulty obtaining lock dur-
ing its initial tracking period because of inaccurate
pointing angle data in the received predicts. This was
caused by an outdated set of antenna angle correction
coefficients for DSS 42 in the predict program.
The signal levels received at the DSIF stations are
shown in Fig. V-8. They compare favorably with the
predicted values. Only the portion of the mission before
the midcourse maneuver is shown because the post-
midcourse data is very inaccurate owing to spacecraft
tumbling. A 2-db discrepancy between DSS 11 signal
levels and other stations (notably DSS 42) is apparent.
This discrepancy is a result of the AGC calibration
method used and is consistent with results obtained in
the Surveyor I mission. The 7-db drop in signal level
recorded by DSS 51 between 16:00 and 18:00 is due to
a 3-db increase in carrier suppression of the 137.5 bit/sec
data rate and a 4-db spacecraft antenna loss due to
spacecraft orientation. Note that this occurred prior to
star acquisition. Because of its larger reflector (210 ft),
the received level of DSS 14 is 8 db above the received
signal levels of other prime stations. Since DSS 72 uses
a smaller reflector (30 ft), its received signal levels are
10 db less than the other stations. These levels were
consistent during the respective station tracking periods.
The variations in the DSS 72 signal level data are due to
the increased carrier suppression at 17.2 bit/sec coupled
with a loss of calibration accuracy at low signal levels.
When the attempted midcourse maneuver caused a
nonstandard spacecraft tumbling condition, low space-
craft signal strength coupled with S-band frequency
variation of 200 to 300 cps resulted in diflficult tracking.
In order to continue tracking the tumbling spacecraft,
DSS 51 and 42 modified their DSIF standard (Goldstone
duplicate standard: GSDS) receivers in real-time during
the second pass at each station. These modifications
extended the tracking loop bandwidths of DSS 51 and 42
by factors of 2 and 4, respectively.
During the second pass at DSS 51, when severe com-
munications problems existed, DSS 61 (designated a
training station) was called up. DSS 61 was able to
interrupt a Pioneer track, reconfigure for Surveyor, and
was in two-way lock ready for commanding in approxi-
mately 13 min. This operation was accomplished with a
very limited crew in the middle of the night. No high-
speed data was obtained from DSS 61 during this pass
because of a manning problem complicated by an OSDP
hardware problem.
JPL TECHNICAL REPORT 32-1086
109
-70
-80
90 - A
E
Si
■o
>
-100
S2 -110
cr
LlJ
tr
cr
<
z
o
CI
-120
-130
-140
X
DSS 51
II
42
72
14
DOUBLE LINE
DENOTES
STATION IN
TWO -WAY
LOCK
D DSS 1 1
O DSS 14
O DSS 42
A DSS 5 1
A DSS 72
-150 -
-160
UPPER MARGIN
NOMINAL PREDICTED
SIGNAL LEVEL FOR
STATIONS WITH 85 -ft
ANTENNA (TOTAL
POWER)
LOWER MARGIN
I
LAUNCH
ATTEMPTED
MIDCOURSE
CORRECTION
1
_L
12:00
i8;oo
SEPTEMBER 20
oo;oo
TIME (GMT) , 1966
Fig. V-8. DSS received signal level
06;oo
SEPTEMBER 21
12.00
A number of minor equipment anomalies and pro-
cedural problems occurred but were readily corrected
by station personnel without affecting the mission.
Although the masers at all stations performed well dur-
ing this mission, the prelaunch experience caused some
concern about their reliability. The low angle rate and
high signal level performance of DSS 51 require investi-
gation and possibly some engineering action or replace-
ment with a standard S-band system.
2. GCS/NASCOM
For Surveyor missions, the Ground Communications
System (GCS) transmits tracking, telemetry, and com-
mand data from the DSIF to the SFOF and control and
command functions from the SFOF to the DSIF by
means of NASCOM facilities. The GCS also transmits
simulated tracking data to the DSIF and video data and
base-band telemetry from DSS 11, Goldstone DSCC, to
the SFOF. The links involved in the system are shown in
Fig. V-9.
a. Teletype (TTY) circuits. Teletype circuits (four
available to prime stations) are used for tracking data,
telemetry, commands, and administrative traffic. The
teletype circuits were exceptionally reliable, the weakest
circuits (DSS 51) showing approximately 97% reliability.
The most serious problem associated with teletype cir-
cuits developed at Goldstone, where the total number of
teletype lines available in support of Surveyor U, Lunar
110
JPL TECHNICAL REPORT 32-7086
Fig. V-9. DSN/GCS communications links
JPL TECHNICAL REPORT 32-1086
m
Orbiter, and Pioneer Missions was inadequate. The lack
of circuits to Goldstone was a constraint during the
August-September testing periods. However, during
launch and midcourse, the system operated faultlessly.
JPL communication engineers added one additional TTY
circuit to the Goldstone Microwave System, increasing
the total to nine prior to the Surveyor II Mission. This
number, however, is still insufficient during SFOF multi-
mission periods. NASCOM provided two additional tem-
porary TTY circuits from JPL to Goddard to assure
sufficient lines during the launch phase of the mission.
h. Voice circuits. The voice circuits are shared be-
tween the DSIF and the Surveyor Project for administra-
tive, control, and commanding functions. The NASCOM
voice circuits provided for the Surveyor II mission per-
formed perfectly except for failures which occurred in
the circuits to the South Africa station (DSS 51) and the
Ascension Island station (DSS 72).
The DSS 51 voice circuit was above average during
the prelaunch phase. Just after launch, however, this
circuit failed because of expected high-frequency propa-
gation conditions.
The "backdoor" circuit via Australia was also inter-
rupted during this same critical period. A commercial
telephone link was established with DSS 51 via the New
York Overseas Operator, and voice operation was trans-
ferred to this temporary circuit. The commercial circuit
also failed periodically, though at different time intervals
from the NASCOM circuits. This additional capability
was maintained for approximately two hours after launch
and proved quite helpful. On September 21, the same
high-frequency radio propagation condition occurred
during certain critical portions of the DSS 51 view period
following the midcourse phase. Activation of the Com-
mercial Overseas Operator circuit was again required
for a period of approximately two hours. NASCOM pro-
vided all possible support in attempting to restore DSS 51
circuits during the mission critical launch phase and
during emergency operations following midcourse.
Because of high-frequency radio propagation, the
DSS 72 voice circuit was continually up and down, as
were most of the DSS 72 circuits during both premission
tests and the mission.
The Surveyor Control Net within the SFOF was over-
loaded during prelaunch tests and mission operations,
causing much slower voice communication. The problem
developed because of insufficient and limited configura-
tion of the Operational Voice Communication System- -
(OVCS). The intercom portion of the OVCS has since
been reconfigured to allow point-to-point coordination
of prepass events and to overcome the communications.-
overload.
c. High-speed data line (HSDL). One HSDL is pro-
vided to each prime site for telemetry data transmission
to the SFOF in real-time. This part of the communica-
tions system performed well during the mission, provid-
ing better-quality high-speed data with less downtime
than during the Surveyor I mission. Crossed lines at
GSFC were responsible for a 10-min loss of data from
AFETR, and a mispatch in the JPL Communications
Center resulted in the loss of some data from Goldstone.
Both modem* types (NASCOM and Hallicrafter) were
required during the Surveyor 11 mission, since the Johan-
nesburg station is not equipped with NASCOM modems.
Modem lA, one of the two Hallicrafter modem re-
ceivers, was again found to be considerably less reliable
during testing than was Receiver IB (as was the case
during the Surveyor I mission) and was not used during
the Surveyor II mission.
Hallicrafter Modem Receiver IB was used exclusively
for DSS 41 high-speed data. Reliability was considered
high, and much superior to that of the previous mission.
The changeover from Hallicrafter to NASCOM modems
at DSS 51 is still highly desirable, but it is doubtful that
this can be accomplished prior to the next mission, as
change of modems will be required at RCA New York,
Tangier, and Pretoria, South Africa, as well as at DSS 51.
The use of NASCOM modems in the Surveyor II mis-
sion proved to be highly successful, with fewer line
outages and with higher quality data received from all
stations.
d. Wideband microwave system. The wideband micro-
wave link between DSS 11, Goldstone DSCC, and the
SFOF consists of two 6-Mc lines for video, and one 96-kc
duplex line. The microwave link between Goldstone and
the SFOF performed with nearly 100% reliability on the
Surveyor II Mission. The microwave circuits were in-
volved, however, in numerous line level and patching
discrepancies during the Augu.st-September testing pe-
riods. Most patching problems experienced during testing
involved patching of mission nonstandard simulated data
*A modem (modulator-domoclulator) is a device for converting a
digital signal to a signal which is compatible with telephone line
transmission ( e.g., a frequency-modulated tone ) .
112
JPL TECHNICAL REPORT 32-7086
- through station tracking and video systems and return to
SFOF. After JPL communications engineers and Western
Union personnel had succeeded in obtaining proper
•equalization on these circuits, most of the problems dis-
appeared during the mission.
3. DSN in SFOF and DSN/AFETR Interface
The DSN supports the Surveyor missions by providing
mission control facilities and performing special func-
tions within the SFOF. The DSN also provides an inter-
face with the AFETR for real-time transmission of
downrange spacecraft telemetry data from Building AO
at Cape Kennedy to the SFOF.
a. Data Processing System (DPS). The SFOF Data
Processing System performs the following functions for
Surveyor missions:
(1) Computation of acquisition predictions for DSIF
stations (antenna pointing angles and receiver and
transmitter frequencies).
(2) Orbit determinations.
(3) Midcourse maneuver computations and analysis.
(4) On-line telemetry processing.
(5) Command tape generation.
(6) Simulated data generation (telemetry and tracking
data).
The DPS general configuration for the Surveyor U
mission is shown in Fig. V-10 and consists of two PDP-7
computers* in the telemetry processing station (TPS),
two strings of IBM 7044/7094 computers in the Central
Computing Complex (CCC), and a sublet of the input/
output (I/O) system.
The DPS performed in a nominal manner, with only
minor hardware problems which did not detract from
•Manufactured by Digital Equipment Corp.
TELEMETRY
PROCESSING
STATION
CENTRAL COMPUTING COMPLEX
Fig. V-10. Genera! configuration of SFOF data processing system
in TECHNICAL REPORT 32- J 086
113
mission support. The two PDP-7 computers were used
extensively to process high-speed telemetry data for the
Surveyor II mission. This processing consisted of decom-
mutating and transferring the data to the 7044 computer
via the 7288 data channels, generating a digital tape for
non-real-time processing, and supplying digital-to-analog
converters with discrete data parameters to drive analog
recorders in both the Spacecraft Analysis Area and the
Space Science Analysis Area. Two PDP-7 problems oc-
curred when the telemetry modes were changed. From
all indications, the computer appeared to be functioning
properly. However, data was not being supplied to the
7288 data channels. The problem was quickly resolved
by switching to the second PDP-7 as the prime processor.
It has not been determined if this was a hardware or
software problem.
The IBM 7044/7094 computer string dual configuration
successfully processed all high-speed data received from
the TPS and all teletype data received from the com-
munications center, as well as all input/output requests
from the user areas. A dual Mode 2 string was utilized by
the Surveyor Project until 00:00 GMT on September 22.
The problems experienced in the Central Computing
Complex were quite easily remedied and had little or
no effect on the mission.
The Input/Output System provides the capability for
entering data control parameters into the 7044/7094
computers and also for displaying computed data in the
user areas via the various display devices. The Input/
Output System performed quite adequately with only a
few reported problems.
b. DSN Intracommunications System (DSN/ICS). The -
DSN/ICS provides the capability of receiving, switching,
and distributing all types of information required for
spaceflight operations to designated areas or users within . •
the SFOF. The system includes facilities for handling all
voice communications, closed circuit television, teletype,
high-speed data, and data received over the microwave
channels.
In general, the DSN/ICS performance was well within
the expected reliability parameters. There were some
problems but these were not of a critical nature. During
the ORT's and Surveyor II mission, the communications
status display in the SFOF proved to be inadequate to
support multimissions. Fewer voice line patching errors
occurred during this mission than during the previous
mission. Seventeen tie lines were available from the
communications center to the telemetry processing station
during the Surveyor II mission.
c. DSN/AFETR interface. The DSN/AFETR inter-
face provides real-time data transmission capability for
both VHF and S-band downrange telemetry. The nom-
inal switchover time is after the spacecraft S-band trans-
mitter is turned to high power. The interface with the
Surveyor Project is at the input to the Command and
Data Handling Console in Building AO. The output of
the CDC is then interfaced with the Ground Communi-
cations Sy.stem for transmission to the SFOF. It is also
possible to go directly from the range data output to the
GCS, bypassing the CDC.
This real-time telemetry transmission interface per-
formed quite well. Good data was received at the SFOF
whenever good data was received at Building AO.
114
JPL TECHNICAL REPORT 32-7086
VI. Mission Operations System
A. Functions and Organization
The basic functions of the Mission Operations System
(MOS) are the following:
(1) Continual assessment and evaluation of mission sta-
tus and performance, utilizing the tracking and
telemetry data received and processed.
(2) Determination and implementation of appropriate
command sequences required to maintain space-
craft control and to carry out desired spacecraft
operations during transit and on the lunar surface.
The Surveyor command system philosophy introduces
a major change in the concept of unmanned spacecraft
control: virtually all in-flight and lunar operations of the
spacecraft must be initiated from earth. In previous
space missions, spacecraft were directed by a minimum
of earth-based commands. Most in-flight functions of
those spacecraft were automatically controlled by an
on-board sequencer which stored preprogrammed in-
structions. These instructions were initiated by either an
on-board timer or by single direct commands from earth.
For example, during the Ranger VIII 67-hr mission, only
11 commands were sent to the spacecraft; whereas for a
standard Surveyor mission, approximately 280 commands
must be sent to the spacecraft during the transit phase,
out of a command vocabulary of 256 different direct
commands. For Surveyor I, 288 commands were sent
during transit and over 100,000 commands were sent
following touchdown.
Throughout the space-flight operations of each Surveyor
mission, the command link between earth and spacecraft
is in continuous use, transmitting either fill-in or real
commands every 0.5 sec. The Surveyor commands are
controlled from the SFOF and are transmitted to the
spacecraft by a DSIF station.
The equipment utilized to perform MOS functions falls
into two categories: mission-independent and mission-
dependent equipment. The former is composed chiefly of
the Surveyor T&DA system equipment and has been
described in Section V. It is referred to as mission-
independent because it is general-purpose equipment
which can be utilized by more than one NASA project
when used with the appropriate project computer pro-
grams. Selected parts of this equipment have been as-
signed to perform the functions necessary to the Surveyor
Project. The mission-dependent equipment (described in
Section VI-B, following) consists of special equipment
which has been installed at DSN facilities for specific
functions peculiar to the project.
JPL TECHNICAL REPORT 32-1086
115
The Surveyor Project Manager, in his capacity as
Mission Director, is in full charge of all mission opera-
tions. The Mission Director is aided by the Assistant
Mission Director and a staff of mission advisors. During
the mission, the MOS organization is as shown in Fig.
VI-1.
Mission operations are under the immediate, primary
control of the Space Flight Operations Director (SFOD)
and supporting Surveyor personnel. Other members of
the team are the T&DA personnel who perform services
for the Surveyor Project.
During space-flight and lunar surface operations, all
commands are issued by the SFOD or his specifically dele-
gated authority. Three groups of specialists provide •
technical support to the SFOD. These groups are special-
ists in the flight path, spacecraft performance, and scien-
tific experiments, respectively.
1. Flight Path Analysis and Command Group
The Flight Path Analysis and Command (FPAC) group
handles those space-flight functions that relate to the
location of the spacecraft. The FPAC Director maintains
control of the activities of the group and makes specific
recommendations for maneuvers to the SFOD in accord-
ance with the flight plan. In making these recommenda-
tions, the FPAC Director relies on five subgroups of
specialists within the FPAC Group.
TSDA
MANAGER
FOR
SURVEYOR''
MISSION
DIRECTOR
ASSISTANT
MISSION
DIRECTOR
MISSION
ADVISORS
DSN
PROJECT
ENGINEER'
SFOD
ASFOD
r
I
t
I
SUPPORT
PROJECT
ENGINEER*
COMMUNICATIONS
PROJECT
ENGINEER"
DATA
PROCESSING
PROJECT
ENGINEER
DATA
CHIEF*
DPS
OPERATORS'
TPS
OPERATORS*
FPAC
DIRECTOR
TPG
TDA
GROUP
0D6
SPAC
DIRECTOR
PA
GROUP
CP
GROUP
"MISSION-INDEPENDENT
MAG
CS
GROUP
SSAC
DIRECTOR
TELPAC
SCTV-GDHS
DSIF
SOC
CDC
CREW
n
DSIF
OPERATIONS
PLANNING
PROJECT
ENGINEER*^!
DSIF
STATION
MANAGER"
ON-SITE
COGNIZANT
PROGRAMMER
SCTV-GDHS
DSIF
EQUIPMENT
OPERATORS*
J
Fig. VI-1. Organization of MOS
116
JPl TECHN/CAt REPORT 32-J086
(1) The Trajectory Prediction Group (TPG) determines
the nominal conditions of spacecraft injection and
generates lunar encounter conditions based on
injection conditions as reported by AFETR and
computed from tracking data by the Orbit Deter-
mination Group. The actual trajectory determina-
tions are made by computer.
(2) The Tracking Data Analysis (TDA) group makes a
quantitative and descriptive evaluation of tracking
data received from the DSIF stations. The TDA
group provides 24-hr/day monitoring of incoming
tracking data. To perform these functions the TDA
group takes advantage of the Data Processing
System (DPS) and of computer programs generated
for their use. The TDA acts as direct liaison be-
tween the data users (the orbit determination
group) and the DSIF and provides predicts to the
DSIF.
(3) The Orbit Determination Group (ODG), during
mission operations, determines the actual orbit of
the spacecraft by processing the tracking data re-
ceived from the DSN tracking stations by way of
the TDA group. Also, statistics on various param-
eters are generated so that maneuver situations can
be evaluated. The ODG generates tracking predic-
tions for the DSIF stations and recomputes the
orbit of the spacecraft after maneuvers to deter-
mine the success of the maneuver.
(4) The Maneuver Analysis Group (MAG) is the sub-
group of FPAC responsible for describing possible
midcourse and terminal maneuvers for both stand-
ard and nonstandard missions in real-time during
the actual flight. In addition, once the decision has
been made as to what maneuver should be per-
formed, the MAG generates the proper spacecraft
commands to effect these maneuvers. These com-
mands are then relayed to the Spacecraft Perform-
ance Analysis and Command Group to be included
with other spacecraft commands. Once the com-
mand message has been generated, the MAG must
verify that the calculated commands are correct.
(5) The Computing Support Group acts in a service
capacity to the other FPAC subgroups, and is re-
sponsible for ensuring that all computer programs
used in space operations are fully checked out
before mission operations begin and that optimum
use is made of the Data Processing System facilities.
2. Spacecraft Performance Analysis and
Command Group
The Spacecraft Performance Analysis and Command
(SPAC) Group, operating under the SPAC Director, is
basically responsible for the operation of the spacecraft
itself. The SPAC Group is divided into three subgroups:
(1) The Performance Analysis (PA) group monitors
incoming engineering data telemetered from the
spacecraft, determines the status of the spacecraft,
and maintains spacecraft status displays throughout
the mission. The PA group also determines the
results of all commands sent to the spacecraft. In
the event of a failure aboard the spacecraft, as indi-
cated by telemetry data, the PA group analyzes
the cause and recommends appropriate nonstand-
ard procedures.
(2) The Command Preparation (CP) group is basically
responsible for preparing command sequences to
be sent to the spacecraft. In so doing they provide
inputs for computer programs used in generating
the sequences, verify that the commands for the
spacecraft have been correctly received at the DSS,
and then ascertain that the commands have been
correctly transmitted to the spacecraft. If non-
standard operations become necessary, the CP
group also generates the required command
sequences.
(3) The Engineering Computer Program Operations
(ECPO) group includes the operators for the DPS
input/output (I/O) console and related card punch,
card reader, page printers, and plotters in the
spacecraft performance analysis area (SPAA). The
ECPO group handles all computing functions for
the rest of the SPAC group, including the main-
tenance of an up-to-date list of parameters for each
program.
In order to take maximum advantage during the mission
of the knowledge and experience of the various personnel
who are not a part of the "hard-core" operations teams
(FPAC, SPAC, and SSAC) but have been engaged in
detail design, analysis, or testing of the spacecraft, a
Spacecraft Analysis Team (SCAT) has been established.
The SCAT group, located in a building adjacent to the
SFOF, has appropriate data displays showing the current
status of the mission. The SCAT is available upon request
for immediate consultation and detailed analysis in sup-
port of the SPAC.
JPL TECHNICAL REPORT 32-1086
117
3. Space Science Analysis and Command Group
The Space Science Analysis and Command (SSAC)
group performs those space-flight functions related to the
operation of the survey TV camera. SSAC is divided into
two operating sub-groups:
(1) The Television Performance Analysis and Com-
mand (TelPAC) group analyzes the performance of
the TV equipment and is responsible for generat-
ing standard and nonstandard command sequences
for the survey TV cameras.
(2) The Television Science Analysis and Command
(TSAC) group analyzes and interprets the TV pic-
tures for the purpose of ensuring that the mission
objectives are being met. The TSAC group is under
the direction of the Project Scientist and performs
the scientific analysis and evaluation of the TV
pictures.
The portion of the spacecraft TV Ground Data Han-
dling System (TV-GDHS) in the SFOF provides direct
support to the SSAC group in the form of processed
electrical video signals and finished photographic prints.
The TV-GDHS operates as a service organization within
the MOS .structure. Documentation, system checkout, and
quality control within the system are the responsibility
of the TV-GDHS Operations Manager. During operations
support the TV-GDHS Operations Manager reports to
the SSAC Director.
4. Data Processing Personnel
The use of the Data Processing System (DPS) by
Surveyor is under the direction of the Assistant Space
Flight Operations Director (ASFOD) for Computer Pro-
gramming. His job is to direct the use of the DPS from
the viewpoint of the MOS. He communicates directly with
the Data Chief, who is in direct charge of DPS personnel
and equipment. Included among these personnel are the
I/O console operators throughout the SFOF, as well as
the equipment operators in the DPS and Telemetry
Processing Station (TPS) areas.
Computer programs are the means of selecting and
combining the extensive data processing capabilities of
electronic computers. By means of electronic data pro-
cessing, the vast quantities of mission-produced data are
assembled, identified, categorized, processed and dis-
played in the various areas of the SFOF where the data
are used. Their most significant service to the MOS is
providing knowledge in real-time of the current state of "
the spacecraft throughout the entire mission. This service
is particularly important to engineers and scientists of the
technical support groups since, by use of the computer'
programs, they can select, organize, compare and process
current-status data urgently needed to form their time-
critical recommendations to the SFOD. (See Section V-C-3
for a description of the DPS.)
5. Other Personnel
The Communications Project Engineer (PE) controls
the operational communications personnel and equipment
within the SFOF, as well as the DSN/GCS lines to the
DSIF stations throughout the world.
The Support PE is responsible for ensuring the avail-
ability of all SFOF support functions, including air
conditioning and electric power; for monitoring the dis-
play of Surveyor information on the Mission Status Board
and throughout the facility; for directing the handling,
distribution, and storage of data being derived from the
mission; and for ensuring that only those personnel
necessary for mission operations are allowed to enter the
operational areas.
The DSIF Operations Planning PE is in overall control
of the DSIF Stations at Goldstone, Johannesburg, and
Tidbinbilla; his post of duty is in the SFOF in Pasadena.
At each station, there is a local DSIF station manager,
who is in charge of all aspects of his DSIF station and
its operation during a mission. The Surveyor personnel
located at each station report to the station manager.
B. Mission-Dependent Equipment
Mission-dependent equipment consists of special hard-
ware provided exclusively for the Surveyor Project to
support the Mission Operations System. Most of the
equipment in this category is contained in the Command
and Data Handling Consoles and Spacecraft Television
Ground Data Handling System, which are described
below.
1. Command and Data Handling Console
The Command and Data Handling Console (CDC)
comprises that mission-dependent equipment, located at
the participating Deep Space Stations, that is used to:
(1) Generate commands for control of the Surveyor
spacecraft by modulation of the DSS transmitter.
118
JPL TECHNiCAL REPORT 32-1086
(2) Process and display telemetered spacecraft data
and relay telemetry signals to the on-site data
processor (OSDP) for transmission to the SFOF.
(3) Process, display, and record television pictures
taken by the spacecraft.
The CDC consists of four major subsystems:
(1) The command subsystem generates FM digital
command signals from punched tape or manual
inputs for the DSS transmitter, and prints a perma-
nent record of the commands sent. The major units
of the command subsystem, which can accommo-
date 1024 different commands, are the command
generator, the command subcarrier oscillator, the
punched tape reader, and the command printer.
Outgoing commands are relayed to the SFOF and
logged on magnetic tape by the DSS.
(2) The FM demodulator subsystem accepts the FM
intermediate-frequency signal of the DSS receiver
and derives from it a baseband signal. The base-
band signal consists of either video data or a com-
posite of engineering subcarrier signals. Depending
upon the type of data constituting the baseband
signal, the CDC processes the data in either the
TV data subsystem or the telemetry data subsystem.
(3) The TV data subsystem receives video data from
the FM demodulator and processes it for real-
time display at the CDC and for 35-mm photo-
graphic recording. In addition, telemetered frame-
identification data is displayed and photorecorded.
A long-persistence-screen TV monitor is mounted
in the CDC. The operator, when requested, can
thus evaluate the picture and, upon the SFOD's
direction, initiate corrective commands during
lunar television surveys.
(4) The telemetry data subsystem of the CDC sepa-
rates the various data channels from the baseband
signal coming from either the FM demodulator or
the DSS receiver phase-detected output and dis-
plays the desired data to the operators. Discrimi-
nators are provided for each subcarrier channel
contained in the baseband signal. The output of
each discriminator, in the case of time-multiplexed
data, is sent to the pulse code modulation (PCM)
decommutator and then relayed to both the OSDP
computer for subsequent transmission to the SFOF
and to meters for evaluation of spacecraft per-
formance. In the case of Continuous data trans-
missions, the output of the discriminator is sent to
an oscillograph for recording and evaluation.
The CDC contains built-in test equipment to ensure
normal operation of its subsystems. A CDC tester, con-
sisting of a spacecraft transponder with the necessary
modulation and demodulation equipment, ensures day-
to-day compatibihty of the CDC and DSIF stations.
a. Network configuration. Table VI-I lists the CDC
mission-dependent equipment provided for support of
Surveyor II at the DSIF stations. CDC's were located at
DSS 11, 42, 51, 61, 71, and 72. Stations 11, 42, 51, and 72
were the prime Surveyor stations. However, Station 61,
at Madrid, was used for some command transmissions.
DSS 12, the Echo site at Goldstone, was configured and
checked out to provide command backup to DSS 11.
The Echo Station transmitter was set to the Surveyor
frequency and a patchable interface established via the
intersite microwave link for the command subcarrier
from the CDC at DSS 11. A return link was also estab-
lished from the DSS 12 receiver back to the DSS 11 CDC
for purposes of checking the command transmission.
DSS 14, the Mars site at Goldstone, was configured to
record both pre- and post-detection signals on magnetic
tape. The added capability of this station was used to
increase the probability of obtaining data during critical
mission phases. DSS 71, the Cape Kennedy spacecraft
monitoring station, was provided for a DSIF compati-
bility test with the Surveyor II spacecraft several weeks
prior to launch. DSS 71 was also used to obtain space-
craft telemetry data during the prelaunch countdown and
immediately after launch.
Table VI-1. CDC mission-dependent equipment
support of Surveyor II at DSIF stations
DSS 11
Goldstone
..Prime station with command, telemetry,
and TV
DSS 42,
Canberra.
..Prime station with command, telemetry,
and TV
DSS 51,
Johannesburg. . .
. . Prime station with command and telem-
etry
DSS 61,
Madrid . .
. . Backup station with command, telem-
etry, and TV
DSS 71,
Cape Kenr
edy. .
. . Station used tor spacecraft compati-
bility tests and pre- and post-launch
telemetry monitoring
DSS 72.
Ascension
. . Prime station with command and telem-
etry
DSS 12,
Goldstone
. . Station configured for command backup
DSS 14,
Goldstone
. . Station configured to record critical
mission phases
JPL TECHNICAL REPORT 32-1086
119
Table VI-2. Surveyor // command activity
station
Commands transmitted
DSSll
854
DSS42
535
DSS51
187
DSS61
3
DSS71
DSS72
1579
b. CDC operations. During the mission, CDC opera-
tions were conducted at six of the DSIF stations. Table
VI-2 hsts the number of commands transmitted by each
station during the mission. Only seven CDC anomalies
occurred during the Surveyor 11 Mission (countdown and
flight operations). No detrimental effects on the mission
resulted from these anomalies.
(1) DSS 11, Goldstone. The Pioneer Station, at Gold-
stone, participated in two passes. This station
used a full CDC with command, telemetry, and
TV equipment. Three additional interfaces identi-
cal to those used for Surveyor 1 were established,
as follows:
(a) During telemetry sequences, the received
telemetry subcarriers were transmitted to the
SFOF from the CDC via the "96-kc" line.
Signal-limiting and level adjustments were pro-
vided by the CDC.
(b) As mentioned earlier, an interface was estab-
lished with DSS 12, Echo Station, for command
backup. If necessary, the CDC command sub-
carrier oscillator could be patched to the inter-
site microwave link for transmission to the
Echo site, where the S-band transmitter would
be modulated. A detected signal from the
Station 12 transmitter was fed back to the
Pioneer Station CDC via another microwave
channel for checking command transmissions
in the CDC.
(c) Two dataphone links were established with
Hughes Aircraft Company (HAC), El Segundo,
California. One line carried the reconstructed
telemetry PCM waveform to HAC from the
CDC's decommutator; the second line carried
the command waveform obtained at the CDC
system tester.
(2) DSS 42, Canberra. This station participated in two" '
passes. The CDC configuration was standard with
full capability available, except that the spare TV
monitor was connected in parallel with the prime '
monitor for better on-site TV monitoring. DSS 42
commanded RADVS turn-on and retro firing.
(3) DSS 51, Johannesburg. This station participated in
two passes. The CDC configuration was standard
although the interface with the DSIF was modified
to use the L/S-band receiver during the second
pass. Because of the tumbling spacecraft, the
Surveyor suitcase receiver (S-band receiver for
Surveyor telemetry) was unable to lock onto the
spacecraft signal since the suitcase receiver track-
ing loop bandwidth of 12 cps was too narrow to
track the doppler rate caused by tumbling. The
L/S receiver, with a bandwidth of 152 cps (later
changed to about 300 cps), was able to maintain
lock most of the time. Thus, telemetry data recep-
tion was possible.
(4) DSS 61, Madrid. This station participated on a
backup basis during two separate periods of the
second view period. No station countdown was
performed prior to the tracking periods, but two-
way lock with the spacecraft was achieved and
several commands transmitted.
(5) DSS 71, Cape Kennedy. This station was equipped
with a CDC in the period between Surveyor 1 and
U missions and includes only command and telem-
etry equipment. A DSIF compatibility test with
Surveyor 11 spacecraft was conducted in mid-
August to establish RF, command, and telemetry
compatibility. During the final portion of the pre-
launch countdown and the first several minutes
after liftoff, spacecraft telemetry data was pro-
cessed by this station and sent to the SFOF via one
high-speed data line and one teletype line.
(6) DSS 72, Ascension. The CDC at DSS 72 is limited
to pulse code modulation telemetry operation and
manual command transmission. No television or
analog telemetry equipment is provided. This sta-
tion acquired the spacecraft signal on the first pass
and started processing telemetry data and trans-
ferring the data to SFOF via high-speed data and
teletype lines. It also monitored the spacecraft two-
way acquisition by DSS 51. The station again
acquired the spacecraft during the loss-of-view
period between DSS 51 and DSS 11. After trans-
ferring the spacecraft to DSS 11, DSS 72 termi-
nated participation in the Surveyor U mission.
120
iPl TECHNICAL REPORT 32-1086
2. Spacecraft Television Ground Data Handling System
The Spacecraft Television Ground Data Handling Sys-
tem (TV-GDHS) was designed to record on film the
television images received from Surveyor spacecraft. The
principal guiding criterion was photometric and photo-
grammetric accuracy with negligible loss of information.
The system was also designed to provide display infor-
mation for the conduct of mission operations and for the
production of user products such as prints, enlargements,
duplicate negatives, and catalogs of ID information.
The system is divided into two major parts which are
located at DSS 11, Goldstone, and at the SFOF, Pasadena.
At DSS 11 is an on-site data recovery (OSDR) subsystem,
and an on-site film recorder (OSFR) subsystem. These
subsystems are duplicated in the media conversion data
recovery (MCDR) subsystem and in the media conversion
film recorder (MCRF) subsystem which are located at
the SFOF along with additional equipment making up
the complete system.
The complete TV-GDHS was committed for the Sur-
veyor 11 mission. For the Surveyor I mission, only por-
tions of the system at the SFOF had been committed
because of implementation constraints.
a. Equipment at Pioneer Site, Goldstone (TV-ll).
Data for the TV-GDHS is injected into the system at the
interface between the Station 11 receiver and the OSDR.
At this point, the signal from Surveyor has been down-
converted to a 10-Mc FM-modulated signal. The OSDR
further down-converts it to 4 Mc* for the 600-line TV
mode (500 kc, and 70 kc from the CDC, for the 200-line
mode), inputs this signal into the station's FR-800 video-
tape recorder, and provides an output signal to the
Microwave Communication Link for transmission to the
SFOF. The station's FR-1400 records the baseband video
signal only during 600-line mode and the 500- and 70-kc
during 200-line mode. The OSDR further processes signals
to obtain television image synchronization, telemetry
synchronization, and the baseband video signal. This
information is then used by the OSFR to record the video
image and the raw ID telemetry in bit form on 70-mm
film, together with an internally generated electrical gray
scale and "human readable" time and record number.
Prior to the Surveyor II mission, replacement of the
cathode ray tube in the OSFR required an additional final
For the Surveyor I mission, the center frequency of the FM signal
was 5 Mc.
calibration of the film recorder. No new major problems
occurred during the mission. Since the mission was ter-
minated prior to spacecraft landing, no results were
obtained from TV-ll.
h. Equipment at the SFOF (TV-1). The signal pre-
sented to the Microwave Terminal at DSS 11 is trans-
mitted to the SFOF where it is distributed to the MCDR.
The MCDR processes the signal in the same manner as
the OSDR. An FR-700 video magnetic tape recorder
records the predetection signal in the same manner as
the FR-800 at DSS 11. In addition, the MCDR passes the
raw ID information to the media conversion (MC) com-
puter, which converts the data to engineering units. This
converted data is passed to (1) the film recorder, where
it is recorded as "human readable" ID, (2) the ID wall
display board in the SSAC area, (3) the disc file, where
the film chip index file is kept, and (4) the history tape.
The scan converter accepts the slow-scan image infor-
mation from the film recorder and converts it to the
standard RETMA television signal for use by the SFOF
closed-circuit television and the public TV broadcast
stations.
The MC film recorder provides two films. One of these
films is passed directly to the bimat processor; the other is
accumulated in a magazine and is wet-processed off-line.
The bimat processor laminates the exposed film with
the bimat imbibed material, producing a developed
negative and a positive transparency. The negative is
used to make strip contact prints, which are delivered to
the users. The negative is then cut into chips and entered
into the chip file, where they are available for use in
making additional contact prints and enlargements.
No major problems associated with the TV-GDHS
occurred during the Surveyor II mission. Prior to the
mission, a hardware malfunction in the disc file and its
interface to the MC computer prevented the Real-Time
System Program from working. Another problem, a non-
linear demodulated video signal in the MCDR, prevented
expeditious video verification and produced degraded
output recordings. The problems with these two sub-
systems were corrected, and the total system was opera-
tional prior to launch. However, since the mission was
terminated before spacecraft landing, there were no
results from TV-1 at the SFOF.
JPL TECHNICAL REPORT 32-1086
121
C. Mission Operations Chronology
Inasmuch as mission operations functions were carried
out on an essentially continuous basis throughout the
Surveyor U mission, only the more significant and special,
or nonstandard, operations are described in this chron-
icle. Refer also to the sequence of mission events pre-
sented in Table A-1 of Appendix A.
1. Countdown and Launch Phase
No significant problems were reported in the early
phases of the countdown, and the spacecraft operations
were ahead of schedule at times during this period, which
included the Spacecraft Readiness Test. During the count-
down, MSFN and AFETR encountered only temporary
difficulties with Rermuda and Trinidad radars and the
communication links with the RIS Coastal Crusader.
Approximately one hour prior to scheduled launch,
SPAC reported spacecraft Receiver R AGC indication
was 26 db below that of Receiver A. Launch Operations
personnel at Cape Kennedy indicated that this anomaly
was not due to spacecraft receiver failure as it had been
observed before owing to gantry movement, although
Surveyor I showed only a 5-db drop as a result of gantry
movement. On the basis of this report, the countdown
was continued. (Later, during the early cruise phase of
the mission, a Receiver R threshold test was conducted
which revealed that the anomaly was due to a faulty
telemetry indication.)
After proceeding normally to the built-in hold at T-5
min, additional unscheduled holds were required because
of launch vehicle problems associated with a low Centaur
hydrogen peroxide temperature indication, failure of the
Atlas liquid oxygen boil-off valve to close at the start of
flight prcssurization, and failure of the Atlas automatic
topping system to maintain satisfactory liquid oxygen
level while holding for the preceding problem. As a result
of these problems, which are described in greater detail
in Section H, liftoff was not achieved until the final
seconds of the available launch window.
Liftoff occurred at 12:31:59.824 GMT on September
20, 1966, with all systems reported in a "go" condition.
All "mark event" times were received from AFETR,
although the reports were somewhat late. Launch vehicle
performance appeared to be nominal, with no significant
anomalies on either the Atlas or Centaur. Injection of the
spacecraft into. the prescribed lunar transfer orbit was
well within established limits, and the required retro
maneuver was successfully perfonned by the Centaur.
A description of launch vehicle performance and sequence' *
of events from launch through injection is contained in
Section III.
Spacecraft performance during the launch-to-injection
phase appeared nominal. The Centowr-commanded
spacecraft events just prior to separation were monitored
as they occurred by observing spacecraft telemetry trans-
mitted to Cape Kennedy from downrange in real time.
With the exception of a 1-min (approx) transmission
dropout at L + 1 min due to a Cape-wide power failure,
a total of 44 min of in-flight spacecraft telemetry was
received at the SFOF from Cape Kennedy, including data
retransmitted from downrange stations in real time.
Following separation at 12:44:32 GMT, the spacecraft
executed the planned automatic sequences as follows.
Ry using its cold-gas jets, which were enabled at separa-
tion, the Flight Control Subsystem nulled out the small
rotational rates imparted by the separation springs, and
initiated a roll-yaw sequence to acquire the sun. After a
minus roll of approximately 72 deg and a plus yaw of
16.5 deg, acquisition and lock-on to the sun by the space-
craft sun sensors were completed at 12:48:13 GMT.
Concurrently with the sun acquisition sequence, the
A/SPP stepping sequence was initiated to deploy the solar
panel axis and roll axis 85 and 60 deg, respectively. At
12:54:46 GMT, the solar panel was in its proper transit
position. All of these operations were confirmed in real-
time from the spacecraft telemetry.
Following sun lock-on, the spacecraft coasted, with its
pitch and yaw axes controlled to track the sun and with
its roll axis held inertially fixed.
2. DSIF and Canopus Acquisition Phase
DSS 72 (Ascension) was the first DSIF tracking station
to "see" the spacecraft, and it achieved one-way lock with
the spacecraft at approximately L + 00:16:50. However,
by prior mission operations planning, initial two-way
acquisition was reserved for DSS 51 (Johannesburg).
Approximately 23 min after launch, the spacecraft be-
came visible to DSS 51, and the initial DSIF acquisition
procedure was initiated to establish the communication
and tracking link between the spacecraft and the ground
station. DSS 51 acquired one-way lock at L 4 00:25:00
and, less than 10 min later, confirmed that two-way lock
had been established with the spacecraft at L +00:32:58.
122
JPL TECHNICAL REPORT 32-1086
The first ground-controlled sequence ("Initial Spacecraft
Operations") was initiated at L +00:45 (13:17 GMT) and
consisted of commands for (1) turning off spacecraft
* equipment required only until DSIF acquisition, such as
high-power transmitter and accelerometer amplifiers,
(2) seating the solar panel and roll axis locking pins
securely, (3) increasing the telemetry sampling rate to
1100 bit/sec, and (4) performing the initial interrogation
of all telemetry commutator modes. All spacecraft re-
sponses to the commands were normal. As a result of
assessment of the data, it was determined from the star
intensity telemetry signal that an object (which was
believed to be the earth) was in the field of view of the
Canopus sensor. Therefore, it was recommended that the
roll axis be held in the inertial mode and the cruise mode
command (which would have caused the spacecraft roll
attitude to be slaved to the position of the earth) not be
sent to the spacecraft. It was also recommended that
Transponder A not be turned on, since the Receiver A
AFC indicated that this receiver was tracking the ground-
station signal.
The spacecraft continued to coast normally, with its
pitch-yaw attitude controlled to track the sun and with
its roll axis held inertially fixed. Tracking and telemetry
data was obtained by use of Transponder B and Trans-
mitter B operating in low power. The spacecraft telemetry
bit rate/mode profile for the complete mission is shown
in Fig. IV-5.
Approximately four and one half hours after launch,
control of the spacecraft was transferred to DSS 72 to
provide additional tracking data for FPAC. The additional
tracking data from DSS 72 was important in that it
provided confirmation of the DSS 51 data and greater
confidence in the premidcourse orbit determination.
Transfer to DSS 72 necessitated a decrease in telemetry
rate from 1100 bit/sec to 137.5 bit/sec owing to the lower
antenna gain available at DSS 72. At approximately
L + 05:13, control of the spacecraft was returned to
DSS 51, and shortly thereafter the telemetry rate was
increased to 1100 bit/sec.
At about L + 06:06, a spacecraft roll maneuver was
initiated for making a star map and then acquiring the
star Canopus in order to fix the roll attitude of the space-
craft to a precise position from which the midcourse
maneuvers could be initiated. At the recommendation of
SPAC, the maneuver was made with the spacecraft trans-
mitting Mode 5 data* at 1100 bit/sec by means of Trans-
*See Section IV-A for data content of telemetry modes.
mitter B high power via Omniantenna B with the
transponder off. Two complete revolutions (one using
Omniantenna B and one using Omniantenna A) were
used to generate the star map. The earth, moon, and
stars Shaula and Ras Alhague were identified on the map
in addition to Canopus, which appeared after 240 degrees
of roll. However, for reasons unknown at the time, the
relative intensities of the stars were not as had been
expected. (It was later determined that reflected earth-
light caused the abnormal intensities.) As was the case in
the Surveyor I mission, a Canopus lock-on signal was
not generated as the star sensor swept past Canopus,
because the Canopus intensity signal was above the upper
threshold of the lock-on range. As the vehicle continued
to roll, the time for sending the proper command to
achieve manned lock-on to Canopus was computed.
Manual lock-on was achieved successfully at approxi-
mately L +06:40. In this mode the spacecraft roll attitude
is controlled so that the Canopus sensor remains locked on
the star.
3. Premidcourse Coast Phase
About 15 min after Canopus lock-on, a gyro drift check
was initiated by commanding the spacecraft to inertial
mode. The vehicle continued to coast as before, but with
its attitude held inertially so that the sun and star sensors
continued to point at the sun and Canopus, respectively.
At L + 09:03, the gyro drift check was terminated by
commanding the return to Canopus lock-on.
DSS 51 lost visibility of the spacecraft at L + 09:46.
A gap of about 40 min would have occurred in spacecraft
visibility before spacecraft rise at DSS 11 owing to the
geometry of the trajectory which resulted from the high
value of launch azimuth (114.361 deg). In order to cover
this period, a deviation was made in the Standard Se-
quence of Events, permitting transfer of control to DSS 72
and reduction of telemetry rate again to 17.2 bit/sec.
After acquisition, DSS 72 had considerable difficulty in
providing good data (SPAC estimated that 80% of the
data was bad). After the spacecraft became visible to
DSS 11 at L + 10:12, transfer was made to this station
and the bit rate was commanded back to 1100 bit/sec.
The first of two premidcourse maneuver conferences
was convened about 7 hours after launch to present to the
project managers a preliminary set of maneuver alterna-
tives and a comprehensive spacecraft status report. The
maneuver alternatives are discussed in Section VII with
summary data presented in Table VII-2. All subsystems
JPL TECHNICAL REPORT 32-1086
123
of the spacecraft were reported as performing well with
two exceptions.
(1) The spacecraft Receiver B up-link sensitivity was
found to be approximately 18 db lower than
expected.
(2) During Canopus acquisition the star intensities did
not agree with predicted values, causing inter-
mittent lock-on of the star sensor and requiring, as
with Surveyor I, the transmission of the manual
Canopus lock-on command. Star positions did ap-
pear as predicted.
The SPAC director indicated that the spacecraft would
be capable of supporting any maneuver choice selected
from among the alternatives presented by FPAC.
The second premidcourse maneuver conference was
held about 12 hours after launch. The Mission Director
approved the following recommendations made by the
SFOD and Project Scientist relative to the midcourse
maneuver plan and target site selection, respectively;
(1) The maneuver to be a roll-yaw maneuver. Such a
maneuver was favored over the possible pitch-yaw
or yaw-pitch maneuvers because of reduced exe-
cution errors, and over the possible roll-pitch
maneuver because of a somewhat better antenna
profile and improved gain margins.
(2) Spacecraft transmission during the maneuver to be
over Omniantenna B exclusively.
(3) Spacecraft transmission during the maneuvers to
be in the one-way mode. (This recommendation
was based on the as yet unresolved problem of
reduced spacecraft Receiver B sensitivity. It was
subsequently determined by means of an up-link
sensitivity threshold test, described below, that the
malfunction was caused by a telemetry anomaly,
and the maneuvers were actually executed in the
transponder mode.)
(4) Spacecraft transmission rate to be 4400 bit/sec
throughout the maneuver sequence.
(5) The maneuver execution time to be at approxi-
mately launch plus 16% hr (05:00 GMT, September
21), leaving open the possibility of a launch-plus-
40-hr maneuver if the necessity should arise.
(6) The target site to be shifted slightly to 0.55 deg
latitude, 359.17 deg longitude.
Final computation of midcourse parameters was con-
ducted following approval of the above plan and target
site selection. (Refer to Section VII for a discussion of " '
the factors considered in selecting the midcourse cor-
rection magnitude and final aiming point.)
Because analysis of Receiver B AGC telemetry data
obtained during star verification and acquisition indicated
a signal strength which was approximately 18 db below
the predicted value, a special test for performing an in-
flight calibration of this data channel was recommended
to determine whether the receiver had a malfunction or
the telemetry calibration had shifted. This test was
required to establish the feasibility of utilizing the trans-
ponder for two-way tracking during the midcourse cor-
rection, inasmuch as a degradation of 16 db in Receiver B
sensitivity would be indicative of a receiver malfunction
and would preclude such utilization. Following satis-
factory completion of the scheduled premidcourse low-
power engineering interrogation, the special calibration
test was conducted. During this sequence, the DSS 11
transmitter power was reduced in 2-db steps until the
command threshold level, as indicated by an indexing of
the receiver-decoder-select unit, was reached. This oc-
curred after a total reduction of 24 db at telemetry-
indicated signal strengths of - 133 dbm for Receiver B
and - 121 dbm for Receiver A. The conclusion reached
was that the calibration of Receiver B had changed from
the pre-mission data, and that the signal strength could
be lowered by 24 db without causing a receiver index,
and by 30 db without causing a loss of carrier signal in
Receiver B. Therefore, it was recommended that the
midcourse correction be done in two-way lock.
At approximately L +14:19, the scheduled premid-
course engineering interrogation was initiated. This
sequence was executed using low-power transmitter
operation since this mode permitted a data rate of 1100
bit/sec to be obtained. As part of this sequence, the gyro
speeds were measured and were found to be exactly
nominal at 50 cps.
4. Midcourse Maneuver Phase
The midcourse correction sequence was initiated at
L + 15:42 with an engineering interrogation which indi-
cated that the spacecraft was in satisfactory condition for
the midcourse operations. This was followed by com-
mands to turn on transmitter high power and increase
the telemetry sampling rate from 1100 to 4400 bit/sec.
Starting at L +16:12, the required roll-attitude ( + 75.4
deg) and yaw-attitude ( + 110.6 deg) maneuvers were
executed satisfactorily, thereby aligning the spacecraft
axes in the desired direction for applying the midcourse
thrust. Next the spacecraft was prepared for midcourse
124
JPL TECHNICAL REPORT 32-1086
thrusting by sending commands to (1) turn on strain
gages, (2) pressurize the vernier propulsion system, and
, (3) load the desired thrust time in the flight-control pro-
grammer magnitude register. Then, at L -|- 16:28:02, the
command was sent to thrust the vernier engines for apply-
ing the midcourse velocity correction. Following this
command, the strain gage of Vernier Engine 3 indicated
that this engine was not thrusting properly, and the gyro
error signals became saturated (pitch error negative, yaw
error positive, and roll error negative). Based upon the
previously commanded time increment, the vernier
engines shut off after a thrust duration of 9.8 sec. How-
ever, a check of the DSIF receiver AGC recording in the
SPAC area showed that the vehicle was rotating at a rate
of approximately 1.22 rev/sec, with a secondary motion
having a period of approximately 12 sec.
The spacecraft was in the inertial mode for the mid-
course firing but, about 4 min after midcourse firing, it
was commanded to the rate mode. The gas-jet system is
active in both modes and was operating to reduce the
spin rate. However, the rate mode was preferred under
the existing conditions because in that mode the gyros
are less likely to be damaged because of the tumbling
motion, and angular rate data can be obtained more
directly from telemetry.
Approximately 14 min after midcourse firing, when it
became evident that the gas jets could not stop the
spirming (approximately 60% of the gas had been used,
and the spin rate was still 0.97 rev/sec), the gas jets were
inhibited to conserve the remaining gas supply.
A 2-sec firing was recommended to attempt to clear the
Vernier Engine 3 problem and determine if the spacecraft
could be stabilized by firing the vernier engines. This
sequence, using the midcourse thrust level, was attempted
at L -1-18:56 and again at L -1-19:18 without success.
5. Post-Midcourse Phase
Since the spacecraft was rotating such that solar panel
output was zero, the only sources of electrical power for
spacecraft loads were the main and auxiliary batteries.
In an attempt to conserve energy, a sequence was ini-
tiated in which the flight control coast phase power was
cycled on and off periodically. Power was left on for
approximately 40 min, then off for approximately 90 min,
the cycle being based on the gyro and electronic tem-
peratures of the flight control system having reached
limits of -t-70°F and 0°F, respectively.
An interrogation of Modes 2 and 4 at hourly intervals
was also initiated. In addition, the auxiliary battery mode
was commanded on when the auxiliary battery tempera-
ture dropped to -F35°F to utilize the energy of this
battery and to keep it above its lower operational limit.
At L -f- 28:40, a special post-midcourse conference was
convened by the Mission Operations System Manager to
consider current spacecraft and trajectory status and to
formulate specific plans directed towards attainment of
the primary mission objectives. The FPAC director first
reported that the spacecraft unbraked target coordinates
were computed to be approximately 7 deg latitude,
353 deg longitude. It was also stated that good two-way
doppler tracking data was received during the abortive
maneuver attempt and throughout the period following
it. Ground receiver lock had not been lost during the
unbalanced thrusting of the spacecraft. The SPAC
Director then reported the following major spacecraft
status items:
(1) Present spacecraft tumbling rate was approximately
0.95 rev/sec with the flight control system in the
rate mode. All gyros were saturated.
(2) Cold gas jets were inhibited and nitrogen gas
remaining was 2.16 lb, about 50% of normal.
(3) The spacecraft was operating in the phase-lock
mode over Omniantenna/Receiver B and trans-
mitting in the low-power mode at a telemetry rate
of 137.5 bit/sec.
(4) The spacecraft was operating in the auxiliary
battery mode and flight control power was on. Esti-
mated battery life remaining was approximately
25 hr at L 4- 27:54.
(5) Spacecraft temperatures were within acceptable
limits with the following exceptions:
(a) Compartment B was out of the sun and was
showing a constant loss of heat through the
radiator.
(b) The lower part of the spacecraft was absorbing
considerable amounts of energy as manifested
by much higher than normal temperatures of
the AMR nozzle (-l-150°F vs -190°F) and
retro attach points (-|-140°F vs -120°F for
Attach Point 2).
(c) The RADVS temperature was abnormally low
and out of tolerance.
(d) All shock absorber temperatures were low
(No. 2 was at — 15°F, and Nos. 1 and 3 were
at -30°F).
JPL TECHNICAL REPORT 32-1086
125
After presentation, by HAC, of a detailed analysis of
the Vernier Engine 3 anomaly, the Mission Director
placed highest priority on attempts to restore the vernier
engine system to normal operation and assigned Mission
Operations system personnel with the task of detailed
design and execution of a vernier engine remedial com-
mand sequence.
The sequence which was prepared provided for pulse-
firing the engine five times (0.2-sec duration per firing,
5-min intervals between firings), followed by firing of
engines for a 2-sec period. This sequence was initiated at
approximately L + 31:12 with the result that there ap-
peared to be no firing of Engine 3.
At L -1- 35:15, a final post-midcourse conference was
convened by the Mission Operations System Manager to
present a general plan for the further conduct of the
mission. This plan had been formulated by the Mission
Operations System Manager and SFOD and was con-
curred in by the Mission Director. Primary elements of
the general plan were as follows:
(1) Because of continued spacecraft abnormal behav-
ior, DSS 42 and 51 were to assume the mission
control function during their respective visibility
periods. The level of mission control organization
staffing at the SFOF was to be reduced during
these periods commensurate with the shift in
responsibility.
(2) A vernier engine thrust sequence consisting of five
short thrusts, 0.2 sec duration and 5 min apart,
followed by a 2-sec thrust, was to be executed in
1-hr intervals until the spacecraft either recovered
or failed entirely.
(3) If, upon evaluation of telemetry data at the remote
DSIF stations, it was determined that all three
vernier engines had ignited as a result of command
sequence execution, the Mission Director, Mission
Operations System Manager, SFOD, and Techni-
cal Analysis Area Directors were to be notified
immediately to assume full control from the SFOF
as soon as possible.
(4) Under such circumstances, the spacecraft would be
stabilized as soon as possible and normal spacecraft
transit orientation would be accomplished. Space-
craft power would be conserved to the extent
feasible. Providing there was sufficient time for a
trajectory correction before the spacecraft reached
the moon, another maneuver would be prepared
and executed. Detailed maneuver requirements
would be established, depending on the circum-
stances existing at the time.
Using this plan, four additional attempts to achieve
Vernier Engine 3 thrusting were made, beginning at
L + 36:28, but all proved to be ineffective. Between the
second and third vernier engine firings, two attempts were
made to command the deployment of the planar array
upward from its launch position so that the solar panel
would be lowered to a position where some illumination
of the solar panel would occur. Solar panel illumination
was desirable for two reasons:
(1) To obtain energy for the spacecraft.
(2) To achieve illumination of one or more of the
secondary sun sensor cells, which are mounted on
the face of the solar panel, so that the actual orien-
tation of the spacecraft could be established. The
two attempts to move the planar array were un-
successful, apparently because of the opposing
force created by the spacecraft rotation.
Following these unsuccessful attempts, the Project
Management replaced the above plan with one designed
to achieve a higher thrust level with less rise time by
placing the flight control system in the post-retro-eject
condition. The objective was to be accomplished by
commanding retro sequence mode on and emergency
retro eject prior to turning on the flight control thrust
power. This would prevent the ejection of the main retro
engine while placing the flight control programmer in the
desired state. The sequence was completed at L + 41:ll
with the commanding of a vernier engine firing of about
2.5-sec duration. Engine ignition and shutoff were both
effected by ground command. Again the results were
negative. With each attempt to fire the engines, the
rotation rate of the spacecraft continued to increase so
that, at completion of the post-retro-eject thrusting, the
spin rate was approximately 1.54 rev/sec as determined
from the DSS AGC variation.
After the failure of these attempts to salvage the
mission, a final four-part plan was implemented. The
necessary sequences were prepared in order to:
(1) Attempt to step the solar panel in an effort to
illuminate its active face and the secondary sun
sensor cells.
(2) Dump the helium to obtain a curve of pressure
decay as a function of time in order to determine
whether a zero-shift had occurred in the helium
pressure telemetry signal.
126
iPl TECHNICAL REPORT 32- J 086
(3) Perform an evaluation of the capability of the main
battery to continue to supply power reliably under
the heavy terminal descent load conditions (flight
control thrust phase power on, high-power trans-
mitter on, RADVS on, etc.) when the remaining
battery energy is on the order of only 15 to 30
amp-hr.
(4) Fire the main retro motor in the normal terminal
descent mode.
At L + 42:22, a squib was "blown" by ground com-
mand to unlock the solar panel. The solar-panel position
telemetry signal showed a change of approximately
23 deg, indicating that the force on the panel created by
the spacecraft spin had caused the panel to move. Further
attempts to step the panel were unsuccessful.
Beginning at L + 43:13, a sequence was executed for
pulse-firing the engine 5 times (0.2 sec for each firing,
with 1 min between firings) followed by a 20-sec firing
in the post-retro-eject mode. Although the temperature of
Vernier Engine 3 rose approximately 24 °F during the
20-sec firing compared to about 100 °F for Engines 1 and
2, the strain gage on Engine 3 indicated no thrust.
The helium dumping sequence was initiated at
L + 44:41, and appeared to confirm that a zero-shift in
the helium pressure telemetry had occurred. The ob-
served zero-shift would account for the relatively large
decrease in pressure which was noted when the system
was pressurized prior to execution of midcourse thrusting.
Flight control thrust phase power and RADVS were
turned on at L +44:47 when the energy remaining in the
main battery was estimated to be 10 amp-hr. The bus
voltage dropped from 19.4 to 17.3 v when this load of
47 amps was placed on the battery. The RADVS was
then turned off at the direction of the acting SFOD.
At this time, the spacecraft spin rate was approximately
2.3 rev/sec. A profile of the spacecraft spin rate following
attempted midcourse correction is presented in Fig. IV-11.
The Emergency AMR signal was sent to the spacecraft
to initiate the retro engine firing sequence at L + 45: 02: 17
(09:34 GMT, September 22, 1966). Ignition of Vernier
Engines 1 and 2 as well as the main retro motor was
verified. Approximately 30 sec after the retro motor
ignited, contact with the spacecraft was lost, terminating
the Surveyor II Mission.
JPL TECHNICAL REPORT 32-1086
127
VII. Flight Path and Events
For Surveyor II, the landing site selected prior to
launch for targeting of the launch vehicle ascent trajec-
tory was near the center of the Apollo zone of interest at
0.0 deg latitude, 359.33 longitude (0.67 deg west longi-
tude). The following factors influenced the selection of
this site: predicted terrain smoothness, desire to land
within the Apollo zone, off-vertical approach angle of
near 25 deg, and good post-landing lighting. An unbraked
impact speed was selected so that the Goldstone arrival
visibility constraints would be satisfied for all launch days
in the launch nerinri
A. Launch Phase
The Surveyor II spacecraft was launched from AFETR
launch site 36A at Cape Kennedy, Florida, on Tuesday,
September 20, 1966. The launch was held until almost the
close of the launch window owing to difficulties experi-
enced with the Atlas (see Section II). Liftoff occurred at
12:31 : 59.824 GMT. At 2 sec after liftoff, the Atlas/Centaur
launch vehicle began a 13-sec programmed roll that
oriented the vehicle from a pad-aligned azimuth of
105 deg to a launch azimuth of 114.361 deg. At 15 sec, a
programmed pitch maneuver was initiated. All event
times for the launch phase were nominal or within the
3-<r tolerance. The launch phase sequence is discussed in
greater detail in Section III. Nominal and actual event
times for all phases of the mission are summarized in
Table A-1 of Appendix A.
B. Cruise Phase
Separation of Surveyor from the Centaur occurred at
12:44:32.4 GMT on September 20, 1966, at a geocentric
latitude and longitude of 12.9 and 309.8 deg, respectively.
The spacecraft was in the sunlight at separation and
never entered the earth's shadow during the transit
trajectory.
The Johannesburg station (DSS 51) reported good one-
way data at 12:55:17, only seconds after predicted rise
over the station horizon mask. Good two-way data was
reported by DSS 51 at 13:05:07. The DSIF stations pro-
vided continuous tracking data coverage from this initial
acquisition until loss of signal occurred at approximately
09:35 GMT on September 22, 1966. The station tracking
periods are presented in Fig. V-7.
The nominal earth-moon transfer trajectory and events
are shown in Fig. VII-1. A plot of the actual Surveyor II
trajectory projected on the earth's equatorial plane is
provided in Fig. VII-2. The earth track traced by Sur-
veyor 7/ appears in Fig. VII-3. Specific events such as
sun and Canopus acquisition and rise and set times for
the DSIF stations are also noted.
JPL TECHNICAL REPORT 32-1086
129
LAUNCH
INJECTION
INITIAL DSIF
ACQUISITION
TO SUN
POSITION OF MOON AT IMPACT,
TOUCHDOWN .
VERNIER DESCENT \
(FROM 35,000-ft ALTITUDE) \
TO SUN RETRO INITIATED \ \ .3,
(60 ml FROM MOONl
STAR ACQUISITION
(^+ 6hr)
SUN ACQUISITION
« I hr AFTER
ACQUISITION)
PRERETRO MANEUVERS
( 30 min BEFORE
TOUCHDOWN)
MIDCOURSE CORRECTION
FOLLOWED BY REACQUISITION
OF SUN AND STAR ^ ■▼
(/.+ I5hr) "
Fig. VII-1. Earth-moon trajectory and nominal events
N)vU'
POSITION OF
MOON AT LAUNCH
2 I -I -*
DISTANCE ALONG JT AXIS OF THE EQUATORIAL PLANE, 10* km
Fig. VII-2. Surveyor II trajectory in earth's equatorial plane
130
JPL TECHNICAL REPORT 32-1086
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JPt TECHNICAL REPORT 32-1086
131
APPROXIMATE FINAL
IMPACT POIMT .
10 I
PRELAUNCH
TARGET SITE
WO
■0 E
10 '
UNCORRECTED,
UNBRAKEO
IMPACT POINT
10
Fig. VII-4. Surveyor // target, uncorrected impact, and final impact points
132
JPL TECHNICAL REPORT 32-1086
The proximity of the uncorrected, unbraked impact
point (-0.0837 deg latitude, 354.658 deg longitude) and
the original aiming point (0.0 deg latitude, 359.33 deg
•longitude) is shown in Fig. VII-4. The uncorrected,
unbraked impact point is located on the western edge of
Sinus Medii just northeast of the crater Mosting. The two
points are approximately 142 km (88 mi) apart on the
surface of the moon. Also shown in Fig. VII-4 is the
approximate final impact site of the spacecraft.
C. Midcourse Maneuver Phase
The original aim point was selected assuming the 99%
landing dispersions to be a 50-km-radius circle on the
lunar surface. However, primarily because of the small
midcourse correction which was determined to be re-
quired, the 99% dispersion computed in-flight was found
to be an ellipse when mapped on the moon's surface with
a semimajor axis of 53.9 km and a semiminor axis of
17.17 km. Because of this smaller dispersion and based
upon a detailed examination of Lunar Orbiter I photo-
graphs, it was decided to bias the aiming point a little to
the north northwest to maximize the probability of soft
landing. The enlarged area of the moon shown in Fig.
VII-5 illustrates the initial aim point, the final aim point,
the 99% dispersion associated with premidcourse track-
ing and execution errors, and the approximate final
impact site. The latitude and longitude for these locations
as well as for the uncorrected impact point are given
below:
Longitude,
deg
Latitude
deg
Original aim point
0.0
359.33
Uncorrected impact point
Computed in-flight
-0.0837
354.658
Computed post-flight
-0.0519
354.710
Final aim point
0.55
359.17
Final impact point (approx)
4.0
349.0
Table VII-1 presents the injection and uncorrected
encounter conditions. These are the results of the final
post-flight calculations.
Table VII-1. Injection and uncorrected encounter conditions
Ceordinol*
system
Iniection
conditions September 20, 1966, 12:43:13.670 GMT
Inertia! cartesian
X =
-4360.9041 km
Y =
4616.8513 km
Z =
1896.4001 km
DX =
-8.7282187
km/sec
DY =
-4.5253066
km /sec
DZ =
-4.7559824
km/sec
lnAr»:»l *^u — ; 1
RAD =
6627.9056 km
DEC —
16.626021 deg
133.36699 deg
VI -
lu.yz laiy
km/sec
OTI
6.4654237 deg
11 9.42063 deg
Earth-fixed spherical
RAD =
LAT =
LON =
VE =
10 523257
PTE =
AZE =
6627.9056 km
16.626021 deg
303.60264 deg
km/sec
6.71 121 66 deg
120.66787 deg
C3 =
Orbital elements
-1.0001392
ECC =
INC =
TA =
LAN =
APF =
kmVsec'
0.98358240
33.423575 deg
13.039233 deg
340.26840 deg
135.66564 deg
Uncorrected encounter conditions September 23, 1 966, 03
:1 9:54.426 GMT
Selenocentric
RAD =
1738.5 km
LAT -
-.051930781 deg
LON =
354.70985 deg
VP =
2.6614642
km/sec
PTP =
-69.779146 deg
AZP =
90.222833 deg
Miss parameter earth
equator
BTQ =
1321.7977 km
BRQ =
-1793.8803 km
B= 1333.9151 km
Miss parameter moon
equator
BTT =
1333.9064 km
BRT =
5.1242575 km
B= 1333.9162 km
JPL TECHNICAL REPORT 32-1086
133
WEST
Fig. VII-5. Surveyor // impact locations
134
JPL TECHNICAL REPORT 32-J086
The maximum midcourse correction capability is
shown in Fig. VII-6 for three values of unbraked impact
speed, V,„,p. Also shown are the expected So- Centaur
injection guidance dispersions, the eflFective lunar radius,
and the maximum maneuver which could have been
executed if thrusting had been terminated automatically
by a spacecraft timer signal. The midcourse capability
contours are in the conventional R-S-T coordinate system.*
*Kizner, W. A., A Method of Describing Miss Distances for Lunar
and Interplanetary Trajectories, External Publication 674, Jet
Propulsion Laboratory, Pasadena, August 1, 1959.
A midcourse correction of 9.587 m/sec was selected for
execution by the spacecraft. The velocity component in
the critical plane, to correct "miss only," was 1.185 m/sec.
This component is referred to as the critical component.
The velocity component normal to the critical component
is referred to as the noncritical component since it does
not affect the miss to first order. Figure VII-7 presents
the variations in flight time, main retro burnout velocity,
and vernier propellant margin with the noncritical veloc-
ity component f/a. The propellant margin and flight time
were acceptable within the limits shown. However, it was
desirable to (1) provide backup midcourse correction
o
E
-2
kmx 10'
Fig. VII-6. Midcourse capability contours for September 20 launch
12 14
JPL TECHNICAL REPORT 32-1086
135
13
<
z
z
<
UJ
Q.
O
(t
Q.
UJ
>
50
40
30
20
10
-10
-20
>-
o
o
_i
>
3
o
z
(T
3
O
O
(T
I-
UJ
K
<
S
800
700
600
500
400
300
200
100
-100
-60 -40 -20 20
NONCRITICAL VELOCITY COMPONENT i/j, m/sec
Fig. VII-7. Effect of noncritical velocity component on terminal descent parameters
capability in the event the first midcourse correction was
unsatisfactory, (2) not exceed a main retro burnout
velocity of 450 ft/sec for flight control stability con-
siderations, and (3) keep the midcourse maneuver small
in order to reduce the execution errors, which are pro-
portional to the magnitude. Consideration of these fac-
tors led to the selection of a value of 9.5 m/sec for the
noncritical component. The predicted results of the
selected midcourse correction and alternatives considered
are given in Table VII-2.
If the maneuver strategy had been simply to correct
miss and flight time to the new aim point, the required
noncritical component would have been 4.325 m/sec,
giving a total correction of approximately 4.48 m/sec.
However, to properly evaluate the performance of the
Centaur guidance system, the original aiming point must
be used in computing the correction, in which case the
"miss only" correction is 1.015 m/sec and the "miss-plus-
flight-time" correction is 4.44 m/sec.
Execution of the midcourse correction was initiated at
05:00 GMT on September 21, 1966. Owing to failure of
Vernier Engine 3 to produce thrust during the midcourse
maneuver, a major, nonstandard flight condition devel-
oped wherein the spacecraft entered a tumbling mode.
D. Post-Midcourse and Mission Termination
The nonstandard condition which resulted from the
attempted midcourse correction precluded a normal soft
landing since the tumbling rate was too great to be over-
come with the attitude control gas-jet system, and re-
peated attempts to resolve the vernier system anomaly
by commanding short thrust bursts were unsuccessful.
136
JPL TECHNICAL REPORT 32-1086
Table VII-2. Midcourse maneuver alternatives
SeleclMl
midcoun*
Alternate condderctions
Execution time from launch, hr
16.5
No
midcourse
14.5
14.5
14.5
38.8
38.8
Critical component, m/$ec
1.18
0.51
0.51
0.51
1.0
2.24
Noncriticol component, m/sec
9.50
2.0
15.0
-33.4
1.7
17.0
Total correction, m/sec
9.59
2.1
15.0
33.4
2.0
17.2
Propellant required, lb
7.96
1.6
12.0
26.4
1.6
14.26
Unbraked impact velocity, km/sec
2.658
2.663
2.662
2.654
2.681
2.662
2.656
Main retro burnout velocity, ft/$ec
450
515
505
400
400
505
408
Vernier propellant margin, lb
30.5
31
31
31
33.4
31
30.0
Arrival time 9/23/66, GMT
03:42
03:20
03:25
03:57
03:42
03:22
03:38
Visibility
Time before landing
04:41
04:19
04:24
04:56
04:41
04:21
04:37
Time after landing
03:25
03:47
03:42
03:10
03:25
03:45
03:29
landing errors (3(r)
Semimajor axis, km
53.9
20
20
33.5
63
5.5
16.3
Semiminor axis, km
17.7
5
5
27
55
3.7
15.6
Orientation angle, deg
-57
-56
-56
-55
-60
-46
-52
The spacecraft batteries could not provide sufficient
power for the full duration of the transit phase since the
spacecraft was unable to obtain solar power in the un-
stable mode. Before power failure would have occurred,
a final command to ignite the retrorocket was transmitted
to the spacecraft at 09:34 GMT, September 22, 1966.
Mission termination resulted about 30 sec later with loss
of spacecraft signal. The best estimate of the impact loca-
tion of the Surveyor II spacecraft is 4 deg latitude,
349 deg longitude.
JPL TECHNICAL REPORT 32-7086
137
Appendix A
Surveyor II Flight Events
Table A-1. Mission flight events
Event
liftoff (2-in. rise)
Initiate roll program
Terminate roll, initiate pitch program
Moch 1
Max. aerodynamic loading
Booster engine cutoff (BECO)
Jettison booster package
Admit guidance steering
Jettison Centaur insulation panels
Jettison nose fairing
Start Centaur boost pumps
Sustainer engine cutoff (SECO)
Atlas/Centaur separation
Centour main engine ignition (MEIG)
Centaur main engine cutoff (MECO)
Vehicle destruct system safed by ground command
Extend Surveyor landing legs command
Extend Surveyor antennas command
Surveyor tronsmitter high power on command
Surveyor /Cenfaur electrical disconnect
Surveyor /Cenfaur separation
Solar panel unlocked and start stepping
Start Cenfour 180 deg turn
Start Centaur lateral thrust
wiufi »uri QCCfuiiiiion roil
Cutoff Centour lateral thrust
Primary sun sensor lock-on
Start Cenfour retro (blowdown tanks)
Solar oxis locked; start roll oxis stepping
Cutoff Centaur retro and power off
Roll axis locked in transit position
Initial DSIF acquisition (two-way lock) completed
Mark No.
Mission time
(predicted)*
Liftoff to DSIF acquisition
9
10
11
12
13
14
15
16
17
18, 19
L + 00:00:00.00
I + 00:00:02
t + 00:00:15
L + 00:02:22.72
L + 00:02:25.82
I + 00:02:31
L + 00:03:36
t + 00:03:55.35
L + 00:04:06.85
1 + 00:11:23.44
L + 00:11:50.85
1 + 00:12:01.35
I +00:12:21.85
L + 00:12:27.30
I + 00:12:32.85
I + 00:12:37.85
L + 00:13:17.85
L + 00:13:37.85
L + 00:16:32.85
1 + 00:20:42.85
DSIF acquisition to star acquisition
Initial commanded spacecraft operations
1. Command transmitter from high to low power
2. Command off accelerometer amplifiers, solar panel deployment logic, and strain gage power
3. Command rock solar panel back and forth to seat locking pin
4. Command rock roll axis bock and forth to sect locking pin
"Tlie predicted values were computed postfligtit utilizing actual
and lime of liftoff.
Mission time
(actual)
1 + 00:00:00.00
L + 00:00:58
1 + 00:01:15.7
L + 00:02:22.2
L + 00:02:25.3
L + 00:02:56.0
L + 00:03:22.9
L + 00:03:55.1
L + 00:03:57.0
t + 00:04:06.6
1 + 00:11:26.3
L + 00:11:36.2
L + 00:11:51
L + 00:12:01
L + 00:12:21
1 + 00:12:27
1 + 00:12:32.6
I + 00:12:34
1 + 00:12:38
L + 00:13:18
1. + 00:13:18
L + 00:13:38
L + 00:16:13
L + 00:16:33
L + 00:18:34
L + 00:20:43
1 + 00:22:46
L + 00:32:58
L + 00:44:33
1 + 00:46:45
L + 00:48:16
1 + 00:49:44
GMT
(actual)
(September 20, 1966)
12:31:59.824
12:32:58
12:33:15.5
12:34:22.0
12:34:25.1
12:34:55.8
12:35:22.7
12:35:54.9
12:35:56.8
12:36:06.4
12:43:26.1
12:43:36.0
12:43.51
12:44:01
12:44:21
12:44:27
12:44:32.4
12:44:34
12:44:37
12:45:17
12:45:18
12:45:37
12:48:13
12:48:32
12:50:34
12:52:42
12:54:46
13:04:58
13:16:33
13:18:45
13:20:16
13:21:44
launch oiimuth, tanked propellant weights, and atmospheric data which depend on day
JPL TECHNICAL REPORT 32-1086
139
Table A-1 (contd)
Event
Mission time
(actual)
DSIF acquisition to star acquisition (contd)
5. Perform engineering interrogation at 11(X) bit/sec
6. Command transfer to Transmitter A low pwr
Command telemetry rate reduction from IICK) to 137.5 bit/sec
Star verification/acquisition
1. llOO-bit/sec engineering interrogation over Transmitter A
2. Command Transmitter B liigh power turn on prior to star verification
3. Command transponder power off and fiigtit control preparation
4. Command execution of positive roll
5. Command transfer to Omniantenna A during roll
6. Command return to Omniantenna B during roll
7. Manual (commanded) star lock achieved
8. Command transponder power on and return to low-power operation
Pramidcourse coast phase
Command gyro drift check
Command telemetry rate reduction from 1100 to 17.2 bit/sec
Command return to 1100 bit/sec
Low-power engineering interrogation at 11(X) bit/sec
Low-power engineering interrogation at 1100 bit/sec
Command gyro speed check
Midcoursa correction
Midcourse correction sequence
1. Low-power engineering interrogation at 1100 bit/sec
2. Command transmitter high-power on
3. Command increase telemetry role from 1100 to 4400 bit/sec
4. Command roll maneuver magnitude and direction (positive roll of 75.4 deg)
5. Command roll execution
6. Command yaw maneuver magnitude (positive yaw of 110.6 deg)
7. Command yaw execution
t + 00:54:19
to
1 + 01:07:24
1 + 01:09:19
t + 04:06:38
I + 05:29:26
to
L + 05:52:35
L + 05:56:59
I + 06:01 :01
L + 06:05:34
L + 06:22:45
t + 06:34:37
L + 06:39:57
L + 06:42:21
t + 06:54:24
to
L + 09:03:22
I + 09:11:54
I + 10:35:54
L + 11:04:31
to
I + 11:11:46
L + 14:18:44
to
I + 14:29:41
L + 14:35:43
to
1 + 14:41:18
GMT
(actual)
L + 15:42:00
to
t + 15:46:10
I -f 16:04:43
t + 16:05:36
L + 16:09:16
L + 16:12:00
L + 16:15:16
I + 16:16:05
(September 20, 1966)
13:26:19
to
13:39:24
13:41:19
16:38:38
18:01.26
to
18:24:35
18:28:59
18:33:01
18:37:34
18:54:45
19:06:37
19:11:57
19:14:21
19:26:24
to
21:35:22
21:47:54
23:11:54
23:40:31
to
23:47:46
(September 21, 1966)
02:54:44
to
03:05:41
03:07:43
to
03:13:18
04:14:00
to
04:18:10
04:36:43
04:37:36
04:41:16
04:44:00
04:47:16
04:48:05
140
JPL TECHNICAL REPORT 32-1086
Table A-1 (contd)
Event
Mission time
(actual)
GMT
(actual)
Midcourse correction (contd)
8. Command propulsion strain gage power on
9. Command pressurizalion of vernier system (helium) and unlock Engine 1 to permit roll control
10. Command tlirust phase power on
n. Command desired thrust duration (9.8 sec)
12. Command midcourse thrust execution
13. Turn off thrust phase power
14. Command off power for propulsion strain gage auxiliary acceleration amplifiers, and
touchdown strain gages
15. Command return to 1100 bit/sec
Nonstandard posl-midceurse phase
Command flight control rate mode on
Command gas jet system off
Command telemetry rate reduction from 1100 to 550 bit/sec
Command transmitter high-power off
Command telemetry rate reduction from 550 to 137.5 bit/sec
Low-power engineering interrogation at 137.5 bit/sec
Postmidcourse vernier Firings 1 and 2
1. Command .transmitter high-power on
2. Command telemetry rate increase from 137.5 to 1100 bit/sec
3. Command postmidcourse vernier Firing 1 (I.975-sec duration)
4. Command telemetry rote reduction from 1100 to 550 to 137.5 bit/sec
5. Command transmitter high-power off
6. Command transmitter high-power on
/. CciMMiCiriu telemeiry rate increase from 137.5 to ilOO bit/sec
8. Command post-midcourse vernier Firing 2 (1.975-sec duration)
9. Command telemetry rate reduction from 1100 to 550 to 137.5 bit/sec
10. Command transmitter high-power off
Command flight control power off
Low-power engineering interrogation initiated at 137.5 bit/sec
Command auxiliary battery mode on
Low-power engineering interrogation initiated at 137.5 bit/sec
Command flight control power on
Command transfer to Omniontenno A followed by return to Omnianlenno B
Low-power engineering interrogation initiated at 137.5 bit/sec
Command flight control power off
Command flight control power on
High-power engineering interrogation at 1100 bit/sec
L + 16:20:22
L + 16:21:38
L + 16:22:20
I + 16:22:47
L + 16:28:02
I + 16:28:41
I + 16:28:53
t + 16:28:55
L+ 16:31:48
L + 16:42:29
L+ 16:47:23
1 + 16:51:02
I + 16:57:20
I + 16:59:46
to
1 + 17:16:43
I + 18:46:28
L+ 18:49:08
L + 18:56:25
L + 18:57:53
L + 19:02:47
I + 19:14:36
t + 19:15:31
L + 19:18:03
L + 19:19:00
1 + 19:26:16
I + 21:47:43
1 + 21:49:05
L + 23:09:09
I + 23:09:34
I + 23:33:57
L + 23:40:12
I + 24:50:14
L + 25:15:16
L + 26:37:24
L + 26:58:48
to
L + 27:16:51
(September 21, 1966)
04:52:22
04:53:38
04:54:20
04:54:47
05:00:02
05:00:41
05:00:53
05:00:55
(September 21, 1966)
05:03:48
05:14:29
05:19:23
05:23:02
05:29:20
05:31:46
to
05:48:43
07:18:28
07:21:08
07:28:25
07:29:53
07:34:47
07:46:36
07:47:31
07:50:03
07:51:00
07:58:16
10:19:43
10:21:05
11:41:09
1 1 :41 :34
12:05:57
12:12:12
13:22:14
13:47:16
15:09:24
15:30:48
to
15:48:51
Jn TECHNICAL REPORT 32-1086
141
Table A-1 (contd)
Event
Mission lime
(actual)
GMT
(actual)
Nonstandard post-midcourse phase (contd)
Command flig)it control power off for 24-min period
Post-midcourse vernier Firings 3 through 8
1. Command post-midcourse vernier Firing 3 (0.225-sec duration)
2. Command post-midcourse vernier Firing 4 (0.225-sec duration)
3. Command post-midcourse vernier Firing 5 (0.225-sec duration)
4. Command post-midcourse vernier Firing 6 (0.225-sec duration)
5. Command post-midcourse vernier Firing 7 {0.225-sec duration)
6. Command return to main battery mode
7. Command transmitter high-power on
8. Command telemetry rate increase to 4400 bit/sec
9. Command telemetry rate decrease to 1100 bit/sec
10. Command postmidcourse vernier Firing 8 (I.975-sec duration)
11. High-pov»er engineering interrogation of Modes 4 and 5 initiated
12. Command telemetry rate reduction from 1100 to 137.5 bit/sec
13. Command transmitter high-power off
Post-midcourse vernier Firings 9 through 14
1. Command postmidcourse vernier Firing 9 (0.225-sec duration)
2. Command postmidcourse vernier Firing 10 (0.225-sec duration)
3. Commend postmidcourse vernier Firing 1 1 (0.225-sec duration)
4. Command postmidcourse vernier Firing 12 (0.225-sec duration)
5. Command postmidcourse vernier Firing 13 (0.225-sec duration)
6. Command transmitter high-power on
7. Command telemetry rate increase from 137.5 to 1100 bit/sec
8. Command post-midcourse vernier Firing 14 (1.975-sec duration)
9. Command telemetry rate reduction from 1100 to 137.5 bit/sec
10. Command transmitter high-power off
Low-power engineering interrogation of Modes 4 and 5 initiated
Post-midcourse vernier Firings 15 through 20
1. Command post-midcourse vernier Firing 15 (0.225-sec duration)
2. Command post-midcourse vernier Firing 16 (0.225-sec duration)
3. Command post-midcourse vernier Firing 17 (0.225-sec duration)
4. Command post-midcourse vernier Firing 18 (0.225-sec duration)
5. Command post-midcourse vernier Firing 19 (0.225-sec duration)
6. Command transmitter high-power on
7. Command telemetry rate increase from 137.5 to 1100 bit/sec
8. Command postmidcourse vernier Firing 20 (1.975-sec duration)
9. Command telemetry rate reduction from 11(X3 to 137.5 bit/sec
10. Command Ironsmiller high-power off
Unsuccessful attempts to command A/SPP polar axis to step positive (240 times)
Post-midcourse vernier Firings 21 through 26
1. Command postmidcourse vernier Firing 21 (0.225-sec duration)
2. Command postmidcourse vernier Firing 22 (0.225-sec duration)
3. Command postmidcourse vernier Firing 23 (0.225-sec duration)
L
I -f 28:56:08
L + 31:12:59
I + 31:35:05
I + 32:03:20
I + 32:23:06
L + 32:43:12
I + 32:55:44
L + 34:48:02
L + 34:50:35
I + 34:56:08
L + 35:01:23
L + 35:02:33
L + 35:08:42
I + 35:11:31
t + 36:28:34
L + 36:33:42
I 4- 36:37:23
1 -f 36:42:41
I -f 36:47:46
I + 36:52:02
L + 36:53:06
L + 36:56:11
L + 36:57:38
I + 36:58:24
1 + 37:08:09
1 -f 37:29:19
I + 37:36:11
I + 37:41:34
L + 37:47:37
L + 37:54:06
L + 38:02:07
I + 38:03:54
L + 38:07:14
I + 38:09:16
1 -f 38:10:21
I + 38:12:58
L + 38:45:24
L + 38:51:53
L + 38:57:07
(September 21, 1966)
17:28:08
19:44:59
20:07:05
20:35:20
20:55:06
21:15:12
21:27:44
23:20:02
23:22:35
23:28:08
23:33:23
23:34:33
23:40:42
23:43:31
(September 22, 1966)
01 :00:34
01:05:42
01:09:23
01:14:41
01:19:46
01:24:02
01:25:06
01:28:11
01:29:38
01:30:24
01:40:09
02:01:19
02:08:11
02:13:34
02:19:37
02:26:06
02:34:07
02:35:54
02:39:14
02:41:16
02:42:21
02:44:58
03:17:24
03:23:53
03:29:07
142
JPL TECHNICAL REPORT 32- J 086
Table A-1 (contd)
Event
Mission time
(actual)
OMT
(actual)
Nonstandard post-midcourse phase (contd)
4. Command postmidcourse vernier Firing 24 (0.225-sec duration)
5. Command postmidcourse vernier Firing 25 (0.225-sec duration)
6. Commond transmitter high-power on
7. Command telemetry rote increase from 137.5 to 1100 bit/sec
8. Command postmidcourse vernier Firing 26 (1.975-sec duration)
9. Command telemetry rate reduction from 1100 to 137.5 bit/sec
10. Command transmitter high-power off
Command flight control power off for 24-min period
Post-midcourse vernier Firings 27 through 32
1. Command post-midcourse vernier Firing 27 (0.225-sec duration)
2. Command post-midcourse vernier Firing 28 (0.225-sec duration)
3. Command post-midcourse vernier Firing 29 (0.225-sec duration)
4. Command post-midcourse vernier Firing 30 (0.225-sec duration)
5. Command post-midcourse vernier Firing 31 (0.225-sec duration)
6. Command transmitter high-power on
7. Command telemetry rate increase from 137.5 to 1100 bit/sec
8. Command post-midcourse vernier Firing 32 (1.975-sec duration)
9. Command telemetry rate reduction from 1100 to 137.5 bit/sec
10. Command transmitter high-power off
Command flight control power off for 31 min period
Postmidcourse vernier Firing 33
1. Command transmitter high-power on
2. Command telemetry rote increase from 137.5 to 1100 bit/sec
3. Command retro sequence mode on (to obtain high thrust level)
4. Command post-midcourse vernier Firing 33 (2.5-sec duration)
5. Command telemetry rate decrease from 1100 to 137.5 bit/sec
6. Command transmitter high-power off
Command flight control power off for 1 hr 46 min period
Attempt to step solar panel in transit locked position
1. Command transmitter high-power on
2. Command telemetry rote increase from 137.5 to 1100 bit/sec
3. Command unlock of solar panel launch lock (5 times)
4. Command solar panel in multiple steps, alternating plus and minus
5. Command telemetry rote reduction from 1100 to 137.5 bit/sec
Unlocking of solar panel transit lock
1. Command telemetry rate increase from 137.5 to 1100 bit/sec
2. Command unlock of solor panel transit lock (panel slips 20 deg)
3. Command solar panel to step minus (87 times)
4. Command telemetry rote decrease from 1100 to 137.5 bit/sec
5. Command transmitter high-power off
L + 39:02:33
1 + 39:07:07
1 + 39:11:52
I + 39:13:03
1 + 39:15:56
t + 39:17:22
I. + 39:18:06
t -f 39:19:34
L + 39:45:31
1 + 39:51:53
1 + 39:57:51
L + 40:03:34
L + 40:09:20
I + 40:20:46
L + 40:21:47
L -f 40:24:12
L + 40:25:15
L + 40:26:19
L + 40:27:16
L + 41:00:59
1 + 41:02:11
1 +41:06:58
1. + 41:11:19
t + 41 :14;CS
I + 41:15:01
I +41:16:51
(September 22, 1966)
03:34:33
03:39:07
03:43:52
03:45:03
03:47:56
03:49:22
03:50:06
03:51:34
04:17:31
04:23:53
04:29:51
04:35:34
04:41:20
04:52:46
04:53:47
04:56:12
04:57:15
04:58:19
04:59:16
05:32:59
05:34:11
05:38:58
05:43:19
05:46:08
05:47:01
05:48:51
1 +42:00:45
06:32:45
I + 42:02:38
06:34:38
L + 42:03:18
06:35:18
I + 42:09:39
06:41 :39
to
to
1. + 42:17:24
06:49:24
L +42:18:33
06:50:33
L + 42:21:54
06:53:54
L + 42:22:33
06:54:33
I + 42:23:06
06:55:06
1 + 42:26:02
06:58:02
1 + 42:27:12
06:59:12
JPL TECHNICAL REPORT 32-1086
143
Table A-1 (conid)
Event
Mission time
(actual)
GMT
(actual)
Nonstandard post-midcourse phase (contd)
Post-midcourse vernier Firings 34 througli 39
1. Command transmitter high-power on
2. Command telemetry rote increase from 137.5 to 1100 bit/sec
3. Command post-midcourse vernier Firing 34 (0.225-sec duration)
4. Command post-midcourse vernier Firing 35 (0.225-sec duration)
5. Command post-midcourse vernier Firing 36 (0.225-sec duration)
6. Command post-midcourse vernier Firing 37 (0.225-sec duration)
7. Command post-midcourse vernier Firing 38 (0.225-sec duration)
8. Commond telemetry rate reduction from 1100 to 137.5 bit/sec
9. Command transmitter high-power off
10. Command Iransmiller high-power on
11. Command telemetry rate increase from 137.5 to 1100 bit/sec
12. Command retro sequence mode on (to obtain high thrust level)
13. Command post-midcourse vernier Firing 39 (21.5-sec duration)
14. Command telemetry rate reduction from 1100 to 137.5 bit/sec
15. Command transmitter high-power off
Command flight control power off for 1 hr 2 min period
Helium dump sequence
1. Command on telemetry Mode 2
2. Command transmitter high-power on
3. Command telemetry rate increase from 137.5 to 1100 bit/sec
4. Command helium dump
Battery power check under load
1. Command battery pressure logic off
2. Command flight control thrust phase power on
3. Command RADVS power on
4. Command power mode switching
5. Command RADVS power off
Main retromotor firing
1. Command on telemetry Mode 2
2. Command retro sequence mode on
3. Transmit emergency AMR signal to initiate retro engine firing sequence
4. Vernier engine ignition
5. Retro motor ignition
6. loss of spacecraft telemetry signal
L + 45:00:19
L +45:01:14
L +45:02:17
I + 45:02:27
I + 45:02:09
L + 45:03:00
(September 22, 1966)
I + 43:09:49
07:41:49
L + 43:10:50
07:42:50
L + 43:13:00
07:45:00
1 + 43:14:12
07:46:12
1 + 43:15:15
07:47:15
I + 43:16:18
07:48:18
1 + 43:17:25
07:49:25
L +43:18:23
07:50:23
1 + 43:19:17
07:51:17
L + 43:28:52
08:00:52
I +43:30:11
08:02:11
L + 43:31:09
08:03:09
1, +43:33:12
08:05:12
I + 43:36:33
08:08:33
t + 43:37:19
08:09:19
1. + 43:38:28
08:10:28
1 + 44:39:44
09:11:44
L + 44:39:50
09:11:50
I + 44:40:34
09:12:34
1 +44:41:16
09:13:16
L + 44:46:42
09:18:42
1 + 44:47:06
09:19:06
1 + 44:47:57
09:19:57
I + 44:50:16
09:22:16
to
to
L + 44:56:01
09:28:01
I + 44:58:09
09:30:09
09:32:19
09:33:14
09:34:17
09:34:27.2
09:34:28.6
09:35:00
144
JPL TECHNICAL REPORT 32- J 086
Appendix B
Surveyor Spacecraft Configuration
JPL TECHNICAL REPORT 32-1086 I45
POSTLANDING CONFIGURATION
-Z
SECONDARY SOLAR SENSOR
LEG I
ANTENNA AND SOLA
PANEL POSITIONER
PLANAR ARRAY
TV CAMERA (SU
CRUSHABLE BLOCK
(3 PLACES)
ALTIMETER/VELOCITY
SENSING ANTENNA
JPL nCHHICAU HeFO^T 31-/0^6
m-
ROLL ATTITUDE JET NOZZLE ASSEMBLY
OMNIANTENNA A
ANTENNA
RVEY)
(APPROACH)
>
!S^
KLYSTRON POWER
SUPPLY AND MODULATOR
OXIDIZER TANK
ALTIMETER/VELOCITY SENSING ANTENNA
COMPARTMENT A
FUEL TANK
VERNIER THRUST
CHAMBER 2 (FIXED)
SOLAR COLLECTOR
(THERMAL CONTROL)
LEG 2
TV PHOTOMETRIC CHART
TV CAMERA (APPROACH)
OXIDIZER TANK
TV CAMERA(SURVEY)
ATTITUDE CONTROL
PRESSURIZATION TANK (Ng)
FUEL SYSTEM PRESSURIZATION TANK (He)-
iti
IRSE CONFIGURATION
STA 166.45
EG
NOTE:
ANTENNA AND SOLAR PANEL
POSITIONER HAS BEEN OMITTED
FROM THIS VIEW FOR CLARITY
VERNIER THRUST CHAMBER I (GIMBALLED)
FUEL TANK
VELOCITY SENSING ANTENNA
COMPARTMENT B
FLIGHT CONTROL SENSOR GROUP
CANOPUS SENSOR
AUTOMATIC SUN
ACQUISITION SENSOR
SPACECRAFT
MAXIMUM
STATIC ENVELOPE
SIGNAL DATA
CONVERTER
OXIDIZER TANK
z
o
I-
<
I-
N
<
o
UJ
o
<
Q.
STA 74.98
ORIGIN OF
SPACECRAFT
COORDINATE
AXES (STA 46.855)
FUEL TANK
-Y
STA 47.48
ACCELEROMETER
AMPLIFIER
STRAIN GAGE
AMPLIFIER
LEG 3
PITCH AND YAW
ATTITUDE JET
NOZZLE ASSEMBLY
(2 PLACES)
TV PHOTOMETRIC CHART
(STA 46.855)
STA 34.45
-OMNIANTENNA B
■VERNIER THRUST CHAMBER 3 (FIXED)
STA
MAIN RETRO
ROCKET
MOTOR —
CENTAUR
HYDROGE
TANK 1
STOWED CONFIGURATION
STA 6.00
(O
I-
<
%
SPACECRAFT/CE/V7V16//?
SEPARATION PLANE
(SPACECRAFT STA 46.63)
STA 125.82
STA 138.00
ALTITUDE MARKING
RADAR ANTENNA
SPACECRAFT
FORWARD ADAPTER
STA 156.45
THERMAL
DIAPHRAGM
r
ADAPTER FIELD JOINT
MOUNTING FLANGE
ATTACH STATION
STA 172.45
SPACECRAFT AFT ADAPTER
ffr-i.
Appendix C
Surveyor II Failure Review Board Recommendafions
Recommendations herein are the result of observations
made by the Failure Review Board (FRB) during its
investigations.
The Board has been unable to determine the exact
cause of the Surveyor II spacecraft (SC-2) failure. In its
search for a single failure mode that would explain the
anomalies observed at all three legs of the vernier pro-
pulsion system, the Board defined to the best of its ability
all possible single-failure-mode causes. This search re-
quired the Board to consider (for commonality to the
three legs) all potential failure causes that could be
defined, although most of these were later assessed by
the Board to be incapable of explaining all observed
anomalies. Thus this search for a failure cause resulted
in the consideration of potential causes that singly or
collectively could explain observed anomalies when oc-
curring either simultaneously or sequentially. Inasmuch
as the Board was unable to determine whether single or
multiple failures were involved, its recommendations are
based on consideration of all potential causes that were
defined — regardless of how they would have had to
occur to produce the SC-2 failure.
Recommendations resulting from considerations of po-
tential causes of the SC-2 failure are primarily associated
with improved test procedures designed to provide con-
fidence that the potential failures will not occur on future
Surveyors. Some of these recommended procedure
changes are required because the Board considered pre-
vious procedures to be inadequate. Others might not
ordinarily be required, but are recommended because
of the unknown nature of the SC-2 failure.
The Board experienced considerable difficulty in inter-
preting much flight telemetry data and preflight test data.
In addition, certain flight data that would have been
useful to the Board was either not telemetered or was
not telemetered in commutator modes used during crit-
ical flight events. Further, the Board found that the
preflight characterization of several data channels was
inadequate to serve as a standard for comparison with
flight data. Therefore, several recommendations framed
by the Board are directed at providing a capability for
improved diagnostics.
Recommendations 1 through 5 are directed at assuring
a proper electrical interface between the vernier propul-
sion system and the remainder of the spacecraft.
Recommendations 6 through 25 are intended to ensure
cleanliness and integrity of the vernier propulsion system.
Recommendations 26 through 40 should provide im-
proved diagnostics, or generate basic data with which
test or flight data can be compared.
Recommendations 41 through 46 are directed at addi-
tional improvements in test procedures.
Additional recommendations, resulting from observa-
tions of the FRB, are included as Recommendations 47
through 51.
The Board has placed each recommendation in a class
of "Mandatory," "Desirable," or "Consider," as an indi-
cation to project management of the importance attached
thereto.
Recommendation 1 (Mandatory)
During each flow bench check of each vernier thrust
chamber assembly (TCA) at ETR, the waveforms of TCA
solenoid current shall be recorded over a range of volt-
ages as the voltage is turned on and off. These measure-
ments shall be niaue under flight pressure conditions.
The corresponding liquid flow response times shall also
be recorded at one throttle-valve position.
Reason: These measurements will ensure that the sole-
noid operated valve (SOV) on each TCA is operating
consistently within specification. Such data, when com-
pared with drive current supplied by the spacecraft flight
control system, will ensure an ample margin in each
solenoid circuit.
Recommendation 2 (Mandatory)
At the flight control sensor group level, a test shall be
performed to ensure an adequate margin in the current
supplied to the TCA solenoids. The current waveform
shall be recorded. Flight-type solenoids shall be used as
loads during this test. Dummy resistors shall not be used.
JPL TECHNICAL REPORT 32-1086
149
Reason: Improper current to the TCA solenoids can
produce erratic engine responses. This test will ensure
current that is proper at the unit level, and will provide
information against which spacecraft-level data can be
compared.
Recommendation 3 (Mandatory)
Spacecraft testing at ETR, and also at HAC, El Segundo,
as a control, should include the measurement and record-
ing of waveforms of the current to each SOV solenoid.
This current should be compared with Engineering Pro-
cessor 4 (EP-4) telemetry data, with flight control sensor
group level data, and with calculations to ensure that it
is correct. Testing should be performed at simulated
minimum-specification unregulated bus voltage to deter-
mine whether circuitry up to the solenoids provides ade-
quate margin; and the minimum acceptable current under
these conditions shall be specified to be sufficiently higher
than SOV pull-in current to provide margin in the sole-
noids.
Reason: Improper current to the SOV solenoids can
produce erratic engine responses. In the past, there has
been only qualitative checking of current to the flight
SOV solenoids.
Recommendation 4 (Mandatory)
Provisions shall be incorporated on the spacecraft to
permit assuring — by means of a low-current resistance
measurement — that electrical continuity exists through
all three SOV solenoids after the last mating of the three
SOV connectors.
Reason: The SOV connectors are last mated after pro-
pellant loading, and safety precautions preclude a check
of the mating by energizing the solenoids in a normal
manner. The recommended check will provide confidence
that proper mating has been accomplished.
Recommendation 5 (Mandatory)
Throttle valve current at each TCA shall be measured
by means of a clamp-on ammeter following the final
mating of each throttle valve connector.
Reason: This will ensure proper mating and confirm
circuit integrity.
Recommendation 6 (Mandatory)
Each TCA shall be flow-bench-checked at ETR prior
to the first mating with a spacecraft at HAC, El Segundo,
and again prior to final mating at ETR.
Reason: This procedure will provide a historical record
of TCA performance, and ensure proper operation of the
TCA immediately prior to final installation aboard the
spacecraft.
Recommendation 7 (Mandatory)
Prior to TCA installation on the spacecraft, the TCA
lines and fittings shall be X-rayed.
Reason: X-rays will detect large pieces of foreign
material that would not be detected in fluid contaminant
checks. This procedure will help establish confidence in
TCA cleanliness.
Recommendation 8 (Mandatory)
At each flow bench check of a TCA at ETR, the first
liquid that flows through each side of the TCA shall be
drawn off and checked for contamination.
Reason: This procedure vdll provide a historical record
of TCA cleanliness and ensure cleanliness shortly prior
to installation of the TCA aboard the spacecraft.
Recommendation 9 (Mandatory)
A vacuum bell jar shall be employed during the TCA
drying operation at ETR following each TCA flow check,
or following any test in which liquids are allowed to
come in contact with TCA flow passages downstream of
the shutoff valve.
Reason: Aluminum oxide was found in engines that
were not properly dried after removal from SC-3. This
improved procedure (already used successfully at Reac-
tion Motors Division, Thiokol) will ensure that all parts
of the TCA are thoroughly dried.
Recommendation 10 (Mandatory)
The last TCA flow bench check at ETR shall be per-
formed two weeks prior to required installation on the
spacecraft, and the TCA installation shall be as late as
practicable before propellant loading.
Reason: Previously, the TCA's were installed on the
spacecraft during the first two weeks the spacecraft was
at ETR. This recommended change will confirm proper
TCA performance at the latest practicable time, and
reduce the opportunity for unnecessary handling while
on the spacecraft— thereby providing increased confi-
dence.
150
JPL TECHNICAL REPORT 32-1086
Recommendation 11 (Mondatory)
Filters shall be added in the hoses used to load pro-
. pellants and gas on board the spacecraft at points as
close to the spacecraft as practicable.
Reason: Filters are presently installed in the ground
support equipment (GSE) upstream of the hoses. Con-
taminents within the hoses are not presently prevented
from entering the spacecraft vernier propulsion system.
Recommendation 12 (Mandatory)
Propellants and solvents loaded into any future space-
craft shall be sampled from the GSE and checked for
contamination before and after the loading is completed.
Reason: This procedure will ensure that the contamina-
tion level has not changed during the filling operation as
a result of flow through various lines and fittings, and
will help provide confidence in the cleanliness of the
vernier propulsion system.
Recommendation 13 (Mandatory)
Each time solvents are down-loaded, off-loaded, or
overflowed from the spacecraft, samples shall be checked
for contamination. In addition, samples of propellants
shall be taken from the overflow lines during propellant
loading and checked for contamination. The last solvent
and propellant samples for the spacecraft and all GSE
Millipore filters shall be preserved for post-mission
analysis.
Reason: This procedure will ensure the cleanliness of
the major portions of the vernier propulsion system at
the time of each off-loading, down-loading, or overflow
operation.
Recommendation 14 (Mandatory)
The two check valves in the helium pressurization
system shall be tested individually, rather than in parallel.
Reason: The present test procedure results in an indi-
cation of proper flow, even if one of the two check valves
does not open.
Recommendation 15 (Mandatory)
A test for blockage or contamination of helium inlet
lines shall be made before each attachment to a TCA.
Reason: Helium line blockage would prevent SOV
actuation. Contamination might prevent proper SOV
operation by obstructing motion of the helium solenoid-
actuated piston, by closing an inlet or exit port, or by
forming deposits on the nylon seats at inlet and exit ports.
Recommendation 16 (Mandatory)
A final helium gas leak check of the solenoid helium
pilot valve seat shall be performed after the last solenoid
actuation is completed.
Reason: The TCA solenoid is not presently checked
for leakage after it is assembled to the spacecraft and
after "click" tests are performed.
Recommendation 17 (Mandatory)
A low-level (2 to 5 psi) gas leak test of all propellant
bladders shall be performed immediately prior to pro-
pellant loading at ETR.
Reason: This test has been performed in the past at
various times during the spacecraft test cycle, but not
immediately prior to propellant loading. Performance of
the test immediately prior to propellant loading will
ensure bladder integrity later.
Recommendation 18 (Mandatory*)
A bladder integrity check shall be conducted after final
propellant loading to ensure that no bladder damage has
been incurred during loading.
Reason: Large amounts of fuel in the helium system
could affect SOV operation.
Recommendation 19 (Mandatory)
Procedures between the start of propellant loading
and encapsulation shall be formalized to require continu-
ous monitoring for propellant leaks by means of probes
inserted in the TCA throats. A vacuum shall be drawn
in an optimum fashion to sense any leakage and eliminate
traces of propellants that may have seeped through the
shutoff valve. This is in addition to monitoring require-
ments already in existence for external vernier propulsion
system (VPS) leak detection and personnel safety.
Reason: This type of testing was done periodically on
SC-2 after loading, but the procedure was not formalized
and resultant data was not recorded. Excessive oxidizer
leakage could result in injector salting, and excessive
fuel leakage could result in gumming of the fuel regulator.
*If feasible, FRB unable to suggest technique.
JPL TECHNICAL REPORT 32-1086
151
Recommendation 20 (Mandatory)
An improved method of propellant leak detection after
encapsulation is required. This method would preferably
draw samples from within the TCA throats, and alter-
nately draw a vacuum in an optimum fashion to sense
any leakage and eliminate traces of propellants that may
have seeped through the shutoff valve. In addition, the
present monitoring system shall be retained.
Reason: The monitoring system used in the past has
adequate sensitivity to detect leak rates that are poten-
tially hazardous to personnel, but does not have sufficient
sensitivity to detect lower-level leak rates that might
produce injector salting or otherwise result in abnormal
vernier propulsion system performance.
Recommendation 21 (Desirable)
In addition to the vernier system pressure-decay test
performed at half of flight pressure, a brief test shall be
performed at flight pressure to ensure system integrity.
In addition, a gas-vs-liquid leak test shall be evaluated
to finalize a test procedure that provides the best evi-
dence of system integrity and demonstrates that adequate
margins exist prior to launch.
Reason: The liquid pressure-decay test that has been
run in the past is performed at half flight pressure for
reasons of safety. Liquid leaks that might occur at flight
pressure may not thus be detected. The Board recognizes
that special safety precautions will be required to per-
form this test.
Recommendation 22 (Desirable)
The flow time required to load or unload solvents or
propellants on the spacecraft during each loading or
unloading operation shall be recorded and compared
with calculated values.
Reason: This will provide flow rate data, and has the
potential of detecting gross line blockages that might
otherwise be undetected.
Recommendation 23 (Desirable)
A filter shall be provided at the output of the helium
inlet line to each TCA.
Reason: Contaminants in the helium might affect SOV
operation.
Recommendation 24 (Desirable)
Acceptable levels of propellant leakage at various
joints in the liquid portion of the vernier propulsion-
system shall be established after final propellant loading.
Reason: In the past, the significance of any leak has
been assessed, but no formal criteria have existed.
Recommendation 25 (Desirable)
Corrosion protection shall be applied to the threads of
TCA fill and vent valves.
Reason: Aluminum oxide was found on these threads
after removal of TCA's from SC-3. (Removed TCA's
were not properly dried.)
Recommendation 26 (Mandatory)
Telemetry channels in the several commutator modes
shall be reassigned, so that data from sensors in the ver-
nier propulsion system will be present in commutator
modes during thrusting phases.
Reason: Engine temperatures were not present in
SC-2 Commutator Mode 1, which was used during mid-
course. As a result, the FRB was unable to completely
assess the performance of the vernier engines during
midcourse.
Recommendation 27 (Mandatory)
A filter shall be added between the unregulated current
shunt EP-4 and the X50 amplifier to remove effects of
ripple currents from telemetered current indications.
Reason: This channel is the primary indication of cur-
rent flow to each TCA solenoid, gas jet, and several other
items. The noise level on this channel often exceeds cur-
rent demands of many loads, making it difficult (if not
impossible) to determine whether the loads are properly
energized. In addition, the telemetry system appears to
respond to ripple currents through the EP-4 shunt in such
a manner as to bias the current indication. The FRB
devoted more time to interpreting this channel than it
did to any other telemetry channel.
Recommendation 28 (Mandatory)
Existing strain gage instrumentation on the engine
mounting brackets shall be revised or replaced with a
system that will provide improved, more reliable thrust
measurements at each TCA. The instrumentation should
be calibrated dynamically as well as statically.
152
JPt TECHNICAL REPORT 32-1086
Reason: The existing strain-gage instrumentation sys-
tem does not provide quantitative data. Gages are sensi-
tive to forces in directions not parallel to the thrust axes
and exhibit considerable zero shift. In addition, the gages
are temperature-sensitive and respond to radiative heat-
ing from the TCA's. Dynamic response characteristics of
the strain-gage mounting-bracket combination have not
been available in the past.
Recommendation 29 (Mandatory)
A dynamic calibration of flight-type engine thermal
sensors shall be performed over the expected range of
temperatures and for room and high-temperature sensor
bondings to be used.
Reason: This will permit a better interpretation of
flight data.
Recommendation 30 (Desirable)
All future TCA acceptance testing for any reason shall
include continuous recording of thermal-sensor resistance
during all hot firings and for a period of 5 min (continu-
ously) immediately subsequent to engine shutdown. Each
recording shall be included in the TCA Flight Acceptance
Test (FAT) log.
Reason: This will provide a standard for comparisons
of flight data.
Recommendation 31 (Desirable)
The dynamic thermal performance of the TCA shall
be characterized as seen by the flight-jacket temperature
sensor.
Reason: This will result in a better understanding of
TCA flight data.
Recommendation 32 (Desirable)
The maximum duty cycle of the vernier propulsion
system line and tank heaters shall be determined for
every sun level during the solar-thermal-vacuum (STV)
test.
Reason: This will provide data for more thorough
diagnosis of the VPS heater system performance.
Recommendation 33 (Desirable)
Gas jets shall be inhibited during vernier thrusting
periods either by command or an interlock in flight
control.
Reason: Uncertain conditions of gas jets during the
Surveyor U midcourse thrusting period made unregulated-
current (EP-4) telemetry data diSicult to interpret.
Recommendation 34 (Desirable)
A study shall be performed to determine whether engi-
neering, midcourse, or retro accelerometers might provide
a useful indication of total vernier engine thrust if telem-
etered during midcourse.
Reason: This telemetry information would aid in the
performance analysis of the vernier propulsion system.
Recommendation 35 (Desirable)
Telemetry of fuel-line temperatures and fuel-side pres-
sure (either fuel or helium) shall be added for commu-
tator modes used during thrusting periods.
Reason: Assessment of VPS performance on Surveyor II
was hampered by a lack of data.
Recommendation 36 (Desirable)
Separate current shunts shall be provided for loads
presently on the unregulated-current shunt (EP-4) in the
following manner: one shunt for cyclic loads, one shunt
for the gas jets and roll actuator, and one shunt for the
SOV solenoids and continuous loads.
As an alternative, the following three steps could be
taken: Inhibit gas jets during any thrusting period, telem-
eter yaw gyro heater on-off information (only pitch and
roll were telemetered on Surveyor 11), and return gyro
heater currents directly to the EP-9 shunt without passing
through the EP-4 shunt.
Reason: Either of these changes (the first being the
most desirable) will permit more accurate determination
of SOV solenoid current during both test and flight.
Recommendation 37 (Desirable)
A filter shall be added between the EP-14 current
shunt and the X50 amplifier to remove efl^ects of ripple
from the telemetry of regulated current.
Reason: The EP-14 telemetry channel is very noisy,
and it is difficult (if not impossible) to detect small cur-
rent changes in noise resulting from the response of the
telemetry system to ripple through the EP-14 shunt.
JPL TECHNICAL REPORT 32-1086
153
Recommendation 38 (Desirable)
The signal processing circuitry shall be modified to
eliminate data inaccuracies that exist at 4400 bit/sec
relative to other bit rates.
Reason: Various telemetry channels— primarily the
temperature telemetry channels — provide data that varies
as a function of the value of the preceding word in the
telemetry format at 4400 bit/sec. The phenomenon is
understood and often permits the correction of telemetry
data in non-real-time, but leads to considerable uncer-
tainty in real-time and lowers the confidence that can
be placed in the data during post-flight analysis.
Recommendation 39 (Desirable)
Yaw gyro heater on-off information shall be telemetered
(only pitch and roll have been telemetered in the past).
Reason: This will permit the more accurate correction
of current telemetry data during both tests and flight.
Recommendation 40 (Consider)
A helium tank pressure indication that is not subject to
zero shift shall be provided.
Reason: The helium pressure transducer on Surveyor II
exhibited zero shift at pressurization and at helium dump.
In addition, pressure indications were erratic at other
times during the flight. It is believed that the erratic
indications are related to the zero-shift phenomenon. The
zero shift and the erratic behavior make it difficult to
place high reliance on the helium-pressure data.
Recommendation 41 (Mandatory)
The power system current telemetry channels EP-4
and EP-14 shall be calibrated at all bit rates against
hard-line measurements and compared with expected
responses as each load is turned on and off, one at a
time (in the case of the roll actuator, for null and satu-
rated conditions). This procedure should be performed
at ETR, and at HAC, El Segundo, as a control.
Reason: This will ensure that all loads are drawing
proper currents, and that the telemetry is properly cali-
brated. This procedure — if followed on Surveyor II —
would have guaranteed, for example, that proper current
flowed to the SOV solenoids, and would have enabled
EP-4 flight data to be more accurately corrected for the
effects of several variable loads (e.g., roll actuator, gas
jets, and gyro heaters). In addition, this procedure will
generate power profile data for each spacecraft and
assure that all connections and loads are proper.
Recommendation 42 (Mandatory)
Throughout the system test cycle, when the flight
TCA's are not on board the spacecraft or not connected,
flight-type TCA solenoids shall be used to simulate sole-
noids on the flight TCA's. Dummy resistors shall not be
used.
Reason: Surveyor II test data suggests that the current
flow may be different with dummy resistors than with
inductive solenoids. This procedure will permit better
simulation of flight loads.
Recommendation 43 (Mandatory)
In the event of loss of air conditioning between pro-
pellant loading and encapsulation, standby vacuum pumps
shall be applied to each TCA within 20 min.
Reason: This will prevent injector salting that might
result from high humidity. Using desiccants in place of a
vacuum pump is not sufficient because the oxidizer itself
is hygroscopic.
Recommendation 44 (Mandatory)
During final harness inspection, it shall be determined
(and recorded) that there are adequate service loops for
vehicle vibration.
Reason: Photographs of SC-2 prior to encapsulation
indicate that the harness to the connector at the TCA 3
solenoid was taut.
Recommendation 45 (Desirable)
The resolution of the hard-line current measuring sys-
tem used to monitor spacecraft heaters during STV shall
be increased.
Reason: This will allow improved diagnosis of the
spacecraft thermal control system. In the past, resolution
was inadequate to monitor cycling of thermal heaters.
Recommendation 46 (Desirable)
Provisions shall be incorporated on the spacecraft to
permit assuring, by means of a low-current resistance
measurement, that electrical continuity exists in all VPS
line and tank heater circuits.
154
JPl TECHN/CAt REPORT 32- J 086
Reason: These circuits cannot be checked functionally
at ETR because their thermostats are not actuated by
ambient temperatures.
Recommendation 47 (Desirable)
Consideration shall be given to reducing the level of
oxidizer leakage at the TCA throttle-valve filters and
bellows areas.
Reason: This could prevent damage to the TCA ther-
mal finish.
Recommendation 48 (Desirable)
An appropriate filter shall be placed in the circuitry
driving the roll actuator motor to eliminate harmonic
components of the drive current.
Reason: The drive current to the roll actuator has a
peaked waveform which derives from a square-wave
error signal at the input to the servo amplifier. Large
current peaks pass through regulated-current shunt EP-4,
and produce "noisy" EP-4 telemetry.
Recommendation 49 (Desirable)
Wiring harness redundancy shall be added in critical
circuits such as those to the TCA solenoids.
Reason: This will improve reliability.
Recommendation 50 (Desirable)
In the event of future in-flight failures, TV pictures of
pertinent areas of the spacecraft shall be taken.
Reason: Structural damage remains an unclosed item
in the Surveyor II failure investigation. Concern over
damage in the area of Engine 3 is intensified by the
recovery of the fragment of the nose cone with its peculiar
hole. TV pictures would be invaluable in the assessment
of such damage.
Recommendation 51 (Consider)
Operations personnel shall review criteria for transmis-
sion of the emergency thrust termination command dur-
ing vernier engine thrusting periods. In addition, the
chain of command associated with making the abort
decision should be reviewed with the intent of shortening
the chain.
Reason: Surveyor II midcourse thrusting was allowed
to continue for 9.85 sec without termination. If the emer-
gency termination command had been sent sufficiently
early, the spacecraft tumbling rate could have been cor-
rected by the cold-gas attitude control system. This
probably would not have saved the mission in the case
of Surveyor II; but for some failure modes, a more rapid
response might permit the salvaging of a future mission.
JPL TECHNICAL REPORT 32-1086
155
Appendix D
Surveyor II Temperature Histories
,56 JPL TECHNICAL REPORT 32-7086
10
20
30 40
TIME FROM LIFTOFF, hr
60
70
Fig. D-1. Compartment A transit temperatures
JPL TECHNICAL REPORT 32-7 086
157
MAIN BATTERY
FLIGHT DATA
PREDICTION ENVELOPE
h
MIDCOURSE
160 r
TRANSMITTER B TWT
40-
tl
RETRO FIRE
20
30 40
TIME FROM LIFTOFF, hr
Fig. D-1 (contd)
50
158
JPt TECHNICAL REPORT 32-1086
tij
H
<
a:
UJ
Q.
LlJ
30 40
TIME FROM LIFTOFF.hr
Fig. D-2. Compartment B transit temperatures
JPL TECHNICAL REPORT 32-1086
159
120
/, ALTIMETER/ VELOCITY SENSOR PREAMPLIFIER
<
CE
UJ
O.
120
KLYSTRON POWER SUPPLY MODULATOR (KPSM)
-40
20
30 40
TIME FROM LIFTOFF, hr
70
Fig. D-3. RADVS transit temperatures
160
JPt TECHNICAL REPORT 32-1086
120
UJ
tr.
I-
<
UJ
a.
-80-
I 1_
SIGNAL DATA CONVERTER
-120
120
MIDCOURSE
rL
VELOCITY SENSOR PREAMPLIFIER
PREDICTION ENVELOPE
r
RETRO FIRE
30 40
TIME FROM LIFTOFF, hr
50
Fig. D-3 (contd)
JPL TECHNICAL REPORT 32-1086
161
30 40
TIME FROM LIFTOFF, hr
Fig. D-4. Flight control transit temperatures
162
JPl TECHNICAL REPORT 32-1086
160
CAMERA HOOD
-160
UJ
cc
t -240
<
tu 160
a.
H
MIDCOURSE
r
RETRO FIRE
CAMERA ELECTRONICS
-160
-240
20
30 40
TIME FROM LIFTOFF, hr
50
Fig. D-5. Survey TV camera transit temperatures
70
in TECHNICAL REPORT 32-1086
163
UJ
a:
z>
I-
<
q:
u
s
UJ
t-
30 40
TIME FROM LIFTOFF, hr
Fig. D-6. Vernier propulsion transit temperatures
164
JPt TECHNICAL REPORT 32-7086
280
<
CC 240
Q.
LiJ
200
VERNIER ENGINE 2
10
20
h* — RETRO FIRE
I
30 40
TIME FROM LIFTOFF, hr
50
60
70
Fig. D-6 (contd)
JPL TECHNICAL REPORT 32-J086
165
200
1 1 1
VERNIER ENGINE 3
160
-
120
^PREDICTION ENVELOPE
^^ /-
80
/
^^^^^^^^
40
•^FLIGHT DATA
10
i::
MIDCOURSE
J
w
r
RETRO FIRE
TJI^
30 40
TIME FROM LIFTOFF, hr
Fig. D-6 (contd)
_J
50
60
70
166
JPL TECHNICAL REPORT 32-1086
I r
VERNIER LINE 2
%n/
60
-PREDICTION
ENVELOPE
MIDCOURSE
100
80
60
VERNIER LINE 3
H
RETRO FIRE
30 40
TIME FROM LIFTOFF, hr
Fig. D-6 (contd)
60
JPL TECHNICAL REPORT 32-1086
167
' ■
SOLAR PANEL
'A
PREDICTION ENVELOPE
FLIGHT DATA
160
120
MIDCOURSE
1
J:
RETRO FIRE
<
IT
Ql
ui
RETRO UPPER CASE
RETRO LOWER CASE
30 40
TIME FROM LIFTOFF, hr
Fig. D-7. Miscellaneous transit temperatures
168
JPt reCHN/CAl. REPORT 32-7086
UJ
Q.
AUXILIARY
BATTERY MODE ON
20
MAIN BATTERY MODE ON
RETRO FIRE
30 40
TIME FROM LIFTOFF, hr
Fig. D-7(contd)
50
60
"T3
JPL TECHNICAL REPORT 32-1086
169
UJ
UJ
a.
UJ
30 40
TIME FROM LIFTOFF, hr
Fig. D-7 (contd)
170
JPL TECHNICAL REPORT 32-1086
Glossary
AC
A/D
AESP
AFC
AFETR
AGE
APC
A/SPP
BCD
BECO
BR
CCC
CDC
CDS
CP
CRT
CSP
CSTS
DC
DOD
DPS
DSCC
DSTF
DSS
DVS
ECPO
EMA
ESF
ESP
FC
FCSG
FPAC
FRB
FRT
GCS
Atlas/Centaur
analog-to-digital
auxiliary engine signal processor
automatic frequency control
Air Force Eastern Test Range
aerospace ground equipment
automatic phase control
antenna and solar panel positioner
binary coded digital
booster engine cutoff
boost regulator
Central Computing Complex
command and data (handling) console
computer data system
Command Preparation (Group)
Composite Readiness Test
central signal processor
Combined Systems Test Stand
direct command
Department of Defense
Data Processing System
Deep Space Communications Complex
Deep Space Instrumentation Facility
Deep Space Station
doppler velocity sensor
Engineering Computer Program Operations
(Group)
engineering mechanism auxiliary
Explosive Safe Facility
engineering signal processor
flight control
Flight Control Sensor Group
Flight Path Analysis and Command
Failure Review Board
fine resolution tracking
Ground Communication System
GSE ground support equipment
GSFC Goddard Space Flight Center
HSDL high-speed data line
ICS Intracommunications System
I/O input/output
IRV interrange vector
J-FACT Joint Flight Acceptance Composite Test
KPSM klystron power supply modulator
KSC Kennedy Space Center
LOS loss of signal
MAG Maneuver Analysis Group
MCDR media conversion data recovery (subsystem)
MCFR media conversion film recorder (subsystem)
MECO main engine cutoff
MEIG main engine ignition
MSFN Manned Space Flight Network
NASCOM NASA World-Wide Communication
Network
ODG Orbit Determination Group
ORT Operational Readiness Test
OSDP on-site data processing
OSDR on-site data recovery (subsystem)
OSFR on-site film recorder (subsystem)
OTC overload trip circuit
OVCS operational voice communication system
PA Performance Analysis (Group)
PCM pulse code modulation
PU propellant utilization
PVT Performance Verification Tests
QC quantitative command
RETMA Radio Electronics Television Manufacturing
Association
RFI radio frequency interference
RIS range instrumentation ship
RTCS real-time computer system
SDC signal data converter
JPL TECHNICAL REPORT 32-1086
171
Glossary (contd)
SECO sustainer engine cutoff
SFOD Space Flight Operations Director
SFOF Space Flight Operations Facility
SOCP Surveyor on-site computer program
SOPM standard orbital parameter message
SOV solenoid-operated valves
SRT System Readiness Test
SCAT Spacecraft Analysis Team
SPAC Spacecraft Performance Analysis and
Command (Group)
SSAC Space Science Analysis and Command
SSD subsystem decoder
SSE Standard Sequence of Events
STEA system test equipment assembly
STV solar-thermal-vacuum
TDA Tracking Data Analysis (Group)
T&DA tracking and data acquisition
TelPAC Television Performance Analysis and
Command (Group)
TPS telemetry processing system
TSAC Television Science Analysis and Command
(Group)
TTY teletype
TV-GDHS TV Ground Data Handling System
VECO vernier engine cutoff
VPS vernier propulsion system
172
JPt TECHNICAL REPORT 32-1086
Bibliography
Project and Mission
Surveyor A-G Project Development Plan, Project Document 13, Vol. 1, Jet Pro-
pulsion Laboratory, Pasadena, January 3, 1966.
Clarke, V. C, Jr., Surveyor Project Objectives and Flight Objectives for Missions
A through D, Project Document 34, Jet Propulsion Laboratory, Pasadena, March
15, 1965.
Parks, R. J., Flight Objectives for Surveyor Mission B, Interoffice Memorandum
MA&E 66-122, Jet Propulsion Laboratory, Pasadena, July 21, 1966.
"Surveyor B Post-flight Review Meeting," minutes of meeting held at JPL Octo-
ber 5, 1966.
"Space Exploration Programs and Space Sciences," Space Programs Summary
No. 37-42, Vol. VI, for the period September 1 to October 31, 1966, Jet Pro-
pulsion Laboratory, Pasadena, November 30, 1966.
Surveyor I Mission Report. Part I. Mission Description and Performance, Tech-
nical Report 32-1023, Jet Propulsion Laboratory, Pasadena, August 31, 1966.
Launch Operations
Macomber, H. L., Surveyor Block I Launch Constraints Document, Project Docu-
ment 43, Jet Propulsion Laboratory, Pasadena, June 11, 1965.
Macomber, H. L., and O'Neil, W. J., Surveyor Launch Constraints Mission B-
September 1966 Launch Opportunity, Project Document 43, Addendum No. 2,
Jet Propulsion Laboratory, Pasadena, September 13, 1966.
Macomber, II. L., Surveyor A,-G Mission Operations Plan (Launch Operations
Phase) Mission A, Project Document 58, Jet Propulsion Laboratory, Pasadena
May 16, 1966.
Centaur Unified Test Plan AC-7/SC-2 Launch Operations and Flight Plan (Sur-
veyor Mission B), Section 8.7, Report AY62-0047, Rev. B, General Dynamics/
Convair, San Diego, September 30, 1966.
Test Procedure Centaur/Surveyor Launch Countdown Operations AC-7/SC-2
Launch (CTP-INT-0004K), Report AA63-0500-004-03K, General Dynamics/
Convair, San Diego, September 6, 1966.
Bamum, P. W., JPL ETR Field Station Launch Operations Plan, Surveyor
Mission B, Engineering Planning Document 423, Jet Propulsion Laboratory,
Pasadena, July 27, 1966.
Surveyor Mission B Centaur-7 Operations Summary, TR-432, Centaur Operations
Branch, KSG/ULO, Gape Kennedy, September 12, 1966.
Surveyor B (AC-7) Flash Flight Report, Report TR-438, Centaur Operations
Branch, KSG/ULO, Gape Kennedy, September 23, 1966.
JPL TECHNICAL REPORT 32-1086 ,73
i
Bibliography (contd)
Launch Vehicle System
Galleher, V. R., and Shaffer, J., Jr., Surveyor Spacecraft/ Launch Vehicle Interface
Requirements, Project Document 1, Rev. 2, Jet Propulsion Laboratory, Pasa-
dena, December 14, 1965.
Aths Space Launch Vehicle Familiarization Handbook, Report GD/C-BGJ66-002,
General Dynamics/Convair, San Diego, February 15, 1966.
Centaur Technical Handbook, Convair Division, Report GD/C-BPM64-001-1,
Rev. B, General Dynamics/Convair, San Diego, January 24, 1966.
Centaur Monthly Configuration, Performance and Weight Status Report, Report
GDC63-0495-41, General Dynamics/Convair, San Diego, October 21, 1966.
Preliminary AC-7 Atlas-Centaur Flight Evaluation, (by staff of Lewis Research
Center, Cleveland, Ohio), NASA Technical Memorandum X-52243, NASA,
Washington, D.C., 1966.
Spacecraft System
Surveyor Spacecraft A-21 Functional Description, Document 239524 (HAG Pub.
70-93401), 3 Vols., Hughes Aircraft Co., El Segundo, Calif., November 1, 1964
(with revision sheets).
Surveyor Spacecraft A-21 Model Description, Document 224847B, Hughes Aircraft
Co., El Segundo, Calif., March 1, 1965 (with revision sheets).
Surveyor Spacecraft Monthly Performance Assessment Report, SSD 68202R,
Hughes Aircraft Co., El Segundo, Calif., September 21, 1966.
Surveyor Spacecraft System-Bimonthly Progress Report, mid-August through mid-
October 1966, SSD 68218R, Hughes Aircraft Co., El Segundo, Calif., October 24,
1966.
Surveyor H Flight Performance Final Report, SSD 68189-2R, Hughes Aircraft
Co., El Segundo, Calif., January 1967.
Tracking and Data Acquisition
Program Requirements Document No. 3400, Surveyor, Revision 10, Air Force
Eastern Test Range, Patrick Air Force Base, Fla., July 22, 1966.
Operations Requirement 3400, Surveyor Launch, Revision 3, Air Force Eastern
Test Range, Patrick Air Force Base, Fla., August 26, 1966.
Operations Directive 3400, Surveyor Launch, Revision 3, Air Force Eastern Test
Range, Patrick Air Force Base, Fla., September 1, 1966.
Project Surveyor-Sup})ort Instrumentation Requirements Document, Revision 1,
prepared by JPL for NASA, September 9, 1966.
Surveyor Project/Deep Space Netioork Interface Agreement, Engineering Planning
Document 260, Rev. 2, Jet Propulsion Laboratory, Pasadena, November 22,
1965.
,74 JPL TECHNICAL REPORT 32-? 086
Bibliography (contd)
Tracking and Data Acquisition (contd)
DSIF Tracking Instruction Manual (TIM), Surveyor Mission B, (4 volumes), Engi-
neering Planning Document 391, Jet Propulsion Laboratory, Pasadena, August
1966.
Tracking and Data-Acquisition System Pre-Flight and Post-Flight Review Surveyor
Mission B, Engineering Planning Document 438, Jet Propulsion Laboratory,
Pasadena, December 12, 1966.
Mission Operations System
Surveyor Mission Operations System, Technical Memorandum 33-264, Jet Propul-
sion Laboratory, Pasadena, April 4, 1966.
Space Flight Operations Plan-Surveyor Mission B, Engineering Planning Docu-
ment 180-S/MB, Jet Propulsion Laboratory, Pasadena, August 4, 1966 (and
revision sheets through September 16, 1966).
Final Report-Surveyor SC-2/Mission Operations System Compatibility Test,
Engineering Planning Document 436, Jet Propulsion Laboratory, Pasadena'
September 1966.
Surveyor Mission B Space Flight Operations Report, Report SSD 64257R (2 vol-
umes), Hughes Aircraft Company, El Segundo, Calif., November 1966.
Flight Path
Surveyor Spacecraft /Launch Vehicle Guidance and Trajectory Interface Schedule,
Project Document 14, Rev. 2, Jet Propulsion Laboratory, Pasadena, August 13
1965.
"Design Specification-Performance Ground Rules and Launch Periods-Suroei/or
Mission B," Specification SAO-50504-DSN, Jet Propulsion Laboratory, Pasa-
dena, February 24, 1966.
"Design Specification Surveyor/CentaurTarget Criteria Surveyor Mission A," Spec-
ification SAO-50552-DSN-A, Jet Propulsion Laboratory, Pasadena, July 25, 1966.
Surveyor Station View Periods and Trajectory Coordinates-Launch Dates August,
September, October, November, December 1966, SSD 68073R, Hughes Aircraft
Co., El Segundo, Calif., March 1966.
Cheng, R. K., Meredith, C. M., and Conrad, D. A., "Design Considerations for
Surveyor Guidance," IDC 2253.2/473, Hughes Aircraft Co., El Segundo, Calif
October 15, 1965.
Fisher, J. N., and Gillett, R. W., Surveyor Direct Ascent Trajectory Character-
istics, SSD 56028R, Hughes Aircraft Co., El Segundo, Calif., December 1965.
Winkelman, C. H., Surveyor Mission B Trajectory Design Report, SSD 64144R,
Hughes Aircraft Co., El Segundo, Calif., March 16, 1966.
Pre-Injection Trajectory Characteristics Report AC-7, GDC-BTD66-081, General
Dynamics/Convair, San Diego, July 1966.
JPL TECHNICAL REPORT 32-1086 ,75
Bibliography (contd)
Flight Path (contd)
O'Brian, W. G., Surveyor Mission B Post Injection Standard Trajectories, SSD
68169R, SSD 68157R (Appendix A to SSD 68169R), and SSD 68184R (Addendum
to SSD 68157R), Hughes Aircraft Co., El Segundo, Calif., August 1966.
Ribarich, J. J., Surveyor Mission A Preflight Maneuver Analysis, SSD 68163R and
SSD 68164R (Appendix A of SSD 68163R), Hughes Aircraft Co., El Segundo,
Calif., August 1966.
Davids, L., Meredith, C, and Ribarich, J., Midcourse and Terminal Guidance
Operations Programs, SSD 4051R, Hughes Aircraft Co., El Segundo, Calif.,
April 1964.
AC-7 Guidance System Accuracy Analysis, GDC-BKM66-003, General Dynamics/
Convair, San Diego, August 31, 1966.
AC-7 Firing Tables Data, September 1966 Launch Opportunity, GD/C-BTD66,
General Dynamics/Convair, San Diego, 1966.
Surveyor II Flight Path Analysis and Command Operations Report, SSD 64260R,
Hughes Aircraft Co., El Segundo, Calif., November 1, 1966.
^7^ jpi TECHNICAL REPORT 32-1086
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