/
This work was performed for the Jet Propulsion Laboratory, California Institute of Technology, sponsored by the National Aeronautics and Space Administration under Contract NAS7-100.
Surveyor SURVEYOR II
Spacecraft
System PERFORMANCE
FLIGHT
FINAL JPL Contract
REPORT 950056/January 1967
SSD 68189-2R
J. D. CLOUD Manager System Engineering
and Analysis Laboratory
T.
B.
VAN r
HORNE
Manage Analysis
Department
A.
R. H. LEUSCHNER Head Post Flight Analysis Section
l-
..................
"1
I
I
:HUGHES:
I I_ ..................
HUGHES SPACE AIRCRAFT SYSTEMS COMPANY DIVIB2ON
! .J
CONTENTS
Page 1.0 2.0 SCOPE AND PURPOSE OF SURVEYOR SYSTEM 2-I 2-2 2-2 I-I
DESCRIPTION 2. 1 2. Z 2.3 Surveyor Surveyor Reference SUMMARY
II Mission Objectives II Flight Configuration
3.0
SYSTEM 3. I 3.2 3. 3
Summary of Significant Anomalies System Performance Parameters Conclusions and Recommendations PERFORMANCE ANALYSIS SYSTEM PERFORMANCE ANALYSIS 4. I. 1 4. i. 2 4. i. 3 4. i. 4 General Mission Summary Trajectory Analysis Summary Spacecraft Command of Deviations From Events Summary Log EPD-180 and
3-1 3-3 3-3
4.0
SYSTEM 4. 1
4.1-1 4.1-1 4. 1-7 4.1-14 4.2-1
4.2 4.3
PRELAUNGH LAUNCH, 4. 3. 1 4. 3.2 4. 3. 3
COUNTDOWN INJECTION, AND SEPARATION
Launch Trajectory Profile Spacecraft Performance Evaluation of Vibration Data AG-7/SG-2 Boost Environment
for 4.3-4
4.4
DSIF 4.4. 4.4.
ACQUISITION l 2 Acquisition Spacecraft Predictions Performance CANOPUS 4.5-1 CORRECTION Midcourse Spacecraft Maneuver Performance Analysis
4.5
COAST PHASE AGQUSITION) MIDGOURSE 4. 6. I 4.6.2
I (INCLUDING
4.6
iii
4.7
DYNAMIC PHASE 4.7.1 4.7.2 4.7.3 4.7.4 4.7.5 4.7.6 4.7.7 4.7.8 4.7.9 4.7.10 4.7.11 4.7. 12
ANALYSIS Introduction Conclusion Computer Computer Simulation Simulation Data Simulation Simulation Midcourse
OF
SG-2
MIDGOURSE 4.7-I 4.7-4 4.7-5 4.7-8 4.7-12 4.7-16 4.7-27 4.7-28 4.7-41 4.7-45 4.7 -47 4.7-51
Simulation Simulation Input of Sun Sensor Data of Automatic Gain Control of Gyro Crossover Profile
of Initial Gyro Response Thrust Duration
Moment of Inertia Uncertainty Closed-Loop Analog Simulation References SEQUENCES Rotational
Results
4.8
POSTMIDGOURSE 4.8. 1 4.8.2 4. 8.3 4.8.4
Operational Discussion Analysis of Spacecraft Summary References
Motion
4.8-I 4.8-3 4.8-15 4.8-17
4.9
RELIABILITY ANALYSIS 4.9. l Performance 4.9.2 4.9.3 4.9.4 Performance Versus Predictions Predictions Future Reliability Reference ANAL YSIS CONTROL Introduction Anomalies Summary Analysis Reference
4.9-1 4.9-I 4.9-5 4.9-5
5.0
PERFORMANCE 5. 1 THERMAL 5.1.1 5.1.2 5.1.3 5.1.4 5.1.5 5.1.6 5.2
SUBSYSTEM 5.1-I Data 5.1-3 5. l-46 5. 1-54 5. l-62 5. 1-63
and Failure Support and Conclusions Discussion
Acknowledgements SUBSYSTEM
ELECTRICAL POWER 5.2. i Introduction 5.2.2 5.2.3 5.2.4 5.2.5 5.2.6
5.2-i 5.2-I 5.2-4 5.2-4 5. 2-30 5.2-30
Anomaly Description Summary and Conclusions Analysis References Acknowledgements LINK SUBSYSTEM Introduction Anomaly Summary Subsystem References Description and Conclusions Performance Analysis
5.3
RF DATA 5.3.1 5.3.2 5.3.3 5.3.4 5.3.5 5.3.6
5.3-I 5. 3-13 5.3-22 5. 3-27 5. 3-64 5.3-64
Acknowledgements
iv
5.4
SIGNAL 5.4.1 5.4.2 5.4.3 5.4.4 5.4.5 5.4.6 FLIGHT 5.5.1 5.5.2 5.5.3 5.5.4 5.5.5 5.5.6
PROC ESSING Introduction Anomalies Summary and Recommendations Signal Processing Analysis Reference Acknowledgements CONTROL Introduction Anomaly Description Summary and Conclusions Performance Analysis References Acknowledgement s
5.4-1 5.4-I 5.4-3 5.4.4 5.4-20 5.4-Z0 5.5-1 5. 5-7 5. 5-7 5. 5-13 5. 5-167 5. 5-169 5.6-1 5.6-7 5. 6-11 5. 6-12 5. 6-26 5. 6-Z6 5.7-I 5.7-I 5.7-2 5.7-4 5.7-11 5.7-II
5.5
5.6
VERNIER PROPULSION SUBSYSTEM 5.6.1 Introduction 5.6.2 Anomaly Description 5.6.3 Summary and Conclusions 5.6.4 Subsystem Performance Analysis 5.6.5 References 5.6.6 Acknowledgements MECHANISMS SUBSYSTEM 5.7.1 Introduction 5.7.2 Anomaly Description 5.7.3 Summary and Conclusions 5.7.4 Detailed Analysis 5.7.5 Reference 5.7.6 Acknowledgement
5.7
v
i. 0
SCOPE
AND
PURPOSE
At 12:32 GMT (05:32 PDT) on 20 September spacecraft (SC-Z) was launched from Cape Kennedy. stages of the flight, the overall performance of the with DSIF cessfully acquisition accomplished and Canopus However,
1966, the second Surveyor Throughout the early spacecraft was excellent,
acquisition and verification being sucduring the midcourse velocity correc-
tion sequence, vernier engine 3 did not respond properly, resulting in spacecraft tumbling. Subsequent attempts to correct this condition failed. Communication with the spacecraft was lost approximately 45 hours after launch when the main retro engine was fired to obtain additional engineering data.
The basic purpose of this report is to document the actual performance of this second spacecraft throughout the mission, compare its performance with that predicted by the spacecraft design, and recommend any changes or modifications that should be made in the spacecraft design or prediction models. The report is based on both real-time and postmission data analysis. Special attention will be given to the anomaly that caused mission failure, although this report in no way attempts to present the complete logic leading to the final conclusions regarding This latter task falls rightfully within the jurisdiction Board. the cause of that anomaly. of the Failure Review
l-I
2.0
DESCRIPTION OF SURVEYOR SYSTEM
The Surveyor spacecraft is designed and built by the Hughes Aircraft Company under the direction of the California Institute of Technology Jet Propulsion Laboratory for the National Aeronautics and Space Administration. It has been conceived and designed to effect a transit from earth to the moon, perform a soft landing, and transmit to earth basic scientific and engineering data relative to the moon's environment and characteristics. A brief but complete description of the Surveyor mission objectives and vehicle de._ign is given in the Surveyor I Final Performance Report (Reference l}.,. Thus only principal variations between the first and second Surveyor missions and designs will be discussed in this section. Z. l SURVEYOR II MISSION OBJECTIVES as defined
The basic objectives of the Surveyor spacecraft system, in Reference Z, were as follows: l) Primary objectives
soft landing I landing point. on the moon
Accomplish a
Surveyor
at a site east
of the
Demonstrate spacecraft with an oblique approach mately 25 degrees.
capability to soft land on the moon angle not greater than approxi-
c)
Obtain postlanding dynamics, radar surface.
television reflectivity,
pictures and and thermal
touchdown data of the
lunar
z)
Secondary to support
objective: Demonstrate capability future Surveyor missions.
of DSS-61
and
72
2-I
The secondary objective was subject to resolution of conflicts between Surveyor and Lunar Orbiter for the use of DSN facilities and support. In the event these conflicts could not be resolved, the secondary objective would have been dropped. 2. Z SURVEYOR II FLIGHT CONFIGURATION For a summary description of the major Surveyor functions and design mechanization, see the Surveyor I Final Performance Report (Reference l). All major differences in the SC-Z configuration compared with that of the first spacecraft are discussed in detail in Table Z-1. A complete listing of SC-Z control items, separated by subsystem or function, is given in Table Z-Z. 2.3 REFERENCE l) "Surveyor I Flight Performance Final Report, " Hughes Aircraft Gompany, SSD 68189R, October 1966.
2-2
TAB LE 2- i.
Item
SC-2
MAJOR
CONFIGURATION WITH SC - 1
Description
DIFFERENCES
1 ) Boost
regulator
overload
trip
circuit
In be
SC-1, the disabled The does
overload because 5C-2
trip circuit it would trip boost unless regulator the
in the boost regulator with a g-n_illiseeond has an is overload 20 to trip
had trancircuit
to
sient. that Z) a) Filter AESP Filter amplifier chokes on input to ESP and
not
trip
transient
30-milliseconds
b)
on
A/D in
converter CSP
2
nulling
Both tions on SC
of these design in temperature - 1.
improvements readouts
on
eliminate telen]etry
the which
large were
variapresent
3)
Telemetry signal
of
flight
control
return
In
SC-2
the harness accurate
flight
control voltage data on
return drops flight can control
signal be
is
telemetered to signals. provide
so
that
varying more
corrected telemetry
4)
A/SPP
pin
pullers
A/SPP pin installation All SC-Z
puller modules at AFETR. motors detents on used
were
redesigned
to
simplify
5}
A/SPP
drive
motor
drive of ball
the in
A/SPP all but
have the
roller SC-l
detents roll axis. This
instead is 6) Omnidirectional release mechanism antenna latch and a
design release been
improvement. mechanisms redesigned in and a the SC-I kickout to for omnidir=ctional the deployment antennas problem has been ,% and
SC-Z B have
prevent flight. spring
that occurred broadened,
The clevis opening has been added.
7)
Command
assignments
SC-Z engineering combine functions channels It has not would since been hut already -l he
mechanisms auxiliary of two commands so made determined the been available that for fuel
had that
been modified t_o command oxidizer dump dump.
to
fuel and
and
oxidizer
are
necessary, had SC
engineering accomplished. Command
mechanisms
auxiliary
change
SC Roll and actuator pressurize
-Z unlock VPS
Roll
actuator
unlock
0605
Pressurize Unlock
VPS roll-(lunar)
0607 0633
Spare U_xloek roll elevation (lunar) and
Unlock
elevation
0634
Spare
8)
Boost regulator
regulator filter
flight
control
SC-Z
boost to
regulator eliminate an
has
a new
filter that shunt did
on
the would
flight
control
regulator occur, not have and Vy
oscillations on the
sometimes SC-I it. reduced in SC-Z did
causing this
overload but
regulator. not need been
filter,
apparently loop gains
9)
Vx sensor
and
Vy group
gain
in flight
control
Vx
radar
attitude
have problem
to eliminate greater than units the
a potential 535 fps. had
instability
at velocities
i0)
Solder
splash
in
ESP
and
AESP
All
SC-2
have solder
the
Kit
10
modification (except the
performed spare
to central
eliminate command
splash
problem
decoder).
11
) RADVS
8idelobe
rejection
logic
Two
resistors
in
the
SC-Z
signal at which
data the
converter sidelobe
were signals
removed are
in order rejected
to lower from 28
the point to 25 db.
12)
Canopus
sun
reference
filter
change
SC-I (filter Canopus ments SC-2 1.5 the to
had factor
a
Canopus of 1.5) window,
sun to
filter
with
a
reduction any with
of
50
percent fogging measureof
coFnpensate in accordance at of 1.2.
for
possible recent
sensor of has 1.2 Canopus a filter
brightness factor the
Tucson. This problem of 79 has did °F for been not reduced materialize the SC-1 for flight. SC-2 filter. A/SPP This charge regulator stepping change in from at
because sensor on to charge pulse the prevent the
fogging
Canopus O-rings effort
temperature Ganopus possible window fogging was from dissipation drive replaced motors, on SC-2 65 in
13)
Canopus
window
The an
were of the to 40 the
changed Ganopus reduce
14)
A/SPP
pulse
duration
Battery current reduced and in and
regulator duration power
changed to
milliseconds. battery
the Q4
A/SPP were
15)
Quick
disconnects
Q3 GV4.
by
changing
valves
CV3
and
16
) Auxiliary
battery
cover
paint
pattern
The
paint to
pattern increase
of
the the
auxiliary temperature II for
battery of SC-I,
container this unit,
was which became
changed too low
during
Coast
Mode
Z-3
r
TABLE
2-2.
SPACECRAFT
UNIT
CONFIGURATION
AT
LAUNCH
Subsystem Control
or Item
Classification; Description
Control Part
Item
Number
Serial Number
Telecommunications Transmitter Transmitter Command transponder Command transponder Omnidir antenna Omnidir antenna A B receiver A receiver B 232400 12 and 231900-3 16 and 263220-4 263220-4 231900-3 15 ii 15
ec tlonal A ec tional B double
232400
21
RF single pole throw switch RF Low Low transfer pass pass
283983
13
switch filter A filter B telemetry
283984 233466 233466 290780
15 I 2 13
Amplifier, Buffer A Amplifier, Buffer B Planar Signal array
telemetry
290780
14
antenna
2323O0
15
Processing command signal decoder processor 232000-5 232200-8 233350-7
Central Central
Engineering processor
signal
Auxiliary engineering signal processor Signal processing a uxilia r y Low data rate auxiliary
264900-3
232540-I
264875-2 232106-5
2 11
Television
auxiliary
2-4
Table 2.2
(continued) or Item Classification Description Control Item Part Number Serial Number
Subsystem Control
Electrical Battery Boost Solar Main
Power charge regulator panel battery battery battery control regulator 274100 -4 12 14 2 63 64 16
274200-12 237760-3 2379OO 237921-i 273000-2
Auxiliary Auxiliary unit Auxiliary compartment Main power
battery
263730
5
switch
254112 290080 filter 290390 choke Meter
5 12
Boost regulator unregulated bus Boost regulator unregulated bus Fixed Shunt Wire-Wound Resistors current current output
12
Battery RADVS Unregulated Flight Control
988645-2 988647-i current 988645-3
778002 1 778O07
Flight Canopus
control sensor
sensor
group
235000-9 235300-2 (Part of 235000-9) II
i
Secondary Attitude Attitude Attitude Attitude Roll jet jet jet jet actuator
sun 1 2 3 gas
sensor
235450-I 235700-2 235700-3 235700-3
2 1 4 6 4 7
supply
235600-2 235900-3
2-5
Table 2-2
(continued)
Control Item Part Number Serial Numbe r
Subsystem or Classification; Control Item Description Radars Altitude marking radar RADVS signal data converter RADVS klystron power supply modulator RADVS altimeter/velocity sensor antenna RADVS velocity antenna sensor
283827-i 232908-2
13 AM-7 (9}
2329O9
AM-3
{5}
232910 AM-4
(6}
232911-I AM-3
{5)
232912 AM-3
RADVS waveguide a ssembly
Television Approach camera Survey television
(5)
284302-i
13
television
camera
284312-3 231051 B 230992
II 12
Photometric chart on omnidirectional antenna Photometric landing Propulsion Fuel Fuel Fuel tank tank tank l Z 3 tank tank tank chamber 1 chamber Z l 2 3 gear chart 2 on
12
287000-3 287001-3 287000-3 287002-3 287004-3 287003-3 285063-I
1 4 2 1 3 i 542
Oxidizer Oxidizer Oxidizer Thrust assembly Thrust assembly
285063-2
546
2-6
Table
2-2
(continued) or Item chamber 3 tank and valve 262789-2 Classification; Description Control Item l_a rt Number 285063-3 Serial Number 544
Subsystem Control Thrust assembly Helium a s sembly Retro Thermal
rocket Control
engine
238612
A21
-27
Thermal switches, compartment A
238810 238810-I 238810-3 238810-4
15, 23,
22, 44 12,
8
II, 29 I
Thermal compartment
switches, ]3
238811
7, 16, 17, 18, 19, 20
Thermal shell Compartment Compartment Thermal heater control assembly
assembly A B and
286459 286460
Compartment Compartment Thermal tray Compartment Compartment Thermostat, temperature assemblies: and 3 Thermal Compartment Tray Lower top Support
A B assembly A B heater, sensing legs l, and 2,
232210-I 232210-2
15 18
264334-I 276935
Resistors A 988653-2 988653-2 988653-2 6538 155 6542
Insulation
2-7
_J
Table 2-2 (continued) Subsystem or Classification; Control Item Description
Canister Thermal Thermal inner Thermal outside Thermal Thermal Compartment Thermal Thermal Lower Lower Tray Lower Canister Thermal inner Thermal inner Thermal Spaceframe Upper Upper Leg Retro Retro Retro Harness Crushable spaceframe spaceframe 2 upper attach attach attach web 1 2 3 1 2 988654-2 988654-2 988653-2 988653-2 988653-2 988653-2 988657-i 988653-2 988653-2 113 116 117 196 163 164 106 217 6414 switch face radiator switch contact Resistors, ring 4, switch switch contact switch ring 5, 988653-2 176 2 5, Control Item Part Number 988653-2 988653-2 988654-i0 Serial Number 6276 180 1954
switch Resistors, B switch switch spaceframe spaceframe top support
8
988653-2
6349
1 5 1 2
988653-2 988653-2 988653-2 988653-2 988653-2 988653-2 988653-2 988653-2
6380 6379 171 167 6271 6386 6269 6368
4,
988650-2
6150
tunnel block
Auxiliary battery compartment
2-8
Table 2-2
(continued)
Control Item Part Number Serial Number
Subsystem or Classification; Control Item Description
Mechanisms Instrumentation Spaceframe Landing gear and
subassembly 1 1 y 2 leg 2
264178-I 261278 (263947)
Footpad leg suba s sembl Landing gear
261279 (263947)
Footpad subassembly Landing gear
3 leg 3
261280 (263947)
Footpad subassembly Shock Shock Shock
absorber, absorber, absorber,
leg leg
1 2
264300-I 264300-I 264300-I 988684-I
9
I0 ii 989062
leg 3
Leg position potentiometer Leg position potentiometer Leg position potentiometer Omnidirectional mechanism Omnidirectional mechanism Cartridge pullers,
i 988684-I 2 988684-i 3 antenna A 287300-I 989919 989920
antenna
B
273880-i
actuated mechanical
pin
Omnidirectional Omnidirectional Antenna and positioner solar
antenna antenna panel
A B
236390-5 236390-5 287580
140 142 1
2-9
Table
2-2
(continued) or Classification; Item Description pin Control Item Part Number Serial Numb
Subsystem Control
e r
Cartridge actuated puller - A/SPP s uba s semblie s Roll Roll latch latch
293184-2 293184-3 293184-i 293184-5 293184-4 287490-9 230069-I
I 2 2 2 2 141 28, 3O 29,
Elevation Solar Solar Solar Retro panel panel panel release legs
rocket
mechanisms, i, 2, and3
Separation sensing and arming devices, legs i, 2, and 3 Cartridge pullers, actuated pin legs i, 2, and
293400
1,9,7
236390-7 3
141, 147
144,
Cartridge actuated pin puller, roll actuator Engineering auxiliary EMA board mechanism
236390-7
143
263500-6
12
4
273341 (Part of 263500-6)
F-4
Strain gage assembly, Accelerometer Accelerometer Accelerometer Accelerometer control
amplifier leg 3 I, leg 2, leg 3, leg 1 2 3
238930
239002-I 239002-2 239002-3 239002-4
17 18 19 20
4, flight group 5 A/SPP
sensor
Accelerometer mast
239002-5
21
2-10
Table 2-2
(continued) Control Item Part Number 239002-6 Serial Number 22
Subsystem or Classification; Control Item Description Accelerometer 6, velocity sensor antenna
(RADVS)
Accelerometer compartment Accelerometer compartment Accelerometer Cables Wiring A/SPP Compartment Compartment Retro TV rocket camera battery bus bus cell 1 2 voltage A B engine and B amplifier A 8, 7,
239002
-7
23
239002-8
16
239011
Harnesses
harnesses 286417 286207 286242 286390 276979 264100 3025357 286398 3025155 l 4 4 2 5 3 l 2 2
Auxiliary Basic Basic Battery RF cable Plana
assemblies r array 276828-I 261714 A 276266 261713 261711 261719 261720 switch 261712-I 261712 261714 261719-2 -2 4 4 5 10 l0 9 8 10 11 4 9
Transmitter
Omnidirectional antenna A Transfer
Planar
array
2-II
Table
2-Z
(continued) or Item Classification; Description IB Control Item Part Number 276266 261711-1 261720-1 261719-1 261721-1 coaxial -i Serial Numbe r 2 5
Subsystem Control
Transmitter
Omnidirectional antenna B
8
17 9
Accelerometer cables
239013-8 239013-I 239013-2 239013-3 239013-4 239013-5 239013-6 239013-7
16
17
18
19 20 21 22 23
2-12
3. 0 SYSTEM SUMMARY
3. i SUMMARY OF SIGNIFICANT ANOMALIES The anomalies that occurred during Mission B are summarized in Table 3-i. For this report, an anomaly is defined as an unexpected occurrence that might be indicative of a spacecraft trouble or failure. The anomalies are discussed in detail in the sections noted in this table. Eight spacecraft anomalies were designated for the flight of SC-2. The first six of these, as outlined in Table 3-i, would not have prevented the completion of a successful mission. The failure of vernier engine 3 to ignite resulted in an unsuccessful completion of the mission. The last anomaly, in which a late shutdown was probably indicated for vernier engine i during engine firing 27, could have resulted in loss of spacecraft
control.
3-i
o
I
_m
_
ce
_
m
i_
o_ °R
:
_
I,--I
0_
s_
,-1
® m
0 Z
.,4
[.-,
_
_
,
,4
....
i
_, '_-4 _ :_; r.,r] r
.__._
!
_._
_'_
;. _
o_
_t_
[4 ,q I::q
_._ o_
_a
o
o
!
r-
I_
s)
:-
I
3-2
3. Z
SYSTEM
PERFORMANCE
PARAMETERS
mission beginning
Performance analysis of each
parameters that could be determined through postof spacecraft telemetry data are given in tables near the subsystem part of Section 5. The major or significant
system performance parameters are summarized in Table 3-Z. Required or predicted values for these parameters are included in this summary for comparison purposes.
3. 3
CONCLUSIONS Conclusions Prior to the
AND
RECOMMENDATIONS
3. 3. 1
scheduled
midcourse
correction,
the
SC-2
flight
was for a the
uneventful. few almost
All spacecraft trivial instances
subsystems performed (Table 3. l, number
as designed except I-6). At midcourse,
spacecraft tumbled when only two of the three vernier engines ignited. After repeated nonstandard procedures could not regain control of the spacecraft, ignition of the retro engine (while still 18 hours away from the lunar surface) caused loss of spacecraft contact. The cause of the catastrophic failure of vernier engine 3 is under continuing investigation. 3. 3. Z Recommendations I recommendations. as a result of the 5. The SC-2
Table 3-3 is a summary of the status of Surveyor Additional operational and procedural recommendations Surveyor 55 flight are made in each subsection of Section recommendations are summarized in Table 3-4.
3-3
TABLE 3-2.
SUMMARY OF SYSTEM PERFORMANGE PAR_AMETERS
Sub_ection Paran,etvr Requi red Prt dictt d Sour_ c A _,:al Reference
Ovp
rail ..... t.ds
2) 3)
Missi.n Dry h_n,!,'d
p.yh,
ad _,,ight
_,i_ht
Launch l)
_e Launch
cti___} t1_1,, B,q and w,,_,n 12:32 0 de_s'se, 263:lt:46 (;NIT 224510E J _ <_ 0 dt, g,_s_,c ¸ 224510E _! _, 1,_4 t l3 ,I._ _e_ ¸ (pitch) (yawl s,_c (pitch) 5531 5.54g
2)
SpacecraEt rates durin_
an_uh_r b(,(,st angular at t,, separation null _eparation
< _
,l_'g:_e_ de_
3)
Spacecralt rates
4)
"I'ir,:c rates
<_0
s_.,
,,rids
224510E
5.54,/
Separation 1 ) Sdar de
to panel plr_yn;ent
acqu_slti_,n 624 tll,,e 62O se*onds s(,_ ,,nd_ 5TV _1 rt V t_._r 55,3,1 t,, El' 5843
Z) 3)
Sun Reqmred acquire
acqulsitiucl n_aneuv,,rs sun Roll Yaw
[l:::t"
5531 lq of first Delta DSS till,e 263:1Z:55:00 ' ,!,Hr*:es 5531 5.3,42
4)
"Fizne visibility: to: One Tw*,-
2i,I:_2:5i:55
way way
h, ck h,ck
< i2 < _0
:: italics :: imaes
224510E 224510E It] r_ _,', _mxt,._ ¸,¸,Is and
5342 5.342
' ransmission Trst c_mn_and 5) I'in/e Canopus 6} Tm_e acquire 7) Rull begil_ir_g revohltion 8) Tutal beginning tu Coast I ) Canopus Phase Attitude Mean sun Mean Cam,pus 2) Lin,it nlode R(.ll amplitude Pitch amplitude Yaw-average amplitude Average (time pulses) 3) Gyro Roll Pit Yaw 4) 5) blidcourse [) Mean Mean _olar OCR Correction Prenlidcours¢+ angles Roll Yaw 2) Pointing ac, ura_ v panel _Jt_tput cl_ drmit pe betw,.en r,od average avera_ ,y, h,-optical line err,,r or_.l_:_,t_,_ err,,r _r,,r:: roll an_h. oi initiation acquisition required Ca nopus frotl: o[ to angle of third Canopus to
_!
ct_
n,inutes
224510E
5.342
of
_' hours
prior
t(_
224510E
Launch
+
6H
iEDP
55,4¸4 {L _2 ', .,, ,i i_H_NI_SS} ln1_, ,,,,is 553.1 5.5¸4¸4 s and 5.5¸4,4
h_0
224510E
• 36r)
degrec_
from
<1080 degrees
5¸5,44
n,aneuver lo, kun
224510E
5.53.1 5.53.1
fr_,n_
224510E
5.53.1
(r]
}, de_
ree
224510E
0.44
0
44I
d,._ree
55_3,1
<0
6
degree
Z24510E
044
0
4_0
J,
gr('e
5.5.31
<0
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degree
224510E
0.44
5,53,1
224510E _as jet
80 pu_se
se',
64
_,
(
::.LIs+,
553.1
<1 " I • [ output 89
deg,'hr dt_/hr de_ =_ 5 hr watts
2245[0E £24510E ZA451OE dg4510E ? 3 _,_ t i_ gl hz'hr
5,531 5531 55,31 5.2.4a 52_4Z
l++aneuver
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3 )
Midcvur_e burn tinge
en_;inv
Postnlidc_ur Tunable Initial After with
s e rate midcourse 14 gas 5 minutes jet dan,ping 5.52. 5.52
3-4
TABLE
3-3.
STATUS
OF
SURVEYOR
I RECOMMENDATIONS
Paragraph h*'m_zatL_m in SSD68189R
and N,_mbe rs {see Rec_mLlnendatiun reference for delaiIM S_a_ls
4
Z
Z.
L
D
Redesign t,,e_ hanis,,,
:,Hmid:rect_onal
antenna
extend
Change
ac
c,.mplish,.d
on
Y,C-2
R,.v_se r*,c_.b, er
catibratton A(iC: signals
_,,eth,)ds
f,,,"
_pa,:_'_
raft
_p,.( daring
iat
calibration _lV f,,r _,C-4
pr:)(,.d,tre and f,,t[owing
t,_
he
f_,th_wed spa_ _., rafl
2B/
In(
hLde
AGC
m
ew'ry
cor,_,,_tat,)r
r,lod_'.
In(
t_,i,.d
in
all
:_.,,,h.s e, raft.
except
_ _,,r
_C-_
a,/,]
foltowing Change Change expected 4) Can,_pu_ dependent _en_ur on gain _he_her calibrali,:m f_gging is
_pa,
S_n( gain
_. p_ten_iat was chang,'d
fogging t,, I i
e×_s_ed Canop,l_
f,,r
SC-Z.
R*-_Lse
paint
patt_
rn
_m
allXH_ary
battery.
Paint
changed
:m
";C-Z
4
Z.Z._
I)
Gyrr_ deletl(m
speed dllrin_!
meas,_reRtent test.
llHprow
nlen:
,_r
Gyr,, i,,,pr,,ved
sp,,,,d for
signal 5,C-Z.
pr,_c,,s_lng
channo/_
w,,r,,
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Change scanmrLg
in te,
posll*ndin¢ hniq,;es
l'V
pi,
tnre
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appli,able _ and s,_hse,b,,.nl
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d,.s_r,.d rafl
,%r
0
Add jill_r
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D'-IF AGC
slation h'vels.
rept)rl_
on
phasu
Data
will
he
s:lppli,'d
at
r,'q,L*_1
,,t
'q_AC.
4.
Z.
Z.
_
1}
.SFOF of DNIY
,ligi,Jz,.d tapes
t_p, whenever
s
SILO_ald p,,ssibh,.
he
Ll_ud
inslead
SF(]F
d_git{z*'d
_ap,.s
beHlg
,_sed
R<,du(ti,m pha_e a_elag*,d
in
n,mc.'r pl,)r_
of
66-hotlr
tran_H
66-h,mr _gnals
pl,,t_ fur whi,
n,*t h
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mad*, ph,t_
fur at,-
te_*'a_:t'_rv ,_,_ ,L_,'f,;_
9
C,)pi_*s bulk rll[Hslon
of prinlur
real-lime data analys_s alaade
SF'OF
teletype available for
data posl
and
Teletype Bulk and
data
hay,' dala ,h._;r,.d
be,'r, hay,"
,,,ad_. _n_t he.I,
ava_labt_. available
print,'r a,-_: _11
4A)
Ch)._,r redt_: aw_id ,'d
cuorr]ina_i,m data i:s_ive h,,t_een
and
exchange and of efforls.
of JPL to
In,pr_v,'::_
nt_
_*_.,"
b_e_,
mad,"
and
_11
Htlghes
eX(
dllplicatlon
4B)
Pt_bli_]l FPAC analvsi
Q_I[( a_livitl,.s _ Q,H(
k
I.o_k
rep¢,r_s Klil::l_la{l"
on
SPAC
and
SPAC Ear]ter
and
Ft)AC: p,lht,, _,,,_
a-,.p,,vt_ ,1
are r,.p,Jr_s
being i_
_dltish,.d d,'s:rabh,
post_tliSSll_n
k
[.,_)k
Rep,,rt.
!':lira!hate current
fran_, d_fferen_lal
by-fra_lte a:!:pliflur
t,)rre¢t_ms cahbrati;m_
for
Sin(e slable frame
(,trr(mt _n _C-Z
all:p];fler a_ :: was v.'*'re
gain ,m _t]ll
_as SC-I.
r:,,t
s,l
fran:_'-hy-
,,,rr_,C_L,,n_
r,.,_,_,r*,rl
6A)
Iqlimmal,rec_ll,n_ capacitive fur
a,lt_):_:atL_ telemetry :,,_pu_
,mbafance channets h,_pedance
currer_! having
cora
C_,rre(tL,,,,_
hay,,
h_','n
_lit,L,i,ar.,l
8B}
Ad,lili,,na_ dependent t,,h,t_Je_ ry
,_t_ala_,_ ,,n signal.
,_ c_lrrent switch
c,,rr_et_on ii_erl for
Addll_ma[
correcti.m
n)_
vet
pr.)xided
_ ,,,r_:,lu_at,_r
7AI
Proper when
p]a, d]gi_iztng
_.,,,,,nt raw
,,f
"end telel,,etry
of
fib'" tape_.
marks
Greater l)lace,J_,'nt
cart" of
has lhes,,
been marks¸
laken
In
proper
?B)
Pr,,per 'hi1 tape_. _]ip
d_cri:,_nat,,r ' when ,i_git_zing
adi,lstm_.nt raw
_o _,,lemorrv
av,_id
"Bit
slip"
_s
s_ill
a
pr,,blem,
g)
M,_dify ad,h_on c_ding
reforn,atler of on a fi,:_ioLls d_gitizvd
pr¢,gram time tape _ags Ls
t,_
olin_inale when _i_n,,
Fhis SC-_
pr,)grall_ p,,_tmi_s_,m
fea_,:r_, data
will pr,_¢
be
e]:nnnated _ss.,_
f*:r
nelly,
gA)
Mo,e
ti,,J,'ty
Tran_nLittal
of
al_
DSIF
t_,l_c,,l_-
Time
of
trans_l&_
.
,_as
been
mq)rl)..ed
ana/ysi_
pe
r s,mn,,/,
9B)
Make ,_c_llators
avaHabh_
,lala (freq,_,'_c
on _',
D_IIF when
s/_bcarr,,,r used, et_. ).
Da_a
no_
a_ailable
and
are
still
desired
qC)
FV ,,_a,:n,'r p,'r
data f,,r
be r,, t,_an,:,..
_n_de h_ll,
a_ailahte ,,,,alL, at,'
i,_ sur_cy
a
,u,,re car,,_:ra
_L,,,e/y
N_I
applicable
tu
F,C-Z
Still
desired
t:,r
4,
d.
$
I)
Contm:_e of IIZ
,_se percent rad_r,,eler
of ,)f
a
"high une solar reading
s_n"
solar constant of 105
intensity (an per, ent}.
Plans test SC-1
for chamber and
future solar 5C-2.
S
IV
tests intensities
assume us*,d
the {_n
saRle
Eppley
z) ,
Investigate more ac(
h_ttery urate r_Lodel
parameters of the state
to
perle,it _f charge,
a
Model data
accura(v become
is available.
being
_mpr,_ved
as
more
i ]
FA more vih
r
vibration nearly
l.ve!s approxi:rLate ,_nw r,mn,,-nt.
shoL_ld the
be a,
red,_ced t_al Centaur
to
FAT SC-2
vibralic)n and
levels s:_hsequent
have spacecraft.
been
red,iced
for
r ati,m
4A)
Chang,bL,,,st vih
r* ral
Hahih_y ion ,'ff,',
n,,,d_,l t s.
t_,
allow
f*,r
les_en,,d
Model h,:l ,,,1,,:
in
being availabh..
ilnproved
as
_:ore
data
-IB)
Chang,, standard
r,,l_ahihly pr_, ,,dur,,_,
:,:o,h'l
to
allow
f_,r
non-
All,)wam y*_t included
,-
f,_r in
_/_nstandard model.
pr,,cedur_s
not
Revise
A/SPP
th*'r_Lal
prediction
model.
Revised _C-I
model performance
has
been data.
developed
based
on
6)
Revise diction
sarvey of _ostlanded
TV
them;a1 thermal
model
for
pre-
Model refinance
has
been data.
revised
based
on
SC-I
per-
performance.
3-5
TABLE
3-4.
SUMMARY
OF
SURVEYOR
II RECOMMENDATIONS
5. Ntllllber
t
5._
3 4
5.5 5.5
5
5.5
6
5.5
7
5,5
8
5.5
9
5.5
iO ]1
5.5 55
]Z
56 5.(;
14
t
5.3
2
5.5
3 4
5.5
5,5 5.5
5
6
5.5
,7 8
5.5
5.5
9
5.5
10
5.5
5.5
5
5.5
6
5.5 5.5
7
8
5.4
t
5.1
2
51
3
5.1
4
5
5,5 5.6
6
3-6
4. 0
SYSTEM
PERFORMANCE
ANALYSIS
4. i
INTRODUCTION
4. I. i
GENERAL
MISSION
SUMMARY
At 05:32 PDT on 20 September 1966, the second Surveyor spacecraft (SC-2) was launched from Cape Kennedy. Through the early stages of the flight, overall spacecraft performance was excellent, with DSIF acquisition and Canopus acquisition and verification being successfully accomplished. However, approximately 16 hours and Z8 minutes after launch, when the command to ignite the three vernier engines was sent to the spacecraft as part of the standard midcourse velocity correction sequence, vernier engine did not respond properly. The thrust provided by vernier engines I and 2 resulted in spacecraft spin at approximately 1.2Z rps. 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 it required approximately 60 percent of the available gas supply to reduce the spin rate to approximately 0. 97 rps, 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 mission. Thirty-nine subsequent attempts 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 rps. With the available power decreasing steadily, g + 45HOM. Communication 30 seconds following retro it was with engine decided to fire the main retro engine the spacecraft was lost approximately ignition. at 3
4. 1. 2
TRAJECTORY
ANALYSIS
Specific maneuver,
The earth track traced by Surveyor II is shown in Figure 4. 1-1. events, such as sun and Canopus acquisition, attempted midcourse and rise and set times for the DSIF stations, are also shown. path on 4. l-l, These the stereopremidcourse results were
Figures 4. I-Z, 4. I-3, and 4. i-4 show the trajectory graphic projection of DSS-51, -II, and -42. In Table injection and terminal conditions have been tabulated. obtained several days after the mission and are
considered
final.
4.1-I
68189-2-171
r_
H
o >
!
4
.r-t
4.1-2
--
,,D I 1 -J
W
Z_O,AZ
90"AZ 360"0EC
E
Figure
4. I-2.
AZ-EL
and HA-DEC Coordinates DSS 51, Johannesburg
Stereographic
Projection,
4.1-3
/
W
ZTO'AZ 3soo_[ ¢
Figure
4. i-3.
AZ-EL
and
HA-DEC DSS ii,
Coordinates, Stereographic Goldstone (Pioneer)
Projection,
4.1-4
"
0_ oo 0o
!
!
Figure
4. I-4.
AZ-EL
and
HA-DEC DSS 42,
Coordinates, Canberra
Stereographic
Projection,
--
4.1-5
_n
ii
ii
II
II
Z 0
co [_ .
Z 0 L)
©
--m
A
_
_
tr_
b-
S" ;4
m
_
H
ii
,i
ii
2
Z 0 k)
m
m
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Z g _ o z
o ©
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M
m
ca
I
ul _. . _ ,D _r" _ _) ,.D 0 "_ _ _r'¢, 4
m
ii
i
<
_<
Q
_
_ Ca m
4.1-6
The tracking included
predicted
premidcourse
view
periods The This
for rise table
the three
committed
stations are shown in Table 4. i-2. under the column marked "Event."
and set criteria are shows that Tidbinbilla,
Australia, did not see the spacecraft until late in the flight. Some trajectories yield a small view period for this station during the first Johannesburg pass. The time periods during which each tracking station received data from and controlled the spacecraft are also shown in this table. Figures 4. 1-5 and 4. 1-6 are plots of probe geocentric radius and velocity as a function of time from launch. Figure 4. 1-7 shows the earthprobe-moon, sun-probe-moon, and earth-probe-sun angles versus time from launch. Figure 4. t-8 shows the cone and clock angles as a function of time. The coordinate system is defined on the figure. In the normal cruise mode, the spacecraft -Z axis is aligned to the sun and the -X axis to the projection of Canopus. Figure 4. 1-9 illustrates the Centaur and Surveyor trajectories. The projection of earth trajectory is plotted on the earth's equatorial plane. The best estimate of the Centaur injection conditions was obtained from AFETR. Although considered poor (10- velocity error = 13 m/sec), these conditions are the best available. They were mapped out to 5 hours, and the Centaur/ Surveyor separation distance was calculated to be 680 kilometers. A mission design constraint states that the separation distance must be 335 kilometers by at least 5 hours after injection to eliminate possible Centaur interference during Canopus acquisition. Therefore, using this "poor" set of Centaur injection conditions, the constraint is well satisfied.
4. I. 3
SUMMARY
OF
DEVIATIONS in procedure Engineering
FROM from Planning
EPD-180 the prepared Document standard EDD-180, mission revision
sequence S/MB,
Significant changes documented in were as follows:
1)
L
+ 45M:
did signal view.
not
send indicated
cruise that
mode an
on object
command was in
because the Canopus
star sensor
intensity field of
L + 4H33M: when transferring bit rate to 137. 5 bits/sec due
from DSS-51 to lower gain manual star 17. to lockon
to DSS-72, decreased antenna at DSS-72. to acquire mode. DSS-51 lost Canopus
3) 4) 5)
L was L
+ 6H6M: necessary + 9H46M: of
unscheduled instead reduced spacecraft did an bit
use of
of
automatic rate to had
acquisition when to of DSS-72 receiver
2. bits/sec transfer
visibility L gain + 13H6M: control.
and "in-flight"
calibration
B automatic
4.1-7
,Z£ O
t) ao
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• ....
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O
•
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eq Ln ".ID
LP_ _1_ [_
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oo
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CD O
o .......
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LE_ O ,,D O
eq ",.1
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4.]-8
Figure
4. i-5.
Probe
Geocentric
Radius
Versus
Time
From
Eaunch
4.1-9
7
IH
ii; [?!
F!
fit
tli
t,*
Ht
!i_
!i!
F7 Ht
t;;
it!
i17
1:i
fl!
H.
$
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ih
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t:!
, ,
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if! $
:t:
t:!
HI
i ;+:
tL' H2 H_ !ii HI
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Figure
,t. 1-6.
t;)_-_)l)_, c;,,,:: .... .,d,'ic'
]m_l'tial 4. 1-!0
V,..!,.
;.
' .._,t.s
liu-_e
From
Launch
Figure
4. i-7.
Earth-Probe-Sun, Angles
Versus
Sun-Probe-Moon, Time From
and Launch
Earth-Probe-Moon
4.1-11
_-= ._ .... @i p_
_!ili! ,,,i: i
ii:2::T _
:.!!_ !:.i ....
_:
+
t:!"
_ttltf+_T
........... :t_t:_;,_::_[:t:r t:i:
t*+t _ttt'
;2
i;'_ii_i iili _:i: Ei_i:d,q Eili iiii il
-,!, ..... _i: i
i,. !!H !i!_
tl
ii
...._!!_iil_:,iiii!i!_ :;i!!i}i
i!:i ..........::!_ _:11
........
:_ :_
:I
ii
4_
.....ifit _" "*
_i
i!
_iii !
t:t:!!_:: :!:.ii _*; !_! ','.h i
Ii
i!i;
ITq',_:;:
5!'!1311:
.........
!i
i-i
:T!; rtr
_....
!_i]i:]i il
............ lltll
Figure 4. i-8.
E;arth ('one
:_
and Clock 4, Angles l-lZ ',,"_' ":<_ls Time From Launch
-6
oo OO ,.O
i
-5 X
-.j
<,
ILl
Z
-4
_ SURVEYOR
2 UNCORRECTED
IMPACT
L + 2 d 14 h 47 m 55 s '__ MOON AT
o
<
m
:_
-3
_-o
tr3
LU NAR
o
Z O
_.1
o
\_SURVEYOR
_ _, RAJECTORY I ATTEMPTED MANEUVER APPROXIMATE DISTANCE-758 -1 SEPARATION KM IL+5hr _ _ 4' ,mj _ MIDCOURSE
I
u Z
]
I
\
_
_
L + 16hr 26m I I .CANOPUSACQUISITION L + 6hr 38m
I
L + 16.5m SUN ACQUISITIONI_
I
/CENTAUR/SURVEYOR (_ 0 TRAJIECTORY -I PLANE'S X-AXIS, l -2 -3
I +3 +2 DISTANCE +I ALONG
EQUATORIAL 105 KILOMETERS
Figure
4. I-9.
Surveyor
and Centaur
Trajectory
in Earth's
Equatorial
Plane
4. 1-13
There were no other significant deviations correction was initiated, after which point Mission
until midcourse thrust B consisted of all non-
standard sequences. Sonde of the n_ost important sequences of the postmidcourse period are given in Table 4. I-3. A complete list of all vernier engine firings, with supplerrLentary data on the spacecraft spin rate, is found in Table 4. i-4. Finally, all periods when flight control power was turned off (never occurs in a standard flight) are given in Table 4. I-5.
4. I. 4
SPACECRAFT
EVENTS
SUMMARY
AND
COMMAND
LOG
In the preceding subsection, data concerning nonstandard events was presented, much of which will be required (especially the vernier burn summary) to follow the analyses that are presented in the remainder of the report. In the tables that follow, all other data of general interest will be given. Table 4. i-6 lists major spacecraft events, although detailed event logs will also be found in most of the subsystem analysis sections. Tables 4. I-7 and 4. I-8 give listings of all transmitter high power and thrust power intervals. Finally, the complete postlaunch command sequence, compiled from DSS digital command tapes (and thus accurate only within a l-second interval), will be found in Table 4. I-9. A complete mode and bit rate summary has not been given here, since this will be found in the RF" data link discussion (Section 5. 3), combined with a configuration log of that sabsystem.
4. 1-14
TABLE
4.
1-3.
SUMMARY FROM
OF
POSTMIDCOURSE EPD180
DEVIATIONS
Mission L I. L
Time Two were A
Decision/Output additional commanded. sequence control turned was Z-second firings To engine initiated in which Power minutes 90 To attempt to 3 problem. conserve
Reason clear vernier
+ 18H56M 4 19H18M
+ 26H12M
energy.
flight was
coast phase off periodically. 40
power
on for and off minutes.
approximately for approximately
L L I_ L L L L L
+ 31H12M 4 35ttZM _ 36t128M _ 37II29M + 38H45M _ 39H45M t + 3gH13M 38H19M
)
J
Pulsed with and a a
fire
the
engines period interval
five
times
To engine
attempt 3
to problem.
clear
vernier
0.2-second 5-minute
for firing between
fi rings.
Com_nanded array upward position.
deployment from its
of planar launch
To
illuminate reasons: for illuminate ceils the to
solar
panel 1) to
for get more and sun establishing
following energy 2) to sensor
spacecraft, secondary help in
spacecraft L _ 41HI1M Commanded on and emergency retro sequence retro eject. mode To with flight postretro L _ 421122M By panel ground squib command, was blown. unlock solar In in an attempt another achieve
orientation. a higher thrust level
less rise-time by control subsystem eject to effort sun mode.
placing in the
step solar to illuminate sensor ceils.
panel
secondary
IJ
+ 43ttI3M
Pulse
firing
engines
five
times
To
attempt 3
to problem.
clear
vernier
(0. 2 second for 1 minute between by a 20-second retro L _ 44H41M Initiated eject mode. helium
each firing) with firings, followed firing in the post-
engine
dumping
sequence.
To
obtain
a
calibration
curve function whether in signal. of
of time
pressure in order zero-shift pressure L ÷ 44H48M Flight control and RADVS thrust phase turned on. power To ble determine of
decay as to determine had telemetry if descent
occurred
helium
battery power heavy
was under load
capa-
were
supplying
terminal conditions. L t 44H59M RADVS ance with was the turned direction off in of accordthe SFOD. Bus to 17. amperes L + 46H2M Emergency to retro the spacecraft engine firing AMR signal to initiate sequence, was sent the To fire descent voltage 3
had on the retro
dropped with a load battery. engine
from of 47
[9.4
volts,
main mode.
in
normal
4.
1-15
TABLE
4. i-4.
VERNIER
ENGINE
IGINITION
SUMMARY
Burn Burn Number Ignition Time, day:hr :min: s ec 264:05:00:0Z 264:07:Z8:25 264:07:50:03 4 5 6 7 8 9 10 11 12 13 14 15 16 17 i8 19 20 21 22 23 24 25 26 27 28 29 30 31 3Z 33 34 35 36 37 38 39 40 41 264:19:44:59 264:20:07:05 264:20:35:20 264:20:55:06 264:2 I:15:12 264:23:33:23 265:01:00:34 265:01:05:42 265:01:09:23 265:01:14:41 265:01:19:46 265:01:28:11 265:02:01:19 265:02:08:11 265:02:13:34 265:02:19:37 265:02:26:06 265:02:39:14 265:03:17:24 265:03:23:53 265:03:29:07 265:03:34:33 265:03:39:07 265:03:47:56 265:04:17:31 Z65:04:23:53 265:04:29:51 265:04:35:34 265:04:41:20 265:04:56:12 265:05:43:19 265:07:45:00 265:07:46:12 265:07:47:15 265:07:48:18 265:07:49:25 265:08:05:12 fi65:09:31:59 Tin_e, seconds 9. 8Z5 1.975 i. 975 0. 225 0. 225 0.225 0. 225 0.225 i. 975 O. 225 O. 225 0.225 O. 225 O. 225 i. 975 O. ZZ5 O. 225 O. 225 O. 225 O. 2Z5 1. 975 0. 225 0. 225 0.225 0. 225 0.225 1. 975 0. 225 0. ZZ5 0. 225 0. 225 0. 225 1. 975 Z. 0. 0. 0. 0. 0. 21.
5 _'I:
Station DSS-11 DSS-42 DSS-42 DSS-51 DSS-51 DSS-51 DSS-51 DSS-51 DSS-II DSS-II DSS-II DSS-II DSS-II DSS-[I DSS-II DSS-II DSS-II DSS-II DSS-II DSS-II DSS-ll DSS- i i DSS- 11 DSS- i i DSS- I 1 DSS-II DSS - 11 DSSDSSDSSDSSDSSDSSDSSDSS-42 DSS-42 DSS -42 DSS-42 DSS-42 DSS -42 DSS -42 ii 11 Ii 11 ii 11 11
Telemetry Mode
Bit bits
Rate, 'sec
Transmitter and Power Mode B -Hi B -Hi B -Hi B -Lo B -Lo B-Lo B-Lo B-Lo B -Hi B-Lo B -Lo B-Lo B-Lo B-Lo B-Hi B-Lo B-Lo B-Lo B-Lo B -Lo B -Hi B-Lo B-Lo B-Lo B-Lo B-Lo B-Hi B-Lo B-Lo B-go B-Lo B-Lo B -Hi B-Hi B-Hi B-Hi B-Hi B -Hi B-Hi B -Hi B -Hi
Tumbling Rate, rpm 50 57
4400 [ 100 [ 1 O0 137 137 137 137 137 [ 100
58 60
1/5
1/5 1/5 1/5 1/5 1/5
137 137 137 13'7 13 '7 t l<)() 137 [37 137 137 137 [ [00 137 137 137 L37 137 [ li_0 137 137 137 137 137 I [ 00 [ 100
70.
5
115 1/5 1/5 1/5
1/5 1/5
75
1/5
1/5 1/5 1/5 i/5 i/5 1/5 1/5 [/5 1/5 1/5 1/5
8O
85.6 92. 3
ZZ5 225 225 225 225
1 1 1 1 1/5 6
[ [00 [ [ [ [ [ 100 100 i00 1_/0 100
128
5/2
1 L_)O
'::High
thrust.
4. 1-16
TABLE
4.
1-5.
FLIGHT
CONTROL
POWER
OFF
PE
RI ODS
Mission Time Off, h r : rnin Z1:48 25:15 28:55 38:11 39:20 40:27 41:17 43:38 Time Off, day:hr:min: GMT, sec Time On, hr:min:sec 12:05:57 15:09:24 17:51:50 03:15:49 04:15:55 05:30:18 07:34:49 09:13:00 GMT, Total Off hr:min:sec 1:46:14 1:22:08 0:25:02 0:32:42 0:24:21 0:31:02 1:45:58 1:02:32 Time,
264:10:19:43 13:47:16 17:26:48 265:02:43:07 03:51:34 04:59:16 05:48:51 08:10:28
4.
1-17
TABLE 4. 1-6.
Time, GMT, sec Mission h r :nqin:
MAJOR SPACECRAFT EVENTS
Time, s ec
day:hr:min: 263:12:31:59.8
Event
00:00:00
Liftoff (Note: for simplicity) Insulation Extend Extend Transmitter Separation Solar Start Primary Solar Roll Start End Manual Begin Complete Reacquir Special Gyro of axis axis of panel of sun sun lock; lock roll roll Jot legs omni panel
this
report
will
use
12:32:00
12:34:56 12:43:51 12:44:01 12:44:21 12:44:26 12:44:34 12:45:18 12:48:13 12:50:34 L2:54:46 18:37:34 19:09:38 19:11:57 19:26:24 21:35:22 21:39:23 264:01:38 03:07:43 04:44:00 04:48:05 04:53:38 05:00:02 05:03:48 05:14:29 11:41:09 265:02:44:58 06:54:33 09:13:16 09:19:57 09:22:16 09:30:09 09:30:33 09:32:19 09:34:17 09:34:27.2 09:34:28. 09:35:00 6
00:02:56 00: l 1:51
jettison Centaur Centaur command, Centaur
command, com;mand, high signal unlock _cquisition set, sot bc_in power (M-9) and
00:12:01 00:[2:21 00:12:26 00:12:34 00:13:18 00:16:13 00:18:34 00:2Z:46 06:05:34 06:37:38 06:39:57 06:54:24 09:03:22 09:07:23 13:06 14:35:43 16:12:00 16:16:05 16:21:38 16:28:02 16:31:48 16:42:29 23:09:09 38:12:58 4_:22:33 44:41:16 44:47:57 44:50:16 44:58:09 44:58:33 45:00:19 45: 45: 45: 45: 02:17 02:27 02:29 03:00
step roll lockon roll axis step
star
map mode on)
(cruise
Canopus gyro drift gyro e Canopus receiver speed check sur
acquisition check drift check
test
(AGC
calibration)
Premidcourse Premidcourse Pressurize Midcourse Rate Inhibit Auxiliary Unsuccessful Solar Helium RADVS Begin RADVS Enable Telemetry Emergency Vernier Retro Loss engine engine of data panel mode gas
and
roll
maneuver
yaw helium, velo, on jets battery attempt unloct;CO SGO Rates Off On
Magnitude (0. 25 See) FC Thrust
(5
counts)
Z;_:: 29:07 29:07 29:09 29:[i 2_:22
q
(727 3(,17
¢ Pwr
On
Interlock M/C Correction M/C _ Pwr Off P_rOff (24)
0735(2) 0737(2} 0522 i 0721
Terminate Thrust Prop
StrainGage
4.1-26
Table
4. I-9 (continued)
GMT, hrlmin:sec 03:29:26 29:31 29:36 29:41 33:48 33:54 33:58 34:05 34:06
Conmland 0512 0516 0232 0506 O510 0226 0521 3617 M0005
Function
GMT, hr: rnin: see 03:48:06 48:22 48:23 48:23 48:24 48:39
Con_nland 073_(2) 0522 0512 0516 0205 0232 0506 0504 0204 0220 0500 0107 0130 Otto 3617 0311 0300 0510 0226 0521 3617 MOO05 Thrust propSt Aux T.D. 1100 ESP Mode 1 77 Coast 7. 960 Xmtr Xfr Xmtr Interlock FC FC AESP Mode Prop Interlock Power Power Sw _ 5 kc cps Hi bits _ Accel
Function
Aux T. ESP Mode AESP Mode Pr,_pSt Interlock D,
Accel St rain Off 5 On Off I On rain
Amp Gage
5-8 Pwr
Off Off
¢ I_vr rain
Off Gage Pwr 5-8 Of( Off
Amp
StrainGage bits Off 5 On /sec Clock SCO SCO Volt i,o l_vr Fil Pwr Off Rates Off On Off /see
PwrOft
Gage
PwrOn
48:49 49:22
Magnitude (0.25 Sec) Thrust
(5
counts)
49:28 49:34
34:24 34:33 34:33 34:35 34:35 34:42 34:48 34:54 34:57 35:02 35:07 38:21 38:26 38:30 38:38 38:38
0727 3617 0721 0735 0735 0737 0522 0512 0516 0232 0506 0510 0226 0521 3617 MOO05
FC
_ Pwr
On
49:43 50:06
Interlock M/C Terminate Terminate Thrust PropStrain Aux F.D. ESP Mode AESP Mode Prop Accel StrainGage Of I 5 On Off 1 On St rain Gage Pwr On ¢ Pwr Correction M/C M/C Off Gage Amp PwrOff 5-8 Pwr Off Off (25)
50:12 50:17 51:34 51:34 04:15:55 16:14 16:22 16:29 16:43 16:43
Off On
Off 1 On Gage Pw r On
St rain
Magnitude (0. 25 Sec) FC Thrust
(5
counts)
17:11 17:31
0727 3617 0721 0735(2) 073712) 0232 0506 0510 0226 o5zt 3617 M0005
¢
Pwr
On
Interlock Magnitude (0. 25 Sec) FC Thrust {5 countB)
Interlock M/C Terminate Thrust NSP Mode AESP Mode Prop Interlock Magnitude (0.25 Sec Thrust I _ l_vr On (5 counts) ¢ Off 5 On Of( [ On St rain Gage Pwr On Correction M/C tAvr Off (281
17:31 17:33
38:53 39:07 39:07 39:09 39:14 39:19 39:23 39:26 39:31 39:36 41:34 43:46 43:52 44:44 44:57 45:03 45:40 45:50 46:48 47:03 47:03
0727 3617 07Zl 073513) 0737(2) 0522 0512 0516 0232 0506 0[05 0127 0[03 0502 0216 0205 0510 0226 0521 3617 MOIIO
¢
Pwr
On 17:40
Interlock M/C Terminate Thrust Prop Aux F.D. ESP Mode Xn_tr Xfr Xmtr 960 7. I100 AESP Mode Prop Interlock Magnitude (2. 0 Secl Thrust _ Pwr On (40 counts) eps 15 kc Sw Hi _Pwr Strain Accel StrainGage Off 5 On B Fil B Hi Volt SCO SCO P_r Pwr On Off On On Correction M/C Off Gage Amp Pwr 5-8 Off Off (26)
17:56 18:06 22:42 22:49 22:55 23:05 23:06
PwrOff
23:38 23:53 23:53 23:55 24:00 24:12 24:20 28:26 28:34 Gage Pwr On 2-8:58 28:58
0727 3617 0721 0735 0737(2) 0232 05O6 0510 0226 3617 M0005
FC
Interlock M/C Terminate Thrust ESP Mode AESP Mode Interlock Magnitude 10.25 Sec) Thrust _iC_r On (5 collnts) Off 5 On Off 1 On # Pwr Correction M/C Off IZ9)
bits/see Off 1 On
St rain
29:36 29:51 29:51
0727 3617 0721 0735 0737(2)
FC
47:35 47:56 47:56 47:59
0727 3617 0721 0735(2)
FC
interlock M/C Terminate Thrust ¢ Pwr Correction M/C OH (30)
Interlock M/C Terminate Correction M/C (27)
29:53 30:00
4. 1-27
Table
4. I-9
(continued)
hr
GMT, :i, dn:sec 04:30:21 30:28 34:38 34:45 34:59 34:59
C omn_and 0232 0506 0510 0226 3617 M0005 ESP Mode AESP Mode interlock Off 5
Function
GMT, hr:min:sec
_on_n_and 0130 )110 36[7 !3311 Xfr Xmtr Interlock PC FC Xmtr Xfr Xmtr 960 7, 1100 Mode Reset Interlock Retro l Pr cps _5 kc Sw Hi Power Power B B Sw Fil
Function
04:58:24 On 58:30 5<:16
Lo
l_vr Pwr Off
Off l On
Off On Fil Hi Volt SCO SCO Pwr l%vr On Off On On
05:30:18 (5 counts) 30:52 32:53 @ Pwr On 32:5_! 33:52 Correction M/C ¢ Off 5 On Off I On Pwr Off (31} 34:0b 34:11 34:4q 37:08 38ff_g 38:%% 39:22 39:4:" 15 counts} 40:27 4[:{4
X_00 0105 0127 0103 0_02 0-)16 0!05 O,a07 0720 3617 0724 (1_2 h715 0732, (1727 -,6i7 072l 0735(2)
Nlagnitude {0. 25 See) FC Thrust
35:18 35:33 35:34 35:35 35:41 35:52 36:00 40:22 40:29 40:40 40:41
0727 36[7 0721 0735(2) 0737(2) 0232 0506 0510 0226 3617 M0005
Interlock M/C Terminate Thrust ESP Mode AESP Nh_de Interlock Magnitude (0. 25 Sec) FC Thrust
bits/sec 6 On IV
Group
Sequence _pStrainGage Delay
Mode PwrOn Mode Retro _ Pv_rr On Eject On
On
Manual Emergency FC
Thrust
41:04 4l:19 41:20 41:21 41:26 41:37 41:45 45:05 52:36 52:46 53:28 53:40 53:47 54:26 54:38 55:03 55:03 i i
0727 3617 0721 0735(3) 0737(2) 0232 0506 0[05 0127 0103 0502 0116 0305 05[0 0226 36[7 M0 l l 0
_
Pwr
On 43:1_i Interlock M/C Terminate (after 2. Correction M/C 5 Sec) ¢ Pwr Group Mode St rain Accel StrainGage 5 On bits/sec _Clock SCO SCO Volt Lo Pil Pwr Pwr Off Rates Off On Off On Gage Amp Pwr 5-8 Off Off Off IV (34)
interlock 43:lq M/C Terminate Thrust ESP Mode Xn_tr Xfr Xmtr 960 7. 35 cps kc Sw Hi Off 5 B B On Pil Hi Volt SCO SCO Pwr Pwr On Off On On _ Pwr Correction M/C Off 43:47 44:01 44:20 44:55 45:00 45:0-i 45:20 46:08 46: 46:2b 46:34 47:0_ (40 counts } 47:13 47:i() _ I 4 0504 0204 0320 0500 0107 ,I130 ,1110 5017 03 l I 41 0105 {)127 0103 0503 0216 0205 Xmtr Xfr Xmtr 960 7. _5 Sw Hi cps kc B Fil B Hi Volt SCO SCO Pwr Pwr On Off On On 0737(2) 0720 070I 0521 Ogl2 (3Z) 43:22
Thrust Reset Rate Prop Aux T.D. kIode 137 Coast 7. 960 Xn_tr Xfr Xn_tr Interlock PC Power Sw
PwrOff
1100 AESP Iv_ode Inte
bits/sec Off l On r lock
_5 kc cps Hi
Magnitude (2 0 Sec) FC Thrust
55:26 56:11 56:12 56:15 56:2l 56:30 56:30 56:31 56:31 56:49 56:59 57:15 57:21 57:55 58:05 58:19
0717 3617 0721 0735(Z) 0737(2) 0522 05t2 051b 0205 0Z32 0506 0504 0204 0220 0500 0107
_ Pwr
On 4,_:51
Interlock 4_:51 M/C Terminate Thrust P r_,p Aux F.I). 1100 ESP Mode [37 Coast 7, 960 Xn,tr _;, kc cps Hi _ Pwr St rain Accel St rain Correction lvI/C Off 32:3_ Gage Amp Gage pw 5-8 P_ r Off 32:45 Off 34:1t r Off 34:/_ 34:3_ Off 35:I4 5 On bits / sec ¢Clock _GO SCO Volt Rates Off On Off to 40;27 41:39 to 41:44 4Z:5i t o 43:47 0401 (l[O) Step Minus Solar Panel t_401(10) Step Solar Panel Plus 0631(5) Unlock (Transit) Solar Pane[ (33) Day 265 DSS Off
06:30:2'_
bits/sec
i 100
blts/sec
4. I-28
Table
4. I-9 (continued)
Command 06:45:4l to 45:46 46:07 to 46:1l 46:32 to 46:37 46:55 to 47:00 47:[7 to 47:23 47:43 to 49:35 50:33 50:40 50:54 51:1I 53:32 53:47 53:54 54:33 0402(34) Step 040l(10) Step 0402(10) Step 0401(10) Step 0402(10) Step 0401(10) Step
Function
hr:
GMT, rnin:
sec
Con_ll_and 0721 0735(3) 3617 M0005
Function
07:46:12 Solar Panel Plus 46:14 46:45 Solar Panel Minus 46:45
M/C Tern_nate Interlock
Correction M/C
(36)
Magnitude (0. 25 Sec) Interlock M/C Te Correction rnainat
(5
counts
)
47:14 Solar Panel Plus 47:15 47:17 Solar Panel Minus 47:48 47:49
3617 0721 0735(2) 3617 MO005
(37)
e M/C
Interlock Magnitude (0. 25 Sec) Interlock M/C Terminate Interlock Magnitude (0. 25 Sec) Interlock M/C Terminate Thrust ESP Mode 137 Coast 7. 960 Xrntr Xfer Xn_tr Xmtr Xfer Xmtr 960 7. 1100 Mode Reset Inter Retro PropS[ Manual Enable Emergency FC Thrust cps _,5 kc 35 Off 5 On hits/see ¢ Clock kc cps Hi Sw Fil B Sw Ill SCO SCO Volt B Lo Rates Off On Off I:_,r _ Pwr Correction M/C Off (39) (5 counts) Correction M/C (38) (5 counts)
Solar
Panel
Plus
48:18 48:18
3617 0721 073512) 3617 M0005
Solar
Panel
Minus
48:20 48:50
0504 0204 0220 0500 0502 0216 0205 0635
137 Coast 7. 960 960 7. 1100 Unlock (Lunar) _5 35
bits/sec 48:50 ¢Clock kc cps cps kc SCO SCO SCO SCO Rates Off On Off On 49:25 49:25 49:27 :49:31 49:54 Panel 50:01 50:23 50:29 3617 0721 0735(2) 0737(21 0232 0506 0504 0204 0220 0500 0107 0130 0110 0105 0127 0103 0502 0216 0205 0507 0720 3617 0724 0521 0715 0706 0732 0727 3617 5 0 0721 0735(3)
bits/see Solar
55:06 to 55:49 58:02 58:09 58:24 58:41 59:12 59:21 59:28 07:34:49 39:54 41:42 41:49 42:23 42:38 42:50 43:19 43:26 43:43 44:08 44:06 0504 0204 0220 0500 0107 0130 0110 0300 0105 0127 0103 0502 0216 0205 0510 0226 0521 3617 M0005 137 Coast 7. 960 Xmtr XIr Xmtr FC Xmtr Xfer Xmtr 960 7, _5 II00 AESP i Mode cps kc Sw _5 kc cps Hi Lo FII Power B Sw Hi Fil B Hi bits/see ¢Clock SCO SCO Volt Pwr Pwr On Pwr Pwr On Off On On Off Rates Off On Off 0402(87) Step Sokar Panel Minus 50:47 50:54 51:17 5l:24 51:38 59:04 08:00:46 00:52 01:12 01:26 02:11 02:28 02:46 03:08 03:09 03:23 03:38 Gage Pwr On 03:59 04:16 (5 countB) 04:31 05:12 ¢ Pwr On 05:12. 05:34. Correction M/C (35) 05:45 06:01 (5 counts) 06:10 07:56 08:33
]:%vr Off Fil B Hi Pwr Pwr On Off On On
Volt SCO SCO
hit_/see 6 On Group lock Sequence rain Delay Gas Jets Retro ¢ Pwr Eject On Gage Mode Pwr Mode On On On IV
Volt SCO SCO
bits/sec Off i On
PropStrain Interlock Magnitude (0.25 See} FC Thrust
Interlock M/C Terminate (after Z1, 5 Correction M/C Sec) Off IV On (40)
44:36 44:59 45:00 45:02 45:5l 45:52
0727 3617 0721 0735(2) 3617 M0005
Interlock M/C ] Terminate Interlock Magnitude (0. 25 Sec) Interlock
0737(21 0720 0701 0506 0504
Thrust Reset Rate Mode 137
¢ Pwr Group M{_de 5 On bits/sec
46:I2
3617
4, 1-29
Table
4. I-9
(continued)
GNIT, hr:min:sec
Conu>and
Function
G M T, hr:min:se,
2,:::::and 0"- .!{ 0320 ,15 I 7 I)323 ,)322 1320 332 5323 i617 0630 070i_ ),16 C23 _617 k11500 i tti Current
Function
08:08:39 08:54 09:02 09:19 0q:28 0cq35 10:28 lO:Z8 09:09:41 11:44 11:50 t2:14 12:22 12:34 13:00 13:16 13:16 18:4Z
0204 0220 0500 0107 0130 0110 3617 0311 0105 0127 0103 0502 0216 0205 0300 3617 0610 0302
Coast 7. 96!) Xn,tr Xfer Xrrttr Interlock FC X:_:tr Xfer Xt_tr 960 7 1100 FC Int 5a tps _ kc cps
¢ Clock SCO >;CO IIi 5;w Fi] "&_lt B Lo Pwr
Rates Off On Oft Pwr Off
09:23:4i, 24:24 24:54 2 :.:29 27:0{! 37:27 27:43
Mode Main 13art On
On Mode
Restore Aux lli Hi Batt Current Current
Mode Mode Mode Main Batt
Ot_ On Mode LcJgic Off
Restore Disable Hi Current
Batt Xfer Mode
P<,wer 13 Sw Hi Pil
Off Pwr Pwr On Off On On
28:01 30:09 50:0(? 30:33 30:53 31:12 31:51
Interlock RADVS Enable Manual Reset Interlock Magnitude {8. 0 See} AESP Mode Off 2 On IV (160 counts) Power Gas Lock N_mL Jets On Thrust Bias Oil
13 tli Volt SC() SCO
k(
bits/sec Power t.r]c_ck On
31:51
3Z:ll Dui>p Disable I__}gic FC i Thrust He liun_ 32:19 Battery Pressure 3 p.:q q _ Pwr On 33:14 33:i4 Power Main On Batt Diode 34:17 3%00 _7 .'4 )75 0 720
Reset Interlock Retro Emergency Loss END
Group
19:06 19:56 19:57 22:16
0727 3617 0637 0320
Interlock RADVS Restore
Sequence AMR Coi_llHand MISSION
Mode Signal Link:
On
ul OF
4. 1-30
4. 2 PRELAUNCH COUNTDOWN
The final prelaunch countdown proceeded smoothly with the exception of one reported difficulty. The telemetry indicated approximately l0 to 17db weaker signal strength into receiver B than into receiver A. During prelaunch, it was felt that this failure was due to a change in the RF link with the gantry moved back (i. e., no RF repeater used under these conditions). The failure was thus attributed to a change in the test setup. Subsequent analysis of flight data showed that the failure was probably due to a shift in the receiver B automatic gain control curve. It was felt that the problem was not serious enough to prevent launch. After encountering difficulty in pressurizing the Atlas propulsion system, the spacecraft was finally launched just before close of the daily launch window at 12:32 GMT at an azimuth of 114. 361 degrees.
4.2-i
4. 3 LAUNCH,
INJECTION,
AND SEPARATION
4. 3. 1 LAUNCH TRAJECTORY PROFILE SC-2 was launched from AFETR launch site 36A on Tuesday, 20 September 1966, using a General Dynamics/Convair Atlas/Centaur (AC-7) boost vehicle. The launch was held until near the close of the launch window when difficulties were experienced with the Atlas boil off and LOX topping valve. Liftoff occurred at 12:31:59. 824. Two seconds after liftoff the launch vehicle began a 13-second programmed roll that oriented the vehicle from a pad-aligned azimuth of I05 degrees to a launch azimuth of 114. 361 degrees. At 15 seconds, a programmed pitch maneuver was initiated. The nominal and actual times for the Atlas/Centaur boost phase events are summarized in Table 4. 3-1. All mark times were nominal. The launch phase ascent trajectory profile is illustrated in Figure 4. 3-1. Separation of Surveyor from Centaur occurred at 12:44:32. 6 at a geocentric latitude and longitude of 12. 9 and 309. 8 degrees, respectively. The spacecraft was in sunlight at separation and never entered the earth's shadow during the transit trajectory. 4. 3. 2 SPACECRAFT PERFORMANCE The boost phase was normal and resulted in SC-2 being injected properly, thereby placing the spacecraft on the desired lunar trajectory to the moon target site. Subsequent to injection and prior to its separation from the spacecraft, Centaur issued the preprogrammed commands for extending the spacecraft landing legs (L + IIM51S), extending the omnidirectional antennas (g + 12MIS), and turning on the transmitter high power (L + 12M23S). Normal response was verified from telemetry data;:". A minor spacecraft anomaly occurred during this period when the flight control subsystem switched from rate to inertial mode. However, this anomaly had no effect on the mission since the flight control subsystem was returned to the rate mode by separation of the spacecraft from the Centaur. Separation was initiated by the Centaur at L + 12M27Sby accomplishing electrical disconnect, and was completed satisfactorily at L + 12M33S. ':_Anapparent anomaly was noted in real time, as the leg I position signal (V-5) indicated 17 degrees in the extended position. However, postmission analysis shows that the SFOF computer had an incorrect coefficient for this signal. 4.3-I
o_ oo oo _o
! GO O
SI/C
SEPARATION
%
Figure
4.3-I.
Launch
Phase
Trajectory
Profile
4.3-2
TABLE 4. 3-1.
MARK EVENTS
Nominal Actual Time, seconds 0.0 142. 29
Ma rk N umber
2-inch Booster staging Jettison Jettison Jettison Sustainer depletion) motion
Event (iiftoff 12:31:59. 824 GMT)
Time, seconds 0.0 143. 0
engine cutoff (guidance acceleration 5. 7 g) booster package panels
discrete,
146. 177. 204.
1 0 0
145.75 176.06
insulation nose
fairing cutoff (by propellant
202.90
235.17
engine
236. 0
6 7
Atlas/Centaur Start (SECO
separation
238. 0 Z47. 5
237. 246.
03 58
Centaur main engines + ii. 5 seconds) main engine cutoff (guidance
Centaur discrete)
684.
0
686.
3
9
I0
Surveyor Surveyor command Surveyor Centaur/Surveyor Separate Admit
landing
gear
extend
command extend
715. 0 725. 0
710.7 720.7
omnidirectional
antenna
Ii 12 13 14 15
high
power
transmitter electrical
on disconnect
746. 0 752. 0 753. 7 758. 0 0
741.
4
742. 08 752. 58 754. 7 NA
spacecraft guidance 2 engines mode 2 engines, mode (V), 180-degree
Start H20 turnaround Stop H20 turnaround Start Stop Energize
798.
16
180-degree
818.0
NA
17
retrothrust retrothrust power
(Centaur
tank
blowdown)
993. O 1243.0
992.8 1242. 1242.9 9
18
19
changeover
switch
1243.0
4.3-3
/'I
J
Following separation, the spacecraft performed the designed autosequences. By using the cold gas jets which were enabled at separathe flight control subsystem nulled out the rotational rates imparted by springs after and initiated a roll-yaw sequence to acquire a n_inus roll of approximately 72 degrees and the sun. a plus
matic tion,
the separation At L + 16M15S,
yaw of 16. 5 degrees, sun acquisition and lockon were completed. Concurrent with the spacecraft sun acquisition sequence, the A/SPP stepping sequence was initiated for deploying the solar panel axis and roll axis of 85 and 60 degrees, respectively. At approximately L + 23M, stepping was completed, resulting in positioning of the solar operations were confirmed panel to the desired in real time from transil position. All these the spacecraft telemetry.
4. 3. 3
EVALUATION ENVIRONMENT Instrumentation
OF
VIBRATION
DATA
FOR
AC-7/SC-2
BOOST
4. 3. 3. 1
Two
accelerometer
channels
(IRIG
channels
14 and
17) of vibration
data were recorded launch of SC-2/AC-7
in real time on a direct write oscillograph during the and also for the initial 5 minutes of powered flight.
Telemetry channel 17 transmitted the continuous signal of accelerometer CY 52 0. Channel 14 produced a commutated signal from CY 53 0, CY 54 0, CY 77 0, and CY 78 0. Accelerometers CY 52 0, CY 53 0, and CY 54 0 were located on the spacecraft at the legs i, 2, and 3 column bases, respectively, with their axes of maximum sensitivity parallel to the spacecraft Z axes. Accelerometer CY 77 0 was located on the upper flange of the Centaur adapter adjacent to leg i and was sensitive to motion in a radial direction. Accelerometer CY 78 0 was mounted in the flight control sensor greup and sensed the vertical response of this unit. The SC-2 dynamic instrumentation was identical to the instrumentation aboard SC-1/AC-10. 4. 3. 3. 2 Evaluation of Data CY and 54 Anomalies CY 78 0 were the only SC-2 accelerom53 0, period.
Accelerometer
0 and
eters that operated normally in flight. Accelerometcrs and CY 77 0 produced no intelligible data during the
CY 52 0, CY entire recording
Since the two operating accelerometers were commutated on an equal time basis with two inoperative transducers, the flight environment was monitored only during 36 of the 90 commutator segments or 40 percent of the time. Most of the shock transients experienced on SC-I during various jettisons and shutdown events were not recorded during the SC-2 flight and, therefore, only very vibration SC-2 limited data are available for levels recorded during similar and SD-2 (AC-6). comparison. flight events Table 4. 3-2 presents for SC-I (AC-10),
(AC-7),
4.3-4
%
u'3
oO
o • "_ _ _.,
c_
c_
c_
%
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4.3-5
4. 4
DSIF
ACQUISITION
4. 4. i
ACQUISITION
PREDICTIONS Surveyor reported II rise at DSS-51 at 1Z:55:00 on good one-way data at 12:55:17, auto-
20
Predictions indicated September 1966. DSS-51
track on the SCM (antenna main beam) at 13:00:07, and good two-way data at 13:05:07; thus, DSS-51 required 10 minutes from spacecraft rise to obtain good two-way data. Had the acquisition been optimum, DSS-51 could have had good data, two-way lock at approximately 12:58, or about 3 minutes after spacecraft rise. In comparison, DSS-51 reported good data, two-way lock less than 4 minutes after spacecraft rise in the Surveyor 1 mission. This delay of about 6 minutes from an optimum acquisition was due partly to a hardware problem (the SCM monitor-receiver was initially saturated by high signal strength) and partly to a procedural problem at DSS-51 precipitated by the same hardware problem.
4. 4. 2
SPACECRAFT
PERFORMANCE
At approximately L + 25M, the spacecraft became visible to DSS-51, and the initial DSIF acquisition procedure for establishing the communication and tracking link between the spacecraft and the ground station was initiated. The first ground-controlled sequence (initial spacecraft operations) was initiated at L + 45M, and consisted of commands for turning off equipn_ent required only for the launch-to-DSIF acquisition phase (e. g. , transmitter high-power off, accelerometer amplifiers off, etc. ) for seating the solar panel and roll axis locking pins securely, for increasing the telemetry bit rate to lI00 bits/sec, and for initially interrogating all telemetry commutator modes. All spacecraft responses to commands were normal. As a result of data assessment, it was determined that the star intensity telemetry signal was indicating that an object (thought to be the earth) was in the Canopus sensor field of view. Therefore, it was recommended that the roll axis be held in the inertial mode and that 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 indicated not be turned on, that this receiver since the receiver A automatic frequency was tracking the ground station signal. control
4.4-1
4. 5 COAST PHASE I {INCLUDING CANOPUS ACQUISITION}
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 were obtained by use of transponder B and transmitter B operating in low power. At L + 4H33M, control of the spacecraft was transferred to DSS-72 to provide additional tracking data. This transfer necessitated a decrease in telemetry data rate from ii00 bits/sec to 137. 5 bits/sec due to the lower antenna gain available at DSS-72. At L + 5HZ3M, spacecraft control was returned to DSS-51 and, at L + 5H30M, the telemetry data rate was increased again to If00 bits/sec. At L + 6H6M, a spacecraft roll maneuver was initiated to make a star map and locate Canopus. Per real-time recommendations, the maneuver was begun with omnidirectional antenna B and transmitter B in high power (transponder off}. Mode 5 data were available at if00 bits/sec. Two complete revolutions were made to generate the star map, the first with antenna B and the second with A. The earth, moon, and stars Shaula, Rasalhague, Menkalinan, and Theta Ophiuchi were identified. Canopus was located after 237 degrees of roll. As was the case in the SC-I mission, a Canopus lockon signal was not generated as the star sensor swept past Canopus, since the Canopus intensity signal was above the lockon range upper threshold. As the vehicle continued to roll, the time for sending the proper command to achieve manual lockon to Canopus was computed, with manual lockon being achieved at approximately E + 6H38M. Roll attitude was now precisely determined, a prerequisite for the premidcourse maneuvers. Following this successful lockon, a gyro drift check was initiated (L + 6H54M). 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 + 9H3M, the check was terminated. With the DSS-51 visibility period ending at L + 9H46M, a potential gap existed in the coverage since DSS-II would not yet have visibility. Because DSS-72 had visibility of the spacecraft for part of the gap, spacecraft control was transferred to DSS-72, requiring telemetry bit rate reduction to 17. 2 bits/sec. Unfortunately, DSS-72 had considerable difficulty in
4.5-1
providing good data; it was estimated that 80 percent of the data was bad. At L + 10HIZM, the spacecraft became visible to DSS-11, and two-way lock was achieved by this station at L + 10H35M. The bit rate was increased to ii00 bits/sec at L + 10H40M. Because analysis of the spacecraft receiver B automatic gain control telemetry data obtained during star verification and acquisition indicated a signal strength which was approximately 18 db below the predicted value0 a special test for performing an in-flight calibration of this data channel was recommended. This test was required to establish whether transponder operation two-way tracking could be used during the midcourse correction, since a degradation of 16 db in receiver B sensitivity (i. e. , a receiver malfunction) might cause loss of two-way lock during midcourse. Following satisfactory completion of the scheduled premidcourse low power engineering interrogation, the special calibration test was initiated at L + 13H6M. During this sequence, DSS-II transmitter power was reduced in Z-db steps until the command threshold level (as indicated by an indexing of the receiverdecoder-select unit) was reached. This occurred after a total reduction of Z4 db at a telemetry-indicated signal strength of -133 dbm for receiver B and -121 dbm for receiver A. It was concluded that receiver B calibration had changed, but that the signal strength could be lowered by Z4 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. Also recommended for midcourse was the roll-yaw maneuver pair (plus roll of 75. 3 degrees, followed by a plus yaw of ll0. 5 degrees), primarily from an analysis of the telecommunication performance expected for each of the four maneuver-pair candidates (i. e. , roll-yaw, roll-pitch, yawpitch, and pitch-yaw). At L + 14H27M, the scheduled premidcourse engineering interrogation was inititated. This sequence was executed using low power transmitter operation, since a data-rate of Ii00 bits/sec was still available. As part of this sequence, the gyro speeds were measured and were reading nominal values (i.e., 50 cps).
4.5-2
4. 6 MIDCOURSE CORRECTION
4. 6. I
MIDCOURSE MANEUVER ANALYSIS
A midcourse correction of 9. 587 m/sec was computed to soft land Surveyor II at a desired site, +0. 55 degree latitude and +359. 17 degrees longitude, on the lunar surface. This correction was executed upon ground command at 05:00 GMT on 21 September. Due to hardware failures, the midcourse correction was unsuccessful, and there was no soft landing. Proximity of the uncorrected and the original aiming point is shown in Figure 4. 6-I. The uncorrected, unbraked impact point is located on the western edge of Sinus Medii just northeast of the crater Mosting. The selenographic coordinates of this point are approximately -0. 0837 degree latitude and 354. 658 degrees longitude. The targeted aiming point was 0. 0 degree latitude and 359. 33 degrees longitude. The two points are approximately 142 kilometers (88 miles) apart on the moon's surface. Also shown in Figure 4. 6-1 is the approximate final impact site of the spacecraft. Figure 4. 6-2 shows the prelaunch target site, the in-flight aim point, and the associated dispersions. The 99-percent dispersions are shown as an ellipse on the surface with a semimajor axis of 53. 9 kilometers (1. 77 degrees), a semiminor axis of 17. 17 kilometers (0. 56 degree), and an orientation angle of -57. I degrees (Figure 4. 6-2_. In order to maximize the probability of soft landing, the aim point was biased from the original target value of 0. 0 degree latitude and 359. 33 degrees longitude. The biasing was based on a detailed examination of Lunar Orbiter photographs. The maximum midcourse correction capability, as a function of the unbraked impact speed, is shown in Figure 4. 6-3. The expected 3_ Centaur injection guidance dispersions and the effective lunar radius are also shown. The midcourse capability contours are in the conventional R-S-T coordinate system. The maneuver execution time of 16. 2795 hours after injection was chosen. This time allowed 6 hours and 17 minutes of premidcourse and l hour and II minutes of postmidcourse visibility from the Goldstone tracking facility. Nominally, the midcourse time was 14. 5295 hours after injection, but was delayed l hour and 45 minutes because of operational difficulties.
4.6-I
80 7O 6O
80 7O 60
5O
50
4O
4O
20
20
I0
I0
WO
OE
I0
I0
2O
20
3O
40
4O
70 8O 80
70
Figure
4. 6-1.
Surveyor
II Target
and Uncorrected
Impact
Points
4.6-2
\ 9
_F
FINAL
AIM
P_NT
Figure
4. 6-Z.
Surveyor
II Impact
Locations
....
4.6-3
-12 VIM P ,- k m/sec
(3" 03 Oo
--I0
I l',a I F-, Oo
-8
-6
(,o rr bJ I-bJ 0 u ,i
-4
--2
I
I
o
z (/) o "r t2 |
I
o
tY 4
10 TARGETED UNCORRECTED AIMING 3cr 30" 14 -16 POINT VELOCITY TIME -12 ERROR -I0 SITE = LAT MISS-'LAT :LAT -" 0.00 = 0.084 = 0.55 °, LONG : 359.33°E _'wD _ --
° S, LONG : 354.66°E ° N, LONG:359.17°E -LAUNCH: 20 SEPTEMBER
II
1966 i E L = 115 DEG MIDCOURSE AT i i 8 I0 15 I/2 i 12 hrs 14 2 4 6
IMPACT FLIGHT -14
ERROR = 423m
:9.39m/sec
-8
-6
-4 TQ,
-2 THOUSANDS
0 OF
KILOMETERS
Figure
4. 6-3.
Midcourse
Capability
Contours
4.6-4
The
predicted
results
of
the
selected
_llidcourse
correction
an_t
oIher
alternatives considered are given in Table 4. 6-1. The required velocity conlponent in the critical plane, to correct r_liss only, was 1. 185 resee. The noncritical direction component that resulted from a weighted selection of flight time, n_ain retro burnout velocity, and vernier propulsion systenl fuel margin was 9. 5 m/see. Figure 4o 6-4 shows the possible flight tinws, burnout velocities, and fuel margins for the range of available noncritical component velocity corrections. Since all three were acceptable over a wide range of values, a non_inal burnout velocity of 450 fps was chosen. This gave favorable landing site errors and backup midcourse correction capability in the event the first midcourse correction b(_can_e nonstandard° If the n_aneuver noncritical approximately strategy component 4. 48 were would m/see. to correct have been miss 4. 325 plus flight time, resee, giving the required a total of
Since the air_ point was changed during the flight, correction does not properly evaluate the performance of ance system. Using the results of the last premidcourse ing to the Niiss plus 4. 6. 1. 1 original flight Alternate the and No aim point ti_ne was 4. Considerations premidcourse elinlinated: correction. phase, the following gives a miss 44 m/sec. only require_lent
the above the Centaur orbit and of 1. 015
required guidcorrectm/sec.
were
During analyzed
alternate
possibilities
1)
lnidcourse
This
case
would
have
r¢,sult(_d
in
acceptable values, 359. 17 uncorrected long5tt_dc 2) Injection
burnout velocity, but since a landing degrees longitude site (Iri_urc plus 14.
fuel margin, and site of +0. 55 degree was desired, it was degree latitude
arrival latitude eliminated. and 354.
time and The 66 degrees
was -0. 084 4. 6 2). 5 hour
corrections. velocity correction because the higher than of nominal 2 m/see burnout or 505 was fps,
a)
A minitnum considered velocity
midcourse and eliminated would have been
desired,
b)
Two midcourse velocity corrections that would have resulted in a burnout velocity of 400 fps were considered. The first one, at _15 m/see, was elin_inated because nlission success could be achieved with the selected n_ancuver without tion gives presents was errors requiring greater problems. lower as large a correcti_n. if the possibility, =\ smaller first correc backup The because fuel capability second correction -33 n_/sec, landing site
eli_inated and
of significantly margin.
greater
4.6-5
0 ,,D o0 00
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4.6-6
_v
68189-2-185
O O 0
O0 ! b_ ! Oo
_0
t_ ,r.4
E
O
Z © DLTMC
DI,TMI
TA
+ Total
Maneuver
Time
DLTMC 3
DLTMC 3
+ _ Maneuver
+ Second Time
tD
>
>
_D
0
For
Surveyor
i: TA -- I0 minutes = i0 minutes = I0 minutes = 150. 7 seconds = ZZI. 0 seconds
DLTMI DLTMC First Second maneuver maneuver
Figure
4. 6-6.
Maneuver
Timing
Plan
4.6-10
It is notable that engine burn was begun within 1 second of the nominal ignition time. The earliest first maneuver (GMT) or earliest allowable break of sun lock is shown to be nominal ignition less total maneuver time, less TA (the operational time necessary to transmit and verify spacecraft commands). By previous agreement, a value of 10 minutes was used. The last times of the first and second maneuvers are computed based upon DLTMI, DLTMC, TA, and the maneuver times. These two times serve as guides to proper execution of spacecraft rotations. The the operational maneuve r. resulting data midcourse necessary message, for properly as shown executing in Figure 4. the midcourse 6-7, contains
4.
6.
2
SPACECRAFT
PERFORMANCE
The midcourse correction sequence was initiated at L + 15H42M with the engineering interrogation, which indicated that the spacecraft \_as ready for midcourse operations. At L + 16HI1M53S, the first attitude maneuver (plus roll of 75. 3 degrees) was executed and confirmed as being satisfactory. At L + 16HI6MSS, the second attitude maneuver {plus yaw of 110. 5 degrees) was executed satisfactorily, thereby aligning the spacecraft in the desired direction for applying the midcourse thrust. FolIowing + 16HZ1M) and the the pressurization loading of of the desired vernier thrust propulsion time in syste_ flight-control
(L
the
the
programmer magnitude register (L + 16H23M), thrusting of the vernier engines was commanded at L + 16HZ8M. At this time, vernier engine 3 strain gage 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). After the previously commanded vernier engine thrust duration of 9. 8 seconds, the engines shut off. However, DSIF receiver automatic gain control showed that the vehicle was rotating at a rate of approximately 1. 22 rps, with a secondary motion having a period of approximately 12 seconds. In an attempt to stabilize the spacecraft, the flight control subsystem was commanded to the rate mode. Approximately 10 minutes later, when it became evident that the gas jets were not going to stop the spinning {since approximately 60 percent of the gas had been used, and the spin rate was still 0. 97 rps), the gas jets were' inhibited. The remaining gas supply was thus conserved for use in the event that the malfunction could be cleared and the vehicle stabilized by vernier engine firing.
4.6-11
68189-2-187A
6_
_q
(D OZ_ZZ_ ZZ
0
m<<<<_
DDDDD_oZ
<<
D_
(D cO ZZ mm 0 'O m
_D -/D O O_ O
m , _m >
m
O< mb_
I
,..Q
&
,i:iii_ : ...........
' ;!i-'iK! :ntm.tf_
;ii!
.........ii:i':
........ i-I
....
_;ii:tiJ!t _I
! !fHk._it_t:_ _
Figure
4.7-18.
Sun
Clock
Angle
for
(I, I)2
Thrust
Condition
4.7-18
I
I
] I
IIBUIH |Ng'FRUII_ ENTB
1
Figure 4.7-19. DSIF Receiver Automatic (High Magnification) time: 264:05:00:03. Gain Control Signal
Reference
750
Figure
4.7-20.
DSIF Reference
Receiver time:
Automatic 264:05:00:03.
Gain 750
Control
Signal
4.7-19_
]
ii
j--j+
r--L;;
• ; : : : : : ::i
.L_L. ,.i _ Li_Li
Ill -...... iI
' i[
I,
i +--
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'4J
I ....
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t
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[ I I I I 1 ] illJ I I lt-_ f I I [] I I t_ I"L_-I_ r-]_
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t-L
I I t I I I tt.Llll+ I I'1 I I I l,l
-il_J _It
4
; T
i
i
i
i
÷
-+
-_t ++_ +_
: : : : :
.... + ,J..i;i...i..,
_!_ q_ ....
7T_t-t
I_,
+ IIII.. ,,L.I...I _ IlI+ .......... : :
i
-L
_++
.__
: : : ++_+.._+__
I I J
AJ±[/_ !_ [F1-
I
!T+
t
r!
i-]_lq
:L
I[
!t!tt
: : i i i
-N LLJ;,,_t[
1-
iiiii_ !-!! ! !
Iltli
+
H
I1111
: : i i : : : : : :
.F
!!!!!
i i I h,Lt.
r+ rt+Y
ilhrl -+--t
F-+<-_,
! ! .+Ct-.!! t-r_-_
: -,: : : : : :- : : :
iiiii_
I tii :,_
:
:
:
:
:
TI_t.-I--FI L..I.. ill..l...
._.
-T-!-_-! ! -lMi-r! F-_ ÷-+
iiiii iiiii iJiii
!!l.i!
iliii:
iiiii
I I i IT+
!iiii
J ; : :
qT_
l'Oz_o_
l_,_
B
, f..... "_-":':"
' "
- .
'_ ",, _:-. ........
",=QT F!LMED.
to verify the as a function
calculated of assumed
spacecraft thrusts.
motion During
output from the digital simulation the short interval studied, the
only RF link parameter that should have changed appreciably was omnidirectional antenna gain. This gain variation is a direct result of look-angle (from spacecraft to DSIF) spacecraft transmitting antenna gain is shown in Figure 4. 7-21. The transformed for for spacecraft dynamic into a track of the changes contours due to tumbling; a map of the in spacecraft coordinates
motion from earth vector
the mathematical model in spacecraft coordinates map and
was suitable
superposition the look-angle
on the omnidirectional antenna gain trajectory for final selected thrust
(see Figure 4. 7-22 moments combina-
tion). The intersection of these two functions produces a gain profile in time, which is directly comparable to station automatic gain control variations. The correlation between salient features of the simulated gain profile and the DSIF AGC data was the criterion used to evaluate the various thrust level and moment model matrix was not combinations chosen. Any considered in this analysis. inaccuracy in the simulation
The initial part of this study was mainly concerned with matching the two M-shaped waveforms which occur near times 4 and 5. 5 seconds (reference time is 264:05:00:03. 750) in Figure 4. 7-19. These waveforms can best be reproduced when the earth vector traverses the -6 db contour (located at coordinates @= 84 degrees, ¢= l0 degrees) in such a way that ¢is relatively constant. Many of the thrust level combinations proposed could be immediately discarded since it was obvious that they could not produce the correct waveforms at the proper times. Based on the preliminary study and analysis of sun sensor and gyro data in this 20-second time period, it was decided to investigate in detail certain cases which were still considered reasonable. These included modified thrust levels for the thrust level for the third For where simplicity, the two engines engine, and used that did fire, a non-zero (but small) a modified moment of inertia matrix. are referred to as (TI, T2, T3)M ,
combinations
T l T 2 T 3 M
= ratio = ratio
of assumed of assumed engine moments
and and
telemetered telemetered in pounds matrix
thrust thrust
levels levels
for for
engine engine
I Z
= assumed = modified
3 thrust
of inertia
4.7-21
120
Figure
4.7-ZI.
Omnidirectional
Antenna
B Gain
Contours,
Z295
MHz
4.7-22
,!iI !i!
IUH _" !V,:
i'ii_i s
a)
From
0 to 6 seconds
1!!!!!11! l_i_ k
if!' !!_i!!_ti _H! !il:t!J !i :l_ii!_1]iiii
ri:[ ::HII:I:
i
ili! i_iit[ii_ T T:I
• i:; H
_11 I
i .... i ....
!1 ,:
_:;¢ !fif_fili
r_t
Utt :i;I
i!:_ iii _,;iil ;_:I ii! iii[ ;1:1
:H ::::
ii!i _i!it!i!i
[I; i:H!t;]
¢i;
i:_:
"'
i !ih
"!i!:!! : !ii!: i
.... lil
!!i!
II ;:]: I i:V
!i ....
i:; :IH
;!
i1_;4
Tilt I_THIT!!
........ ::: ::¢: _ ii H.
!
ii:
i;[: i'[
, •
_NT_
PHI ' /-i
_5
-'
b) Figure 4.7-22. Look
From
6 to 9 seconds for {I, 1,2} m Thrust Condition
Angle
Trajectory time:
Reference
264:05:00:3.7
4.7-23
! ,_
ii_! ii!i
!i !i
F:
:1:
:i
c)
From
9 to
II
seconds
ti]It!ili
,._ l-
....iiiii:i_
tt:q:;:t "* ....
!1 P,tt : t': I:'I:!:t!
:ff::
d) 4.7-2Z(continued).
From
11
to
13
seconds Trajectory for (I, 1,2) m Thrust
Figure
Look Angle Gondition time:
Reference
264:05:00:3.7
4.7 -24
e)
From
13
to
15
seconds
f)
From
15
to
18
seconds
Figure
4.7-22(continued).
Look Angle Condition time
Trajectory
for
(l, 1,2) m
Thrust
Reference
: 264:05:00:3.7
4.7-25
!i
g)
From
18 to Z0
seconds
Figure
4.?-2Z(continued).
Look Angle Condition time:
Trajectory
for (I, I, 2)m
Thrust
Reference
264:05:00:3.7
4.7-26
In this way, the six cases investigated
(0. 9, O. 9, 0), (1, O. 9, 0), (1, 1, 0),
in detail can be called (0. 9, I, 0),
(1, 1, 2), and (1, 1, 2) M.
Comparison of the gain and automatic gain control waveshapes (see Figures 4. 7-23 through 4. 7-28) for these six cases resulted in the conclusion that (i, i, 2) M produced the best fit with station automatic gain control. Table 4. 7-3 summarizes the considerations that led to this choice. Two of the cases were immediately rejected, since the spacecraft motion did not have the correct period, producing a time displacement between the gain and automatic gain control curves that was quite obvious after the fifteenth revolution. A choice among the remaining four cases was much more subjective, involving comparison of subtle features of the waveshapes. It should be stressed that the antenna gain tolerances are large (2 db for high gains and 8 db for low gains), and that small changes in look-angle could appreciably alter the waveshape details. _':-" Also, the antenna gain waveshapes had to be "mentally smoothed" during the comparison to compensate for the low-pass filtering used in processing the DSIF automatic gain control oscillographs. Therefore, none of the four cases with the proper period should be considered unequivocally eliminated.
4. 7. 7
SIMULATION
OF
GYRO
CROSSOVER
PROFILE
Using
the
standard
moment
of inertia
matrix
discussed
previously,
the simulated gyro outputs disagree significantly with the flight data for all thrust factors selected which provide a reasonable fit to the sun sensor and automatic gain control flight data. A plot of gyro output telemetry to the same scales used for simulation outputs is given in Figure 4.7-29. A typical example of the simulated gyro output for the standard inertia matrix, [ 0, 0, 0, 0, 0, 0], is given in Figure 4.7-30 with flight data points plotted for comparison. This [I, i]2 thrust case is typical of the set of cases used to fit the sun sensor and automatic gain control data ([0.9,1]0, [I,0.910, [0.9, 0.910) in that the simulated pitch gyro crossover does not occur, the simulated yaw gyro crossover occurs early, and the duration of the negative saturation interval is approximately 2 seconds less than that of the flight data. In order to improve the correlation between gyro crossover, sun sensor, and automatic gain control flight data, percentage variations were made in the initial moment of inertia matrix elements. Although an exact match to the flight data was not obtained, gyro simulated output is presented in Figures 4.7-31 through 4.7-35 in which best match cases obtained are presented. The sun sensor Figures 4. 7-36 and 4.7-37. comparison plots for these cases is given in
*On some Estimated
curves, values
I0 db errors occasionally occur on some have been drawn in for these instances.
of the peaks.
4.7-27
TABLE
4. 7-3. SUMMARY COMPARISON FOR
OF AUTOMATIC THRUST LEVEL Comments
GAIN CONTROL/GAIN COMBINATIONS
Case
(1,
l,
Z) M
Period Peaks
correct. for cycles
Dips in cycles 9 and i0 Ii through 14 correct. ll waveshape too
correct.
(1,
(z, (i,
l,
z,
z)
o)
Period correct. Cycle Peak of cycle 14 late. Period Period instead correct. correct. of two has has wrong wrong
smooth.
Waveshape
of cycle
l0 too
pointed. three
o. 9, O)
Waveshape of first cycle has peaks, but otherwise correct. period. period.
(0. (0.9,
9,
1, 0.9,
O)
Signal
o)
Signal
4. 7. 8
SIMULATION
OF
INITIAL
GYRO
RESPONSE
In an attempt to match the pitch and yaw gyro angle response the first second of the midcourse period, perturbations were made following : Ignition time in ignition level time between engines
during inthe
Difference Engine Moments The moment of inertia
thrust
of inertia changes which were required on the to bracket simulated the gyro gyro outputs
crossover response (Table 4. 7-4).
produced
negligible
effect
A plot of gyro outputs (FC-16 and FC-17) and simulated gyro outputs for the (l, I)Z thrust level condition is plotted in Figure 4.7-40. The effect of the uncertainty in ignition time (see subsection 4.7.9) is indicated in the maximum and minimum values plotted. A close fit to the pitch gyro data could be obtained by appropriately selecting the ignition time; however, a fit to the yaw gyro data could not be obtained. It should be noted that due to the saturation characteristics of the gyro telemetry output, the curve shape is not representative than 4 degrees. of the actual flight motion for gyro angles greater
4.7-28
J
1
lltftittltlti:tti_,_tltiiitll!ttftfttttttltl I iit tit tlt tt_ttt_tH_ttfbtlt_t:tttiiH_tttt!t-tfttt ti[[lt/ttllttitt I ttt_ttttffftfttlI_T_]_l!!N_ !H tt II t!tll!_IfHt t fftt tltttttt:N_tf_Hl_ t!!l!tI!![l!!ttt_ tf t _ _Ht!Itft;!_ ' I itt !'_ : " ' :}_!_[fttlttfi4-tfftHfrttTHfH_}-fi]iH_t__ !
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]
]
Figure 4.7-23. Antenna Gain Versus Time for (i, 1,2) m
J
I
I:
i
t
r_ _' _i
Figure
_'r' 1
4.7-Z4.
Antenna
Gain
Versus
Time
for
(I, 1,2)
__
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i_',Ii i-l-Filllltlll
IIBECI
1
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Iit!_tI¢tt t
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pR_CE'DING
PAGE BLANK.
NOT
FILME'D.
I
"_: i ii_ t!i!if/lll if! !
:|:I
0
Figure
4.7-25.
Antenna
Gain
Versus
Time
for {i, l, 0)
r:T
i'
Figure
4.7-26.
Antenna
Gain
Versus
Time
for (l, 0.9, 0)
'__
_
I
4.7
- 31
!!!!!!
iflI_
ilti
LU_
Lt,'-]- "St ?._
'If
¥o_
_
3
%
f i
] J ] ]
l_'ig,_re 4.7-_,7. A_,I 'r,l_;z Gain Versus £i_:nc £o_ {i,},. O: [, O, )
1
!
i[l!l
!i L
i
'i
_
,_:i
j
Figure
4. 7--?_8,
Antenna
Gain
Vc;]s,_s
rinse,
fo2L (0_9_
O, 9,
0)
TIM_
15EC)
!
,'2 t_O,U:__ }?_',UI
3
.._..,
r_
t-,-
L_
r_
°r-i v
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I
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4.7-35
Figure
4.7-30.
Simulated
Gyro
Outputs
for
Standa,d
Inertia
Matrix
;I
ii !i
Figure
4.7-31.
Simulated
Gyro
Outputs,
Gase
I
4.7-36
9 IIME[SEC]
io
Figure
4.7-3Z.
Simulated
Gyro
Outputs,
Case
2
TJlMEISEC] l°
Figure
4.7-33.
Simulated 4.7-37
Gyro
Outputs,
Case
3
Figure 4.7-34.
Simulated Gyro Outputs, Case 4
if:
_+÷"*_+!:
i:!:!'!!:'!
!!"
.......
ii
I.........................
7.
i
"
i'[ ........
_.........
ii
i:
'!
_ii
'
:iil:'_II:rl_i!l
:!:,
! .:r_!i_!
Figure
4.7-35.
Simulated 4.7-38
Gyro
Outputs,
Case
5
........ I_I:tI_tHHI!
.......... _+ _+t.
l+k_i_ _
_._',_, _ _i
:::17 _4__.
+* t
I!I 1III-F-_I±I fl_ ....
t itLi-Lit ¢+L i', t[ tI#-L :_
L_E
..........
i:i i i' i_ :_!!
I:::::_i[Iftl[[
[[_:[ll ,t r trltt+_t ............... "ttttt
!!!!!!!!! !:_::!::i: iii_i!_!7 :i1!T_i!ii11!1
:::::iiiiiiitH ..+,+'rr_r _***_!!!!!!!!! H!!!
ttll_tH]_tltH _]]_]
Figure
4.7-36.
Solar
Panel
Sun Transits
Versus
Time
f;!f_
i
_i!: = :
Ihrr
t¢_t t _,_I_ _,__
tt_ tm_
i;iiii
ZZ;.....'!i
iittt! .........
!li_t! ::::::
II!i!1
iJ#fTI77i
'r!'!! ,_,iiill :. '.:::I
!it, tttH [':i_i_
_
h _'
l#¢illlt i "_t_, 0"4'.lii _1_,i-_
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,,
il!! ri
,i i
tti+414 {iiHHi!t
lill]l _++f,+t_÷_
_ r_ ,_
!!!!!!
:: ;11 ..... ; : .
::;,,lltll
...., i ..... .....
IfiIill
7!_tt 111
.....
iiiiiiiiii
4.7-37. Entry Solar Panel Sun and Exit Azimuths
Figure Transits
4.7-39
r--'_
f--------_ o o o o o o
Z
I--I
co
t'-.-
,D
0" ,_
Oq o0
tlq e,,l
0
i__._j
r---_
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I
#
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I
k_---g
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o o
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n
oo
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r,4
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oo
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o
r-,._
I I
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b
Z
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cq
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a
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0o
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u3
.
I
#
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o
I
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o
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oo
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o
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I
_i_
O0
o I _-
o Lr_ L_
o
o',
_
_o
o
-;
t.........._ _......-_
I
#
I
._
I
.4
I
_
o
_
t-.-
_
m
_q
#
ca
ol _D
4.7-40
The effect of an engine 2 ignition delay with respect to engine 1 is indicated in l_igure 4. 7-41. ]Engine 1 ignition time was selected to provide a best fit to the pitch gyro output. From this figure, it appears that an ignition delay of engine 2, with respect to engine l, of 50 milliseconds would be required to match the yaw gyro data. In order to obtain a better fit to the curve shape of the yaw gyro, the factors were varied in addition to the above. The best fit case is
thrust
indicated in Figure 4. 7-42. If ignition delay between engines 2 and 1 is not included, the ratio between engines l and 2 thrust factors to obtain a good fit to the yaw gyro data must be increased to nearly 1. 2.
4. 7. 9
MIDCOURSE
THRUST
DURATION
To the events telemetry
establish time intervals within which telemetry data indicate that of vernier ignition and shutdown must have occurred, all signals correlated to these events were examined. Acceleration
error signal, pitch and yaw gyro errors, roll precession command, engine thrust commands, and engine strain gages are affected by ignition and shutdown. However, the gyro error signals do not show a sufficiently large change at ignition and are saturated at the expected time of shutdown. Also, midcourse was conducted in a constant acceleration mode, causing an acceleration error signal to appear as soon as the ignition command was received. The acceleration error, in turn, produced thrust commands; yet none of these events specifically required the vernier engines to be actually firing. Therefore, the signals employed to accurately establish the events of ignition and of shutdown were the strain gage channels, P-18, P-19, and P-20. Since vernier engine 3 strain gage gave no evidence of thrust from that engine, only P-18 and P-19 could be used. The I) following tO' actual specific time times are defined for the thrust phase:
at which
vernier
engines thrust
ignited occurred and thrusts
2)
t I, time when b_gan decaying t2, time when
end of operational toward zero thrust had decayed
3)
to zero
4.7-41
.......
_, _.._;:, _,,!"-;T i"l!,'_¸;
:T'T!:_
i !t:q
i ...... ,
4:: i _'t' ,i:!:i! !.
-11
i
]
' i
k
_
i :.' ;,}' !-~--t::
lIT
....
_' k, i:li;
i i ;i , ,'P_
i
, %
• ...... ;;Li .,, .... , !!ii!:i11:_!
, ,
i!!:_;i:
; :i it
# :,!
i:i i i i!:::
:!i.
H' : H
i
lill
Figure 4.7-40.
7T-
:}':
:, li !! i!
'
]
11
',??i:i-f x:_a Angles ir_First Second 0. 525-second Delay After Midcourse
Pitch and YawGyro For (I, I)2 With
i
i
r7
i,_oi
_""r-r
L
I i'll !
Figure
4.7-41.
Pitch and YawGyro Angles in First Second For (i, i)0 for 0.5-second Delay
After
Midcourse
4.7-43
Figure 4.7-42.
Pitch and Yaw Gyro Angles in First Second After Midcourse For (0.9, 0.79)0 With 0. 525-second Delay
4.7 -44
The occurrence of a given event can be restricted by a single telemetry signal to the interval between successive samples. When telemetry signals are available, this interval can be further restricted, seen in Figure 4. 7-43, by combined estimates. This procedure has used to determine the maximum and minimum values of the following Figure 4. 7-44: Atol = operational to start vernier thrust engine decay at thrust, t1 from ignition at to
two as been in
of
Ato2
=
engine thrust
thrust at t2
period,
from
zero
thrust
at to to zero
Atl2
=
engine thrust
decay at t 2
time,
from
start
of
decay
at
t 1 to
zero
4.
7.
10
MOMENT Two
OF
INERTIA sources
UNCERTAINTY introduce errors weight and, in spacecraft in the moment of inertia
independent
matrix. uncertainties uncertainty source from is the
The
first source includes prelaunch of each spacecraft component (mounting error) relative to the of the roll vernier axis (Z) engine due to
and center of gravity addition, positional frame. The second center velocity of gravity about the away _ axis.
the shift spacecraft
propellant an angular
4. 7. I0. i
Prelaunch
Uncertainties
Values of the moments of inertia and the products of inertia are calculated by the Mission Mass Properties Profile computer program. The accuracy of the computed values for the moments of inertia was demonstrated on SC-2 when actual measurements (accurate to approximately ±l percent) of these moment of inertia matrix elements were performed. For those tests, the computed values were shown to agree the measured moments of inertia within +2. 5 percent for the dry landed weight configuration. No actual measurements of the spacecraft products inertia have been made to determine the accuracy of the mass properties calculations. The program calculates moments of inertia and products from specification drawings or The inertia matrix is determined origin, then axes are relocated gravity, which is also computed. gravity program of weight location are and
with of
the entries to the moment of inertia matrixof inertia-using weights and locations actual measurements for all spacecraft units. with respect to the spacecraft coordinate by the parallel axis theorem at the center of When the actual weight and center of the inputs measured to the values
in the X-Y plane are measured at AFETR, "corrected" to ensure that the computed and center of gravity position agree.
4.7-45
i
Figure 4.7-43. Time Interval Which Vernier Ignition Could Occur
in
Figure
4.7-44.
Possible
Time
for Vernier
Engine
Thrust
4.7 -46
Analysis of this correction procedure has shown that, for the magnitude of correction in weight and center of gravity location required for SC-2, the products of inertia determined in subsequent computations could be in error by as much as 5 to 6 percent. The spacecraft center of gravity vertical location is only known to within 0. 25 inch and is not measured at AFETR when the spacecraft is in its final configuration. Transformation from the spacecraft coordinate origin to the center of gravity as the reference point for the ine'rtia matrix introduces an error in Ixx and Iyy on the order of l percent due to uncertainty in the Z coordinate of the center of gravity.
This discussion is not intended to specify the inaccuracies involved out that they accuracy of the can not be However, test Z to 4 in the moment of inertia matrix entries, but rather to point exist and give a general indication of their magnitude. The products of inertia calculated by the mass properties program specified since there are no measurements for comparison. data indicate that the computed moments of inertia are within percent of the true spacecraft values.
4.
7.
10.
2 As
In-flight the
Uncertainties spins about the roll axis (_), the resultant
spacecraft
centrifugal force causes the vernier engine propellant to shift away from the Z axis (see Figure 4. 7-45). This produces a change in the moment of inertia of each of the six propellant tanks about the Z axis, which, in turn, change the spacecraft moment of inertia (Izz). To compute the maximum change in Izz, assume the^propellant in each tank moves such that its surface is parallel to the Z axis. Figure 4. 7-45 shows how the liquid moves from its prelaunch condition to the position in a zero g field under the influence of a centrifugal force. For ullage of 48.7 in 3 in each tank, the center loading tanks of gravity of 182.4 of 0. 6 slug shift is 0. 28 inch. pounds and produces ft 2 or 0. 3 percent. This shift a change corresponds in Izz from to all a propellant six propellant
4. 7. II CLOSED-LOOP Spacecraft behavior with a closed-loop mixed electronics hardware and dynamics (Figure 4. 7-46). approach is the accurate
ANALOG
SIMULATION
RESULTS has been investigated flight control of vehicle mixed simulation saturation charac-
during midcourse thrusting simulation involving flight-type an analog computer mechanization The major advantage of the representation of all electronic
teristics. Previous simulation studies, involving only analog computer equipment, were not successful at reproducing the thrust command profile during the 1-second period following midcourse ignition. This was due mainly to the difficulty of simulating the sXturation characteristics which have a strong influence on actual vernier engine thrust levels.
4.7-47
7
LIQUID
PRELAUNCN
CONFIGURATION
CENTRIFUGAL
FORCE
CONDITION
Figure
4.7-45.
Effect
of Rotation
on
Eiquids
in Fuel
Tanks
O _ 00
To
I ELECTRONICS I ACCELERATION amPLIFiEr T, _ I _ lJ q _ _ Lr_L,_m_u_LQIr_C,I_IE NO, I THRUST L_._x_ _ [ EQUATION l MIDCO URSE SHORTING SWITCHES I _ I EF3MPLITER ,_'LCHANI_ATION AND ROLL LOOF j_ ' Ft_ 1 T O
I
COM AN .OWN
_z
-II-
L__I_
PITCH
GYRO
TELEMETR_
ROLL GYRO TEL_M
Figure
4.7-46.
Analog
Computer
Mechanization
4.7-48
Also included in the mixed simulation (and not Considered previously) were the effects of pitch and yaw gyro errors at midcourse ignition. Initial errors are possible up to ±0.4 degree per axis due to limit cycle deadband and electronic offset, and have a strong influence on initial vernier thrust levels. The assumed conditions for engine thrust response were zero thrust from engine 3 and normal thrust from engines l and 2. By selecting realistic initial gyro errors and engine ignition delays, a good reproduction of initial pitch and yaw gyro error telemetry signals was obtained (Figure 4. 7-47). Also, simulated vernier engine thrust behavior showed qualitative agreement with strain gage telemetry signal waveforms. These results demonstrate that the initial pitch and yaw motions observed at midcourse ignition are reasonable responses under the conditions listed above. The mixed simulation mechanization is shown in Figure 4. 7-46 with the flight electronics contained in the dashed rectangle. All other equipment used is part of the analog computer. The engine throttle response, gyros, and telemetry lags were mechanized on the analog computer along with the spacecraft rigid body dynamics. The roll loop was mechanized entirely on the computer. Relay switching was set up to start analog computation when midcourse was commanded from the electronics. Vernier engine thrust was delayed from this command to simulate the ignition delays. The computer simulation was run for the following Gyro saturation, degrees condition:
Pitch 15. 5 Yaw 21.6 Roll 16.8 Engine I 0.050 Engine 2 0. I00 0.0 0. 26
Engine ignition delays, seconds Initial conditions, volts Pitch demodulator output Yaw demodulator output
A telemetry recording of the simulation data points superimposed
(0.4
degree) with
results is given in Figure 4. 7-47 on the vernier engine thrust traces (@x and@y). The computer
(T l and T2) and the pitch and yaw gyro errors output matches the flight data reasonably well. 4. 7. 12 SPACECRAFT To describe the MOTION spacecraft motion
for 20
seconds
after
the
midcourse
correction was attempted (t = 0 sec), the digital spin program was used to generate time dependent curves for spacecraft axes motions, angular velocity, angular momentum, and sun vectors in both inertial and body coordinates. All of these quantities have been plotted for the same thrust level/moment of inertia combination of (I, l) 3; (0, 0, Z, -10, -g0, 0).
4.7-49
IO0
STRAIN FCO rr.o IGAGE SIGNAL,P--IB
1
75
_. _,_. _":,__
,_A_
.
_"_"FLIGIT (THRUST CONTROL, COMMAND} FC 25
,,D
! I
50
-I--
,?,
25 =e 0
I00
I.u')
75
.
_
o :_'-Z-c_ _
F- . _
_J
50 z5
nO OC W
81 F C - 16 ,..._.. 4
.=_
o_
rr_ -r L) -4
0
,,u._
-8
BO
nO nr 40
0"o
O-
,FC-
17
-40 -80
_'_
_,...._,_
2O
nO rr I0
o_.
-I0 J 0 rr -20
0
0
SECONDS
Figure
4.7-47.
Mixed
Simulation,
Midcourse
4.7-50
Figure
4. 7-48
represents
the elements
of the direction into inertial coordinates,
cosine coordinates. for example,
matrix (C%) which To obtain the motion
transforms of the roll
a body fixed vector axis (ZB) in inertial
the direction cosines can be plottedat C^T(I, 3), CT(Z, 3), and CT(3, 3) of the Z B axis relative to the inertial XI, YI, and Z I axes, respectively, on a unit sphere. This presents a visual representation of the ZIB axis for the ?.0-second period of interest. (The motion of the spacecraft pitch (J_B) and yaw (YB) axes are obtained similarly. ) Figure 4.7-49 shows the time dependent behavior of angular velocity and angular momentun_. Both the spin and angular n_onlenturn magnitudes increase almost linearly until vernier engine thrust is terminated (9.8 seconds) and are constant from then until the next thrust period. Figures 4. 7-50 through 4. 7-53 describe the spacecraft angular veloc-
ity vector (W) motion shows the components
in spacecraft and inertial coordinates. of W in inertial coordinates. The X,
Figure 4. 7-50 Y, and Z com-
ponents increase when the vernier engines about values of approximately -350 300, Figure 4. 7-5Z is a graph of the components nates. After thrust has been terminated,
are thrusting and then oscillate and -ZZ deg/sec, respectively. of W expressed in body coordithe components are sinusoidal in
nature. Figure 4. 7-53 gives the clock angle ¢ and the cone angle @ for the direction of W in spacecraft coordinates. Figure 4. 7-54 shows how the direction of the angular momentum changes over the Z0-second period of observation in inertial and body coordinates, respectively. In inertial coordinates, both the cone angle and clock angle are essentially sinusoidal from 4 to 9.8 seconds and, as would be expected, are constant from 9. 8 seconds until the next vernier engine firing. In body coordinates, the clock angle completes one revolution, and the cone angle turns 350 degrees in this same Z0-second period. Figure 4.7-55 gives the clock angle and cone angle for the sun vector expressed in body coordinates. When the cone angle exceeds 90 degrees (roughly), the sun ilhminates the solar panel. Figure 4.7-56 is a time plot of the computed gyro output angles. 4.7. 12 REFERENCES
I.
C. T. Dorsett, "Mass Revision C, Surveyor
Properties SC-2," IDC
Monitoring 2225.05/464,
Report, SSD 64226R, 15 September 1966. Current Telemetry for SC-I
2.
J. H. Green, "Correction and SC-2," IDC Z29Z/3Z, "Surveyor Spacecraft 21 September 1966. J. H. Report
of Throttle-Valve 23 August 1966. Assessment
3.
Monthly
Report,"
SSD
68202R,
4.
Sports, "Secondary Sun Sensor Performance No. SEAR-16, 26 October 1965.
Characteristics,"
4.7-51
!i-i
H _ii-£
_;i
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+-: I::
i ........
-.:-,:_!.:::i_:
_¢!,I
"1"J'l'' I "' "/ ....
,, ._ llli_ :I :',t:t:: _,
%,
_i;tt!t,tM I_
_'r'l'_=l_;iIi; X-':-i--[ ;_h--t-_-' ....
,TT '_I_I-N_--t-T
_ !i ' i _ !;_i
:! ....... jTJ i iT
:T
"
'i ........
t t-t_f-1 -
t ,l]
]:t_!_
5TTTF7 i ', Y
4-:! S _!i--i-',!i.
-_-. .....trut
H--_-_-,, t tf Ii
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I ......... P2I-JP-;IY:
'2 2 ._ .... I
T:
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,/
i(
4
I...
_ J _. , "r] i LO TIP_ibFCI I ;i _: I
;:
I
i x-
L
I
:
J
'
i
i:!
a)
CT
(1,i)
b} Figure
GT
(Z, i}
4.7-48. Direction Cosine Matrix (CrF) Elements for El, 12] 3, [0, 0,2, -lO, -20, 02]
4. 7-52
......iiii_.t_ o ,
7!:
it_
1.11.
ii1
,i, lIl'::_h ihi!__ .::i:i
_'
:i'?I
i: '_' '!
!L
i i li.... _]}_t
i iiF:i:!!_ !iil _F_ '_°'' _
[ "f,:
[! ....
Ii:l
I! !if!::Ii'® t1' I : ;l:ltlttt:
.:_ [','i t I:ti[I " ll:l
::::
....
t':l
fi',_ tiiitlit:t_ii
It:ti_1
;i
_!!i
HI;
Rrl
_7l["
1171
i:
:ttl'
ltii!i: iil '! llli Illi
il: IltI } ili_iiiiiii
:'ill
tl ttit i "
Itt:
;!
111
..
,_-..I I.-,
[::[_lti ,,,,:!tit!, itt
Ht-
7!
-
u:_ illi4 ""
i i,
,
'_"
i
[
:
:,
_
_
"
.....
_Ir'i
;:i:_
lf!itil:'i
iiii It
t:',:
:[[: :1:
:N: t!h " ;r!"
;::: :t:! I _l-r
F_:. :_*ii!iu: t
i l 1 q. _. i. I. ii. ii io. T l#411[51FC) 1 ii. i1. ii.
i._.i_]" _ 'tttt! li it!t!{i! it}! i: ::ti li!i "
ii. i I_. ii, if. ii.
iii iiiiii!!
• i _
c)
Figure 4.7-48(continued). for
cm (s,i)
(Cl')Elements
El,
1-]
Direction Cosine Matrix 3, _'0, O, Z, -10, -20--]
li
li
': .... i!i i!ii
;i:[ _::::i _
i:i
E:I
!!:!_:_u
H:I
:].}
'I"_:Z!:
::t =
i[
! II'.tti!:[ ,,_fL .... i ill!l .......i_i.:
t:! .,-, ::::1!] ',:;I ::i
:i ::i;
.i::
:'
: r::r,,i::ri-rT-Fi
!i:!!!!h
::FI
:rx:
""::i!!!:i
....
;_::
i, ;.i
INLJ:
:::_ :;: .......... tq+"[i: :.!:
i4:ii!i!i!!
l_I?:t_ l[::_ud" }!_::i]i]1:::: , .; i_! ii] Iti4_*i_li[[i i!il ,-i:l:i:!:::'.: i:.:i it
..... [ i. i [ _]7 7_] el * _ r : ]
.xl
:::::[
[iii!ii!i
....
:]
: :
:::. :d: -- 2L::: _" :::. ........ _ ...... _ _
if!! !!! !!i!!iii i:!iii{:
t_ ?RT ::::::::
: *:? :x:
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:i[:
M:',
I:':
:
.
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i:: ::i - t ..... !!_i
'::; ::
! :_'I[[::[
_t_iq
. ,.+,.. :l::]
_., ::-! 77:_-_-7:71 T7 7:7777:77 ., .... :::[I]:K i:]: ]L
:: : ::: : : .......................... ii
........
i
....
:!! ;.
!:!
bl ::i i::2i:L !H ; .:: LLi {}!i[i! LL :
:
: '
[ [ t
;
l'
; :::4:::
i : I
'
;
......
.iii ii.! ii.:i2 :. :: {: i!
i!i i:! Hi
.i21/:i.;iiii-:LL_ !i!:!ii ri{::: ....
:!:! !:: ,!' ii:i
:____/L--LL :!
.... i! i .................. i ................ [i: : !!ii!i ; .......... !:i! !ili I:T:.-_L_,+ .E ::.: ?T:: !:.' i.: i':[ :[i!
!_It!P':I_!: I
tB::tl;,,; li:L:it
]L
i!i! :-
Lil
!:i !! +tTt T::
:i:
:: :_ ii! :!: :i!2:!!!!: :i]i ]i[.,
i[! : :-7TtM
/: i
::
i_Li]f:i
,, .... ," :
.... " :::::
::: ...........
:
T-i_
:::'
:i:-!!F!-:::!i!:!
+'T" T:T:_+-:T+7 :'7_:::
-i!
: 1
.......
:
:
...... .....
.
lililt :'
:;t
'ii!i '=
[ii !_;r'i:! _i: :: .... ii ::: :;V :: ;: Li2i i:: !:i :i:; i!! i:: I :+_L
TT:_7_7-T'tTT
_,_÷t"
777
:-'1
---*
: -7
i i;
_., ::::
r'l !
:. !] i:
I
: :!::'
I i i ii! ......
::
: ii
i.
' iiii:i .... , ::!!!!ii ..... !i!! .... ;! :
::1 ....
71";1I +ttl_
K
i-I i:i dh "i'
: .... X
:E
i::!:: ' ........... .... i ...... "+-T +i": "N" ::' [i :;:
xu i _:1 ....
::
',
: :::!
::
;:;
::::
od
•
_ iTiill...ii!!:.:,.:i,, . i_
:i:i
il'i
li:i
l:l
Iiii
::'d
:i
i _ iH!:
: TI_4E{$ECI
;i:ii
: iii:i:
!ili ....iiti .... i_:::: .... li:l[!iii'
ii 1_1.
Figure
4.7-49.
Inertial
Spin
Magnitude
and
Angular
Momentum
4. 7153
Figure
4.7-50.
Angular
Velocity
Vector
Compone_ts,
Inertial Coordinates
i:ti,ii ,
ii11: _r
'i!i:il_
{, i iiii_ii
lTlll;i i_ I: lllI-J_
:1 t!_ !
......
i i ,
t!
i
Figure
4.7-5].
Q
and
@
Inertial
Coordinates
4.7-54
_iittitttitTtti77it_ltt_ittl I!thtltt!hff!t_tttf_t_# h!tihtttt_!h!_ttEhr!ttfh_,l_ ::if_itt!tltt!titt!titf_t!tiiff_
i7
:H ,d _.
_i! I!-':i_i ! !!ll
ill
ih
:d:7:
lii! lh tt.tiHih I I hNdltt t+N_,NH#tttttIt t, i: W; :n ift:;N!tltl!til_tf{itttt_Nh: l] !t::{_ilttit_Fitt_ _t_tt_tllfttiitft_,tf!tittt!ftl!I::it!i::ilit!! !!i!Nitt!t!tt!Httittttit_flf_7_ _itiiiitii_::: :t i!!i_fi_}ti!!it]!i':tlit!tit!!tlI_f_f!tt_71ti_tf!ti!lit_il!t_!lL il::!ilfi _!!ltit]tt]{tttlttitl_t!!ttl !_iirtliiit!iiih?,:;t!';!iti!!_tiH_it_t!H_bt._ii_t!'}ttlttlti!l!!!ltTtT!fii::t iH!t_!t!!i:: iitlhi_,ftf_tt_tL_ttfl i!! 1:!i _i ! l ,U I_1_Lii4'1:_Itl, Frff._flHillHUt'l't .!t,,.L !ftHh:X_ t ,itll_,t,!l b_ !+_t+!_t,t, I ,, !ili_iti_h:-!!:_: iil :!!l::::::-/,:i::ii!_ ,t._ t,! I,. :i
; i " 'i t ' ! !b _i ! il ! _ * J{ ; '_ P i ]_ _ il i !t ii: t [t
f:httlft+ftJltt;t_t_,_,,
;i_i_ii
I_'
Ii
: i_ ..... • _ £ _ i[ L£ ; !_ +.i : fi: tT_ih 7i;_ : ',ii!7_:.+U:
:;i!iti!iit_i i!_!i
iiliili iiilii!i I _iiiTi i!_: !!rl i:!i_
_i!l_t_!t#:!l!f!ml_:!t!!i!hi_it[itdii_!lii,_t_i ! t:: i:f I iiilhi i t i i i _::i _! _ii_ :_ i!i7 i!i; !:Z! 7i71_!i:r 7i2! !i::it!!Tit::fiit_ti!Nitit!!IH!itit_ !H _H t!iiti_4i!il _3 iiT_ i7i17iii7iii7i !ii717ii!!!!77t77;7 iff ii::!tii!itiiiitiii!ti!tit!ltNiiiiti !Ii t,t,Lf,,t,t,ft,fX!:tlT_..4,+_tt!L!t._I!,,t_,_ !: _" H:t_ rlI t!,,_h.,If_tt_,l !! ti!iIi!t!tl!!_t!!flt _! !tltlI_.t! II_N :!:t_t_t fttt:tLt_! ._ i :h:!,tf I t,_+,t=.! tdt.t,!J.t,!Lh!t,, ht..t,..fr_dH_ I ti_ il_:!t!_iiiii::! !_i!t!i!!ti_i!t!!!i i?il t. ,tf,I: It !t, i! ti t f_ti!ici!!tlI itqi_Ltlt:;rTt!H!t!tit l_tt!f_J_lhithi!ii!!i!hi_ _qt:h7t:2"*' @i
i _ f • • ' • i 47 i,t s. _. t. iI. llll ii_ l t [ jill. _l ;;L .i
iiTi iiiitii!iN!!i!::_:
Figure
4.7-52.
Angular
Velocity
Vector
Components,
Spacecraft
Coordinates
[H
._ ,÷
il_l_il
I i? :, r: IT: _:
_f
;!
.I
_iii
i_iliiiil !i
:!
!i:
:i
i: i
Figure
4.7-53.
Q andO
Spacecraft
Coordinates
4. 7-55
'4--
a)
Inertial
Coordinates
b) Spacecraft Figure 4.7-54. Angular
Coordinates Momentum Components
4.7-56
i! li
I: _i! i, !i i:l :i: i
ilill !!!_lii:li_] i T
[_[:i:i]i
:,, , _.74___
!i:.l :!i!i:i
I!ll:i!iV_
i:
i:
It
!.ii_ #!!]
II!lI:
llli);; . :
I Z,/,ill, I
i! _,f I/I ',i_:Tl,
I /I I1/I
11.
!,i.++._,, r111
Figure 4.7-55.
Clock and Cone Angles for Sun Vector, Spacecraft Coordir_ates
:ll;
._+4+
+!.!.,!..
.......... !:;:i::
!i:l
i
i: i:lli
,! !i
........
+t'.
111; .+_
:: :I _
:!!'
,++_+_
t
: .... ili!
i!l
IIU
, ,+.,
:1
] +:= [!
:i
l!t! t!....
4-4_ ,H.
_H
+._+_.
,,++
TTT!Ti ;:,::,=it!!i!t!iI! s_i_: _
_U
HH l[!i
:i
_Si_Iisiilf!!i!!_i!!iliii liil I-
_iiiitiiiii ....
T:T!! _:
:i:ii:]
..... il;i : ,r, 5:
H4444_
f;a -!444
4a.a
$. |.
t4_
t_
7. 15.
ll,
II.
I?.
II.
II.
Figure
4.7-56.
Gyro
Output
Angles
4.7-57
4. 8
POSTMIDCOURSE
SEOUENCES
4. 8. 1
OPERATIONAL
DISCUSSION
Additional 2-second firings were recommended to attempt to clear the vernier engine 3 problem, and, if successful, to possibly restabilize the spacecraft. A firing sequence, using the midcourse thrusting level, was attempted at L + 18H56M and again at L + 19HI8M without success. Since the spacecraft was the only sources of power rotating such that for the spacecraft sohr loads panel were output was the main and
zero,
auxiliary batteries. To conserve energy, flight control coast phase power was turned off periodically (i.e. , power on for 40 minutes and off for 90 minutes) tures above An while maintaining the flight control 70 and 0°F limits, respectively. interrogation of modes 2 and gyro and electronic tempera-
4 at hourly
intervals
was
initiated.
Also, auxiliary battery mode was commanded when the auxiliary battery temperature was 35°F to utilize the energy of this battery and to keep it from approaching its lower operational limit. Since a possible cause of the vernier engine failure was a stuck fuel regulator valve, it was decided to pulse fire the engines five times (with a 0. Z-second period per firing and a 5-minute interval between firings) and then fire the engines for a 2-second interval. This sequence was first used at L + 31HI2M Four were and completed at L +35H2M. Engine 3 did not appear to fire. pro-
cedure
additional attempts to achieve thrusting with the same made at hourly intervals (i.e. , initiated at L + 36H28M,
L + 37H29M, L + 38H45M, and L + 39H45M), but all proved ineffective. It was then decided to try a higher thrust level with less rise time by placing the flight control subsystem in the postretro eject condition. This was accomplished by commanding retro sequence mode on and emergency retro eject prior to turning on the flight control thrust phase power, thereby preventing the ejection of the main retro engine while placing the flight control programmer in the desired state. This sequence was completed at L + 41HIIM with the commanding of vernier engine ignition for approximately 2 seconds controlled manually (i. e. , engine shutoff by ground command). Again, thc results were ne}zative. With each attempt to fire the engines, the spacecraft rotation rate increased so that by the time of the postretro eject thrusting completion, the spin rate was approximately I. 54 rps.
4.8-1
Between the second and third vernier engine firings, the planar array was commanded upward from its launch position to lower the solar panel for partial illumination. This was desirable for two reasons: l) to obtain more energy for the spacecraft, and 2) to illuminate sore( of the secondary sun sensor cells (mounted on the solar panel face) so that the actual spacecraft orientation could be established. Two attempts, at L + 38HI3M and L + 38HI9M, to move the planar array were unsuccessful, apparently due to the opposing force created by the spacecraft spinnlng. Preparations were then made for the follo\ving operations: i) stepping the solar pane{ illuminating its active face and th,. secondary sun sensor cells, 2) determining whether a zero-shift had occurred in the helium pressure telemetry signal by dumping the helium and recording the pressure decay function, 3) evaluating the capability of the main battery to continue to supply power reliably under the heavy terminal descent load conditions (i. e., flight control thrust phase power on, high power transmitter on, RADVS on, etc. ) when the remaining battery energy is low (i. e., on the order of 15 to 30 arnp-hr retnaining), and 4) firing the main retro engine in the normal terminal descent mode. At L + 42H22X4, the unlock solar panel squib was blown by ground command, resulting in a solar panel position telemetry signal change of approximately 23 degrees, indicating that the force on the panel created by the spacecraft spinning caused the panel to move. Further attempts to move the pane] by comnland were mostly unsuccessful. At L + 43H13bl, a new sequence for pulse firing the engines five times (0. 2 second for each firing, with 1 minute between firings), followed by a 20-second firing in the postretro eject mode, <_asexecuted, ending with the 20-second thrusting at Z + 43H33M. Although w_rnier engine 3 temperature rose approximately 24°F (as compared to approximately 100°F for engines 1 and 2) during the Z0-second firing, the. engine did not respond properly. At L + 44H411Yi,the helium dumping seque_ce was initiated, confirming that a zero shift in the helium pressure telemetry had occurred and accounted for the relatively large decrease when the system was initially pressurized. At L + 44H48M, flight control thrust phase power and RADVS were turned on. At this time, the estimated energy remaining in the main battery was l0 amp-hr. The bus voltage dropped from 19. 4 to 17. 3 volts, with a load of 47 amperes on the battery. RADVS was then turned off before proceeding with retro firing. At L + 45H2M, the emergency AMR command was sent to initiate the retro engine firing sequence. Ignition of vernier engines I and 2, as well as the main retro engine, were verified. Contact with the spacecraft was lost approximately 30 seconds after retro engine ignition.
4.8-2
Although there were no more telemetry or tracking data available, the spacecraft continued on its trajectory toward the moon, striking the surface at approximately 265:03:42:54 (flight time was 63. Z hours).
The landing ioc:atic_n is believed to be 0. 55 degree north latitude, 0. 83 degree west longitude. These data were taken from the last trajectory prediction made after rnidcourse, and are not as accurate as data from a normal flight.
4. 8. 2
ANALYSIS
OF
SPACECRAFT
ROTATIONAL
MOTION
Simulation of spacecraft motion during midcourse firing, as discussed in Section 4. _ depends on the integration of the equations of motion under a set of assumptions concerning engine performance and other pertinent variables. This simulation attempts to find the set of assumptions which allow best approximation of the observed data. It would be desirable to determine from independent sources as many of the parameters of spacecraft motion as possible, so that these could be compared to the values from powered flight simulation, of the simulation. Useful spin vector tum vector coordinates spin vector A spacecraft l) 2) 3) 4) 5) 4. 8. 2. l thus providing an additional parameters for this purpose check on the results are the motion of the
in spacecraft coordinates in inertial coordinates. can be determined from moves.
and the location of the angular momenAngular 1_on_enturn in spacecraft the spin vector and will change as the
number of independent types of data rotational motion after midcourse. Spacecraft Gyro temperature shortly received panel distribution after
provide These
an indication of are as follows:
crossovers of DSIF of solar
vernier
shutoff
Variation Behavior Retro Spin
signal
strength servo retro firing Distribution
elevation during
accelerorneter Orientation
output From
Vector
Temperature
The variation of temperature as a function of location on the spacecraft is the most positive indication of the general direction of the spin axis, although these data provide a less precise determination of the direction of this axis. The thermal analysis indicates that the spacecraft-sun vector was in the quadrant bounded by the +X, -Y, and +Z spacecraft axes (Reference i). Figure 4. 8-I shows the bounds of the probable sun vector locations. The spin vector would also have to be within this envelope.
4.8-3
COMPARTMENT A
y ,i
COMPARTMENT B
.T_ CT'lo4,j OF PK #NC I P'#_ L X )('-Y pLANE
@
AUXILIARy ENVELOPE VECTOR OF POSSIBLE LOCATIONS SUN//
"o, = 28._
BATTERY
Figure
4.8-I.
Probable
LocatioF_
of Sun
Vector
From
Thermal
Analysis
4.
8-4
!
4. 8. 2. 2
Oyro
Zero
Crossings
and
Postmidcourse
Tumbling
Dynamics
When the vernier engines were shut off after midcourse, the spacecraft was spinning with a period of about 0. 8 second, and the spin axis was precessing with a period of 12. 5 seconds. About 7 minutes after start of midcourse thrust, the precession was essentially damped out by the nonconservative forces in the spacecraft and the gas jet operation, resulting in a pure spin. The axis of spin was that spacecraft principal axis of inertia with either the largest or the least moment of inertia. In this case, analysis shows that it was the former. This principal axis would also be the axis of precession after midcourse matrix of the spacecraft, to be as follows (Reference O = = and later firings. By diagonalizing the inertia the direction of this principal axis, R, was found 2): -54°15 71o13 ' 01 ' ' or 62 -- _41o50 : 41. 8 ° : 28. 9 °
where@,_5, @'@l'and 02 are defined in Figure 4. 8-2. When the been damped out, the pitch and roll gyro error signals were tively and the yaw gyro error signal was saturated positively and inertial mode). This verifies that the spin vector was in indicated by thermal analysis. Since the spin is about a principal axis, this axis must
precession had saturated nega(both in rate the quadrant
also
contain
the angular momentum vector. Figure 4. 8-I shows the location of this vector relative to the sun direction. If the angular momentum was about the computed principal axis, it can be seen that the projections in the XZ plane of the sun-spacecraft vector and the inertial angular momentum vector were within about 20 degrees of each other. An attempt to determine, independently of the powered flightsimulation, the angular motion and orientation of the spacecraft at vernier shutoff was made by fitting the pitch and yaw gyro zero crossings after vernier shutoff to }_ulers equations for a force-free tumbling body, as follows:
Ii# l
(i - 13) wzw 3 2
Iz# z = (I 3-I l) wBw I i3_¢ 3 : (i 1-I
whe re --angular = moments rates about the about principal these axes axes
z) wlw z
(1)
W1,
W2,
and
W3
I 1 > 12 > 13
of inertia
4.8-5
Z
_, • ×
• _
R-:._ -- cos co,
= COS
-Z
-_
Figure
4. 8-Z.
Definition
of
Direction
Cosines
and
Unit
Vectors
4.8=6
The
solution
of these
equations
is the
following
Jacobian
elliptic
functions:
W 1
=
adn
IP(t-
to) ]
W 2
=
[3sn
[P(t-
to) ]
W 3 A typical in plot Figure of sn(X), 4. 8-3. cn(X)
=
Ycn -
[P(t2 , R are
tO) ] and dn(X) = V/1 which - k2sn(X) depend 2 on is
(2)
= V/1 P,
sn(X) and
shown
or, _ , Y,
constants
I1, I2, and I3, the total angular momentum and total rotational energy. Note that W1 =_dnlp(t - to) ] never changes sign. This is because the spin vector precesses about principal axis 1 and is never more than 90 degrees away from it. W 2 and W 3 do cross zero due to the precession (Figure 4. 8-4). Later, when the precession is damped out, W 1 = W and W 2 = W 3 = 0. The pitch and yaw gyro data following vernier shutoff are shown in
Figure 4. 8-5. The roll gyro was in negative saturation throughout the period. Although the gyros were in the inertial mode and would have normally measured angular position, in the present case the only quantitative data that can be derived are the times of zero gyro rates. The high angular rates precess the gyro output axis against the stops until the angular rate reverses polarity (due to precession). When the angular rate changes polarity, the gyro is precessed to the other stop. If the gyro dynamic lags are ignored, the gyro will leave the stops at exactly the time the rate reverses sign. The gyro telemetry measurement will unsaturate slightly iater because the telemetry range is less than the gyro range(8 degrees versus 15 degrees). Thus, it can be assumed that Wp or Wy = 0 shortly before the pitch or yaw gyro telemetry measurement unsaturates. Figure 4. 8-5 shows that Wp is negative for longer periods than it is positive, and Wy is positive for longer periods than it is negative. W Z is always negative. This results from the fact that the spacecraft X, Y, and Z axes are not coincident with the principal axes. The equations for pitch yaw and roll rate are as follows:
Wp
=
All
W 1 + A12
W 2 + A13
W 3
Wy
=
A21
W 1 + A22
W 2 + A23
W3
W Z
=
A31
W 1 + A3Z
W 3 + A33
W 3
(3)
/
4. 8-7
2K
X
Graphs
of
sn
x,
cn
x,
dnz
(k I
-
0.7).
Figure
4. 8-3.
Jacobian
Elliptic
Functions
Figure
4.8-4.
Precession
of Spin
Vector
About
Major
Principal
Axis
4.8-8
0
0
4_ ,,'4
_4
!
4
°r,,I
4.8-9
where Aij are the direction spacecraft axis. Numerically Wp lwl 0. 5841
cosines
between
each
principal
axis
and
each
_ dn
° I
ip(t
P(t
- to)
l
+ O. 4431
snIP(t
- to) ] + 0. 5357
cn
[P(t - to)]
Wy__=
_
_ to
o _ l6 IP 1-o.832 IPItto _z_o.74 odnlplt_too. 2 2sn t_to 1 cn - l iwI v
It can be seen that these equations are qualitatively consistent with the gyro data. Since _ is a positive constant and dn[P(t - t0)]is always positive, the term_dn[P(t - to) ] produces a bias in each gyro measurement. From Equation 4, pitch and roll rate have the same polarity bias yaw has the opposite bias as observed in the telemetry. Also, the negative bias in roll is larger than in pitch, accounting for tl_e fact that roll rate is always negative. The ratio-_ and P depends only on the
• (Y
(4)
inertia
matrix
which
is
presumed to be known and a constant R. Smce _ depends only on the inertia matrix, if value of R can be found which causes WX, Wy, and W Z to fit the telemetry data, the spacecraft rotational motion can be completely described inspacecraft coordinates for the period following the midcourse maneuver. A computer program that computes Wx(t), value of R has been written. A fairly good (Figure 4. 8-6). However, it is not believed zero crossing times is sufficient to account data and the simulation. Furthermore, it not possible with the inertia matrix used. would be necessary to change the inertia Wy(t), and WZ(t) for a given fit to the data can be obtained that the uncertainty in the gyro for the differences between the can be shown that a better fit is In order to improve the fit, it to shift the principal axes. dynamics whether as well
matrix
A similar conclusion arose in the analysis of the spacecraft during the thrust period (Section 4. 7). It has not yet been determined or not there is enough independent data to refine the inertia matrix, as determine the parameters of the motion.
An attempt was made to fit the solution of Eulers equations to the telemetered strain gage measurements since, with the engines off, the strain gage output is due entirely to spacecraft angular motion. However, it was not possible to fit these data, apparently, because the formula for strain gage output did not include The response of the strain the effect of torsional stress on the gage output. gages to torsional motion is being measured. axes would was such that have made it
there
It should be noted that the location of the principal was not time(when the orientation of the spacecraft
possible)to cancel out the spin rate by firing the vernier engines. However, this was not true in general. If an engine other than W 3 had failed, it would
4.8-10
0
°1-4
oq
T
r h
0
> 0
E
,0
0
1
tn
}
J
i
°,-4
_.°
!
+
4
k
4.8-11
have been theoretically possible to find a time, while the spacecraft was precessing, when the torque vector, due to the thrust imbalance, would have been opposed to the angular momentum vector and thus capable of reducing the spin rate.
DSIF Signal Variation Due to Spin
Another source of information concerning spacecraft motion after the midcourse maneuver is the oscillation in signal strength observed at the DSIF. These data have been used in the analysis of the motion during midcourse thrusting (Section 4. 7). After n lidcourse, these data were used to determine the variation of spin rate with time and to observe the effect on spin rate of the 39 vernier engine firings that followed midcourse (Figure 4. 8-7). Solar On Panel Motion the solar panel was unlocked by radio command. The
265:06:35,
panel subsequently moved from its transit position at 270 degrees to a position of approximately 249 degrees (Figure 4. 8-8). The solar panels were then stepped 87 times in the negative direction, which should have n_oved them i0. 9 degrees to 238. 1 degrees, but the panel only moved to 246 degrees. Thus, it appears very likely that the spin axis projection in the XZ plane is perpendicular to the solar panel when the panel is between 246 and 249 degrees. Then the component of the spin vector in the XZ plane is between 21 and 24 degrees from the -Z axis (]Figure 4. 8-i). This is within the range of values consistent with the resu]ts of the thermal analysis, but is in disagreement with the location of the major principal axis (which should also be the spin axis) whose projection on the XZ plane is 41. 8 degrees from the -Z axis. The disagreement between the location of the principal axis and the spin axis indicated by the solar panel motion could be due to errors in the inertia matrix. However, it should be noted (Figure 4. 8-1) that the XZ component of the spacecraft-sun vector location estimated from the thermal analysis is between 20 and 60 degrees from the Z axis. The spin vector calculated from the solar panel motion is near the extreme end of this band (21 to 24 degrees from the Z axis). It would be expected that the spin vector would be further inside the band of Figure 4. 8-1, and it appears that the results of this figure are more consistent with the spin vector indicated by inertia matrix considerations than with the spin vector from solar panel motion analysis. solar panel motion component can be Retro Unfortunately, the significance ascribed above to the is questionable. Also, no measure of the spin vector determined from the solar panel motion. Data also An produces a centrifugal (Reference acceleration been on to Y
Accelerometer rotation
Spacecraft the retro determine spin rate
accelerometer.
analysis
3) has
performed
the spin vector orientation (92. 3 rpm), would produce
which, combined with the retro accelerometer
the measured output observed
4.8-12
• -----. INCLUDE3 ALL 8uRnS .... Q OMITS o.2 s_'r. BuR_J.S ISg
I.(,7
RpM
,,-'_
. •
TO GAS 3E_3
r._r
j_
l
CUMULATIVE
_URIU
-I-IME_
$_CONJP..T
Figure
4. 8-7.
Spin
Rate
4.8-13
-7--
AFTER £7 "STEP M|_,,/L,,'_'! (OMM _'AFTER 2/4.4_ /_NLOCK
ANDS
PosITiON I I wK INDICATL='D LoCATIoM \ oF PR:o'o"ECTIOP, J OF SPIIv V_'CTOR
iN
XZ.
PLANE
×
Figure 4. 8-8. Spin Axis Location From Solar Panel Motion
4.8-14
during the period just preceeding the 20-second firing. The results are shown in Figure 4. 8-9. The spin axis would have to lie on the contour of the ellipse-like curve to cause the observed retro accelerometer output of 3.42 g. Orientation of the major principal axis is also shown in this figure. This principal axis, which should be the steady-state location of the spin axis, falls reasonably close to the curve. The difference is about 5 degrees in @Iand 4 degrees in @2" The uncertainty in the location of the curve is due to accelerometer errors, errors in spin rate, and errors in the location of the center of gravity relative to the accelerometer are probably considerably greater than the distance between the curve and the point that represents the principal axis. The uncertainty in the calculation of spin axis location from retro accelerometer data will receive more detailed examination. The results presented here are somewhat different than those in Reference 3. The results given in this reference were obtained from data taken just after the 20-second retro firing when the spin axis had been torqued away from its steady-state position and when the spin axis was precessing. The results presented in this report were calculated from retro accelerometer data taken just before the 20-second firing (265:08:02). 4. 8. 3 SUMMARY The values of 81 and @ for the steady-state 2 spin vector (Figure 4. 8-2), derived from the various data sources, are listed in Table 4. 8-i. The angular momentum vector in inertial coordinates appears from the thermal analysis and inertia matrix considerations to be within 20 degrees of the sunspacecraft vector.
TABLE
4. 8-I.
SUMMARY
OF
SPACECRAFT
ROTATION
AXIS
ANALYSES
81, Thermal analysis
degrees ;:-"
82, -30
degrees to -70
Notes
Inertia
matrix
Z8.9
-48. 5
Probably most determination
accurate
DSIF signal strength variation Solar panel motion -66 to-69 Questionable
Retro
accelerometer
-8 to +34
-35
to-55
@l and82are (See Figure
correlated 4. 8-9)
;:-'Cannot be
determined analyzed
from for of @l
this data this purpose. and @2"
source. Data does not appear capable of
;:-_;:-'Has been not improving
estimates
4.8-15
_"
RA_G-E oF ®2. CoNS _-FE_rF
D-
40-
30-
/O-
-/0"
Figure
4. 8-9.
Eoci
of Spin Axis
Deterrninat:iot_s by All Methods
4.8-16
4.8.4 i.
REFERENGES "SC-Z Tumbling Mode 5 October 1966. "Principle Axes Thermal Observations, "
H. E. linudson, IDG 2221. 19/86, L. M. 1966. Bronstein,
2.
of SC-Z,
" IDC
ZZ9Z/II4,
8 November
3.
E. W. White, "Study 7 December 1966.
of SG-2
Retro
Accelerometer
Data,
" IDC
2293/50,
4.8-17
4.9 RELIABILITY
ANALYSIS
4.9.1
PERFORMANCE
Assessment of performance from a reliability standpoint mainly concerns relevant failures and unit operating time. To date, there are four failure modes (TFRs) (Reference I) pertaining to the mission. They are listed in Table 4.9-I with descriptions and current status. Unit operating experience is listed in Table 4.9-Z with unit part and serial numbers. The unit operating time and cycle information was developed by translating commands transmitted to the spacecraft. 4.9.2 PERFORMANCE VERSUS PREDICTIONS
The predicted reliability for the transit phase was 0.66. The growth pattern of reliability estimates prior to launch is shown in Figure 4.9-I. These predictions excluded consideration of the use of nonstandard procedures. For comparison, the SC-I growth pattern is included.
Although two mission attempts, successful or not, cannot in themselves completely justify or vitiate prediction methods, data collected during those missions do serve as a basis for investigation of areas of possible improvement in prediction. 4.9.2.1 Reliability Math Model (Nonoperating Equipment) is that equal to
An assumption electronic equipment 1/100 of the failure
used throughout the reliability math model in the nonoperating state has a failure rate rate during its operating state
koff
=
0.01
k
on
Experience too high.
gained Detailed
on both analysis
SC-1 and SC-2 of this parameter
indicates has
that been
this factor initiated.
may
be
4.9-1
TABLE
4.9-1.
MISSION
TFR
SUMMARY
TFR Number 18247 At legs prior Description approximately extend signal to Centaur the (35 time of the the Status Closed. (on i December Cause pulse 1966)* of failure sensi-
Relevant. to noise
seconds
is attributed
separation) group mode from
flight control sensor reverted to inertial rate mode (should in rate mode until after 18248 Centaur
tivity of flight latches. ECA drawings corrective porated spacec Open. provide
control output 11175 and related
have remained 52 seconds
284544 and 284546 provide action and will be incorand subsequent
separation). failed to maintain midcourse
for SC-3 raft. Failure detailed
Spacecraft
Review
Board
to
stable attitude correction. 18249 Receiver B was
during
analysis.
reading
below of
Open.
Under
investigation.
specification values from 90 minutes before launch to end mission. 18250 Canopus automatic sensor star failed
to achieve
Open.
Affects
only
spacecraft by requirelockon. of failure engineeronly.
acquisition.
operational ing manually Closed. is unknown. ing data on
procedures commanded Relevant. Loss launch Cause
18251
No intelligible received from signal CY 53 0 during
data were commutated launch. data were the commutated
of some vibration
of flight accelerometer
18252
No intelligible received from signal CY 52
Closed.
Relevant. Loss launch
Cause of some vibration
of failure engineeronly.
of flight accelerometer 0. line 2 heater was full
is unknown. ing data on
18253
Vernier
Open.
ECA
113043
initiates heater assembly
on and line temperature still decreased prior to midcourse correction. 18254 Helium (P-l) "zero squib tank pressure sensor
investigation of line unit level tests.
Closed. was P-I shock helium at squib
Relevant. seen release. tank by
Cause the For pressure
of failure the
experienced a -528 psi shift" at helium release actuation.
transducer SC-3, signal
will be displayed in analog form during squib release to permit immediate positive verification to of a "zero shift" prior midcour se correction.
* For
additional
data,
see
subsection
3. i.
4.9-2
TABLE
4.9-2. TIME
MISSION B AND CYCLE
UNIT OPERATING DATA;:'
Time, Part Subsystem and Unit Number Serial Number (or flight n_ber
hours of )
cycles
Telec
atom Central EngHxeering Au×_hary Signal Lo_
unicationa command signal engineering processing data rate auxiliary _uxiiiary antenna antenna mechanism nlechanisnt _ignal A B A r pass pass g filter filter buffer buffer s_/tch s_itch A B A B proceas,,r A B A B decoder processor signal processor 23Z000-5 233350-7 2b4900-3 232540-1 264875-2 232400 232400 287300-1 273880-1 232200-8 263220 263220-4 231900-3 231900 233460 233466 290780 290780 Zg3_84 283983 2 cycles 3 4 45.6 45.6 45,b 45.b 45.b 45.6 45.6 45.6 7._ 3b,6 1.0 29.0 45.b 45.6 1 cycle I cycle 45.6
Omnidirectional Omnidirectional Omnidirectional OnlnidirectionaL Central q ransmitter I ransm_tter elver
Rec Receive Low Lo_,
lelemetry Telemetry RF tzanafer SPD7 Te ie visi_,n Survey Appruach Television Vehicle mechanism_ TherznM Thermal Thermal Thermal Thermal ThermM Thermal Spaceframe Engineering Landing Landing Landing Footpad Footpad Footpad Crushable Shock Shock Shock Wiring Wiring Wiring Wiring Wiring Wiring Wiring Wiring Wiring Wiring Antenna Roll Solar Polar Elevation Separathm sensing mAar gear gear gear leg leg leg absorber absorber _haorher harness harness harneas hornenn harnena harnes_ harness harness h_rness harness camera RF
28431Z-3 284302-1 232106-5
camera auxiliary
sen_ors control c:cmtrol s_itche_ swztche8 _heH shell
(total and and he_ter beater
f_r
231 assembly ausembiy A B A B A B
988653 23ZZI0 232210-2 238810 238811 Z86459 28b460 264178-1 1
1048.8 45.6 45.6 41_.4 273.6 45.6 45.6 45.6 45.6 1 1 1 cycle cycle cycle
cc,mpartment compartment comparmlent compartment
mechanism 1 2 3 1 Z 3 bluckm leg leg leg 1 2 3
auxiliary
263500-6 261278 261279 2blZBO 263947 Zb3947 Zb3947 264300-1 _b4300-I 264300-I A B ZgbZ07 Zgb24Z 3025357 286398 panel posltxuner 266417 264100 276979
1 cycle l 45.6 45.6 45.b 45.b cycle
1 cycle
compartment compartment b_Bic basic antenna auxiliary TV RF retro battery p_sitiuner camera cabling motor cell bus bus 1 2 solar battery
3.0
36.9 286390 voltdge 3025155 287580 700 1001 cycles cycle_ 0.1 15.b
pane[
and
_rming
device
293400
I
cycle each
Propulsion Betro Vernier Vernier Vermer Electrical Battery Boost Auxiliary Main Main A axilt.t Boost Boost Solar Flight controls Flight cnntrol Coast Thrust Radar and Signal Rlystron Altitude Velotity w ,, vegmde Altitude Roll Alt_tude Altitude Altitude Secondary Pin Pin ¢'Data pullers puller sc.arce: cartridges DSS tapes. tnalklng actuatc,r jet Jet jet leg leg leg s_m I 2 3 sensor radar guidance data pha_e phase RADVS converter power velocity sensor supply sensor antenna antenna 23290g-ZAM7 232909-AM3 232910-AM4 Z3Z911-1AM3 23291Z-AM3 283827-1 235900-3 235700-2 984 235700 235700-3 235450-1 3 984 904 22.8 9 9 cycles cycles cycle_ cycles cycle_ 0.Z 0.2 0,2 0.2 0.2 0.1 3.0 sensor group Z35000-9 38,5 3.0 power battery ry battery unregulated unregulated filter choke rocket engine engine engine power charge regulator battery _witch control regulator 274100-4 &74200-1Z 273000-2 254112 237900 Z37921*1 _90080 290390 2377h0-_ syJtem 1 Z 3 238612 285063-1 285063-2 285063-3 A21-Z7 54Z 54b 544
45.6 45.6 45.6 45.6 45.6 3.0 45.6 45.6 45.t_
regulator regulator panel
4.9-3
1.0
0.9
0.8
0.7
A I r
>I,-.,,.,,I
0.6
0.5
i ,-,.I us
0.4
h
0.3
sc_.
0.2
0.1
0
200 SURVEYOR
400
600
800
1000
1200
1400
SPACECRAFT SYSTEMS TEST EXPERIENCE - HOURS
Figure
4.9-I.
Reliability
Estimate
(Flight
and
Landing)
4.9-4
4.9.Z.Z
Reliability
Math
Model
{Boost
Vibration
Effects)
Boost phase {vibration stress) failures in electronic equipment are assumed to occur at a rate equal to 80 times the failure rate during nonboost periods,
kon
boost
=
80
kon
koff boost
=
80
kof f
This factor may failure rates. 4.9.Z.3
also
be
too high
and
is included
in the analysis
of off time
Reliability
Math
Model
{Propulsion
Subsystem)
Data-based estimates of the propulsion subsystem reliability assumed a binomial distribution of successes and failures. Experience gained on SC-I and SC-2 test programs and flights indicates a possible requirement for inclusion of additional parameters. In particular, within the vernier propulsion system, the following are presently under investigation: distribution of the difference between thrust realized and thrust commanded, effect of these dispersions subsystem upon moment success. control, and effect upon probability of propulsion
4.9.3
FUTURE
RELIABILITY
PREDICTIONS
Reliability predictions for future spacecraft will include SC-2 mission unit experience, as well as SC-I transit and lunar phase experience, where there are no significant design differences among units.
4. 1.
9.4
REFERENCE Relevant Failures, " IDC 2258.2/328, 24 February 1966.
"Reliability
4.9-5
4. i0
ADDITIONAL The
REFERENCE
MATERIAL specific acknow3, and 4: Aircraft Company,
ledgement 1.
following reference material was used without in the preparation of many parts of Sections 2, Final Report, " Hughes
"Surveyor I Flight Performance SSD 68189R, October 1966.
Z.
"Spaceflight Operations Plan - Surveyor Revision S/MB, September 1966. "SC-2 Consent to Launch," September 1966. "Surveyor Aircraft "Surveyor Company, SC-2, Company, Hughes
Mission
B,"
JPL
EPD
180,
3.
Aircraft
Company,
SSD
64229R,
4.
ETR Test Phase Data Evaluation SSD 642081, September 1966. Space Flight Operations November 1966. Analysis and SSD 64Z60R,
Report,"
Hughes
5.
Mission B, SSD 64257R,
Report,"
Hughes
Aircraft
6.
"Surveyor If, Flight Path Hughes Aircraft Company, G. A. Young IDC 2292/91, B. D. Love, 292-66-219,
Command November B
Operations 1966.
Report,"
7.
to R. H. Leuschner, I0 October 1966. "Surveyor 13 October II Transit 1966.
"Mission
Command
List,"
8.
Command
Sequence,"
JPL
IOM
4.11
ACKNOWLEDGEMENTS The material or general mention Bronstein, for for in Sections 1, Z, 3, and 4 was and 4. people: Section Section coordinated R. H. Leuschner. 10 which were (and in many
cases addition used,
originated to the special L. Wo L. M.
compiled) by G. A. Young references in subsection is due to the following for final Section the analysis in of on
In frequently
4. 3
8
McIntyre, K. Cooley,
compilation 4.9
reliability labor that through produced analysis,
P. E. Sterba, Section 4.7, programming,
for direction and to all those or writing:
of the tremendous who also contributed
F. R. Fagerlund D. J. Giem J. H. Green J. D. Haller W. R. Heathcote E. R. Kopitzke 4.9-6
J. McFerson G.D. Passey A.Z. Reynolds F.K. Rickman M.R. W einer
5. 0 PERFORMANGE ANALYSIS
5. 1 THERMAL CONTROL SUBSYSTEM
5.1. 1 INTRODUCTION 5. 1. i. 1
Surveyor Thermal Control design passive Techniques uses a variety of temperature control systems are employed to provide the
The techniques.
Surveyor thermal Both active and
required temperature control throughout the transit and lunar phases of the mission. Each spacecraft subsystem is individually controlled, and the thermal coupling between subsystems is minimized by using conduction and radiation isolation wherever advantageous. Subsystem analyses are accomplished by evaluating in detail the thermal environment for each subsystem, with consideration being given to all significant interactions between the subsystems whenever a high degree of isolation is not possible. The following spacecraft: temperature control techniques are used on the
Surveyor
l)
Passive thermal control utilizing combinations metal processes to provide solar absorptance emittance characteristics that produce required temperatures.
of and
paints infrared subsystem
and
2)
Active thermal energy in cases available.
control where
systems utilizing heaters to provide sufficient solar illumination is not
3)
High conduction and radiation for systems having a large reach equilibrium conditions stored heat capacity.
isolation utilizing supe rinsulation heat capacity. Such systems never and therefore depend on their
4)
Bimetallically activated thermal switches perature of the electronics compartments lunar operations. of the above temperature techniques control are system. used on many
that control the temduring transit and
Combinations optimize
of
the
subsystems
to
the
5.1-i
5.1. i. 2
Analysis
Organization
The spacecraft has been divided into a number of subsystems for thermal analysis. The thermal behavior within each of those listed below discussed in subsection 5. 1.4, with comparison to test and SC-1 data. 1) 2) 3) 4) 5) 6) 7) 8 ) 9) i0) ll) 1Z) 13) 14) 15) 16) 17) Con_partments Auxiliary A/SPP Spaceframe Landing Thrust gear chamber tanks line s and crushable assemblies blocks battery A and B
is
Propellant Propellant Helium Main Flight Roll
tank retro engine electronics and Canopus sensor
control actuator tank gas
Nitrogen Attitude RADVS Altitude Television
jets
marking system
radar
Included in subsection 5. i. Z is not only a discussion of the vernier line thermal anomaly, but also considerable analysis, as outlined below, done in support of the vernier engine anomaly: l) Z) 3) Analysis Vernier Tumbling of vernier burn mode thermal thermal system temperature data
inconsistency observations
5.1-2
5.
1.
1. 3
Major
Events
and
Times affected the thermal subsystem are of mission time. A complete tabulathrust power on periods, and all 4. 1. 1 of the system discussion.
tabulated tion of vernier
Some of the major events that in Table 5. 1-1 as a function spacecraft high power periods, burns can be found in subsection
5. 1.2
ANOMALIES
AND
FAILURE
SUPPORT
DATA
Only one primary anomaly existed in the thermal control subsystem: vernier line heater cycling (see subsection 5. I. 2. l). Also included here are two extensive analyses prepared in support of the vernier engine failure investigation (see subsections 5. I. 2. 2 and 5. i. 2. 4). Study of the data from each of the 40 vernier engine firings after midcourse disclosed a second potential anomaly, since engine burns of the same duration sometimes produced different thermal results (see subsection 5. I. Z. 3). This secondary anomaly may possibly have been caused by spacecraft spinning following the midcourse attempt. One of the anomalous burns (number Z7) is probably due to a burn interval that was more than twice as long as commanded. But the remaining burns are discussed here in the absence of other plausible theories. 5.1.2. I Vernier Line Thermal Anomaly
An examination of sensor P-4 thermal data presented in Figure 5.1-33;.'-" indicates that the heaters on the propellant lines feeding vernier engine 2 began to cycle 90 minutes after launch. Thermal data indicate that the heater operational duty cycle increased with mission time during the first 4 hours of flight. The cycling exhibited by the vernier line 2 heater terminated at approximately Z + 3H, and the line heater remained on. Termination of vernier line heater cycling during the course of a mission is considered a thermal anomaly. Since the spaceframe and other subsystems in the proximity of the not reach their respective steady-state equilibrium temperature increase in the line 2 heater operational or "on" is considered normal. The thermal response in the vicinity of the engine 2 propellant lines is P-16, and V-38 (Figures 5. 1-38, 5. 1-44, and time
lines
do
during coast phase I, an as the mission progresses exhibited by subsystems shown by sensors P-10, 5. 1-67). As environment colder, so as to creating increase the spaceframe continue a
to
and cool,
other subsystems in the thermal environment for more to the energy upper
the from limit
propellant of the
lines
line becomes
greater the line
demand temperature
the line heaters (26°F) of the heater
;:-'SC-2 thermal mission plots are located at the predictions superimposed on these plots were 5.1-3
end of the section. The thermal taken from Reference ].
o m 0 0 _ 0 L_ _oo ¢_ 0 ..-..
0
"_
_q _q < b_ _q :> _q < g
,-, r_ 0
._
_:
_
',_
•
._
__
o
,_
_>"
o
._ ,_ o<_--_o_._._o ,_
_ ¢) 0
_= oo_._
_ _ _; ;::1 ._ 0 .,-_ r_ _ _ _ ._ _ ..._,..C "_ _ _ ,.a .,_ _,_ , _ 0O r.,o _>>'_ _.,_ o
U:
=-
0
..........
,.
....................
o
©
I °_
0
_
_
_
_
_
_
_
_
_
_
_
.4 _q ;q <
u O_ 0_00_
.o
_
_
00_
_ 0 _ _00_0_0
O_
_
_
L__O_ o. ............ ,o
, ...........................
__0_
0000000
0
"a
f_lO
I
eq _
I
r,q _1
I
o_
v
o_
_
c_
v
5.1-4
thermostat
range.
Since
the
thermal
dissipation
{P
= E2/R)
of the
propellant on
line heaters is relatively increased energy demand time,
constant for a given supply voltage, the lines' is achieved by increasing heater operational or
The vernier engine g propellant line heater on time is illustrated in Figure 5. I-i as a function of mission time for SC-I and SC-2. Effects of the changing thermal environment and the decay in bus voltage on the line temperature level for SC-I is readily observed. The curves indicate that the line thermal environment continues to cool throughout the mission. The ordinate of Figure 5. I-I is simply the heater on time divided by the total time for one cycle, that is, on time plus off time for any cycle taken at a discrete tinge interval during the mission. The data presented in Figure 5. i-I indicate that the SC-2 vernier line 2 heaters reached a saturated state both during solar thermal vacuum testing and flight. Current data indicate that the line heaters were on and functioning properly; however, the line temperature continued to decrease. SC-I thermal data indicate that the line g heater maintained the propellant lines within the cyclic deadband range of the thermostat throughout the mission and that this heater did not saturate or reach a continuous on condition. The maintain Thermal 5. 1-2 for heater to suggests thermal dissipation capability of the line heaters was sufficient to line temperatures within the thermostat cyclic range of 19 to 26 °F. behavior of the vernier engine 2 oxidizer line is shown in Figure the Mission A transit coast phase. Failure of the SC-2 line 2 maintain line temperatures within the thermostat deadband range that the ernittance of the exterior surface of the aluminum foil value. exterior
wrap covering the line heater may have exceeded the specification The line heaters are designed to give satisfactory operation with surfaces whose emittance is less than or equal to 0. I0.
The propellant lines are thermally controlled by means of active thermal control techniques. Ideally, the thermal design attempts to decouple the lines from other subsystems by using thermal surfaces that are insensitive to infrared radiation interchange. The use of low infrared emittance surfaces also minimizes the net heat loss, by radiation, between the lines and heat sinks in the environment. Increases in the infrared emittance by the wrap covering the line tends to: l) couple the line with its thermal environment, Z) increase the line radiation into space, and 3) increase the power required to maintain the line temperature within the thermostat deadband range of 19 to Z6 °F. Thermal analyses indicate that emittance of the exterior surface of
the vernier engine 2 oxidizer line could have been in the range of 0. 30 to 0. 50. However, because of the many variables affecting the propellant line energy balance, the exact oxidizer line emittance cannot be determined more precisely. Therefore, the range of emittance values presented in this discussion is not to be taken as exact, but serves to demonstrate that the emittance on the exterior surface higher than expected. Ultraviolet of the oxidizer 2 line was significantly (black light} inspections at Cape Kennedy
5.1-5
._-,I
0.o,2,
(_ ._._
I
.0
I
C_
.......... ................
I I I I
+
J" u-I-H-HH+H_+H-tH4ftP 1- ""_ ....................... _ _4-=-:f_t _
"'i
Ell ............ ilI4111111111 I"IIMII +]I, ] _l_l I I I ......... I] III LI I I i
:, _!,,
"[_' 'i 'ii'!T!4}f;4}
i-,,
............ ]11 [I I ILI LI Iii','"
_+
4
i , , , ,M--,I--L[ l I 11. I I y"
.......
bt-FH-Ftd4-H+t+t+_
' '
I _11_
: t
II
iil
I I I I ] I I I , i i I I I i_'_.ll I_I
'
_
_' LI
_
": -
'
:
' -'
I _lll"l '''''';'_'',, _,,,,,,,,.,
II I I I I I I I II I III
.,.
_
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Figure
5. I-2.
Vernier
Line
2 Temperature
Cycling
During
Mission
A
5. I-7
revealed that the exterior surfaces of the vernier engine 2 propellant lines were contaminated with epoxy residue which would not be detected during routine daylight thermal inspections. oxidizer In conclusion, line exterior
Thermal Review Thermal during
it can be said that the high emittance of the vernier surface is the probable cause of the thermal anomaly.
of Vernier System Temperature Data (Failure
5. 1.2. 2
Analysis Support)
manded
data indicate the midcourse
that vernier engines correction maneuver,
1 and but
2 ignited as comvernier engine 3 did
not exhibit any positive indications of ignition. The apparent failure of thrust chamber assembly (TCA) 3 to ignite as commanded resulted in spacecraft tumbling. Table 5. 1-2 is a summary of propellant flow determinations during long burns. Individual burns are discussed in the following a rg ument s. Oxidizer An cated by Flow Arguments of the vernier line temperature on the vernier responses, oxidizer lines, as indi-
examination the flight
sensors
installed
shows
positive temperature changes on oxidizer lines l an(] 2 at midcourse. The thermal data show only a small negative temperature perturbation on the oxidizer line feeding vernier engine 3. The vernier line temperature profiles are shown in Figure 5. 1-3 for the midcourse interval. Vernier line temperature perturbations may result from line heater
cycling, spacecraft attitude other than the nominal transit attitude, or propellant flow which is at a temperature different from the line temperature. Effects of vernier line heater cycling are readily distinguishable in the thermal data because of the constant amplitude cyclic waveform exhibited by the line thermostat on-off duty cycle. Nominal transit line temperatures may change as a result of a misalignment in the vehicle sun attitude whereby the vehicle sustains a yaw or pitch maneuver. Line temperatures may increase or decrease depending upon the relative orientation of the vehicle with respect to the sun vector. In general, decreases in line temperature will be observed be observed by increase in the b[ an increase in the line heater duty cycle; increases can a decrease in the line heater duty cycle and/or a gradual line temperature level. Temperature perturbations resulting
from the flow of propellant through the lines are readily distinguishaLle by: l) rapid change in line temperature, and 2) return of the lines to their nominal temperature level subsequent to engine shutdown. These generalizations are best demonstrated by data obtained during Mission A, as shown in Figure line Z. 5. i-4. Only lines 2 and 3 are shown, since line l is similar to
5.1-8
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a)
Vernier
Line
2
]
T !"
b) Figure 5. I-4. Vernier Line
Vernier
Line
3 at Midcourse, Mission A
Temperatures
5. 1-10
During Mission B, all vernier lines began to warm at completion of the premidcourse yaw maneuver. The rates of temperature change were Z, 3, and 3. 5 deg/min for three lines. At vernier engine ignition (Z64:05: 00:02), the line temperatures were 41, 44, and 51°F respectively (approximate values; see subsection 5. I. 4. 8). Vernier lines 1 and Z temperatures increased to 66 and 60°F, respectively, following vernier ignition. In each case, these values were 6°F below the postmidcourse temperature of the corresponding oxidizer tank. Vernier line 3 did not exhibit the rapid temperature change characteristic of normal propellant flow. Instead, the line 3 temperature decreased by Z degrees at midcourse. The positive temperature increases on lines 1 and 2 to values near the true oxidizer tank temperatures, along with the ignition of engines l and Z, indicate normal oxidizer flow to these engines. The failure of vernier line 3 to experience a similar increase strongly suggests that little or no oxidizer flowed into thrust chamber assembly 3 at midcourse. Table 5. I-3 summarizes the above data, with a comparison to corresponding SC-I values. Fuel Flow Arguments
A second proposed conclusion is that fuel flowed through the engine 3 fuel line _t midcourse. The propellant line temperature sensors are mounted on the oxidizer lines, and thus there is no direct thermal measure of fuel flow. Transit temperature profiles (sensors P-7, P-10, and P-ll) for the vernier engines are shown in Figures 5. 1-35, 5. 1-38, and 5. 1-39 for TCAs l, 2, and 3, respectively. Prior to the midcourse burn, the last recorded thermal data for the TCAs were acquired approximately 43 minutes before vernier engine ignition. Subsequent to engine ignition, approximately 34 minutes elapsed before TCA data was reacquired. TCA i and 2 temperatures were decreasing and TCA 3 temperature
was increasing when vernier engine thermal data was reacquired. The temperature of engine 3 was 71°F when TCA thermal data was obtained following the midcourse burn. A careful observation of Figure 5. I-5, which shows the first 6 minutes subsequent to the reacquisition of TCA data, indicates that the temperature of vernier engine 3 was increasing at the rate of 0. 5°F/ rain. A straight line interpolation of the data back to midcourse results in an engine temperature of 54°F at the time of vernier ignition. However, a straight line interpolation of the data presented in this figure is optimistic and unrealistic. It is reasonable to postulate that the temperature rate of change of TCA 3 is in excess of 0. 5 °F/rain subsequent to the rnidcourse vernier engine burn and prior to the acquisition of engine ther1_lal data. The straight line interpolation suggests a temperature change of 17°F for the interval bounded by engine ignition (264:05:00:02 GMT) and the acquisition of engine thermal data (264:05:34 GMT). turn vacas
The temperature o£ TCA 3 at initiation of the ll0. 6-degree yaw (264:04:48 GMT) was 64 °F. An examination of the SC-Z solar thermal uum test data indicates that the maximum temperature drop of TCA 3,
indicated by the flight sensor (P-ll), was 10 degrees in 12 minutes. Vernier engine 3 was not completely eclipsed prior to the initiation of engine thrusting; to the contrary, the engine was partially illuminated. Therefore, a maximum
5. l-ll
o
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I
5.1-12
TABLE
5. I-2.
SUMMARY
OF PROPELLANT (LONG BURNS
FLOW
INDICATIONS
Oxidizer Burn Number Fuel Flow (Less Than
Flow Normal)
Yes, Inferred 2 3 9 15 21 27 33 34 40 41 Inferred In fe r red Yes Yes Yes Yes Yes Yes Yes Yes No No No
Questionable
Yes Yes Yes Yes Yes Yes Yes determination determination determination
TABLE
5. i-3.
OXIDIZER
TANK MIDCO
AND URSE
LINE (° F)
TEMPERATURES
AROUND
SC-2
L
Temperatures Vernier Vernier lines before lines maneuvers
1 23 37-45 91
Data 2 16
SC3 24 51 49 24 24 91
1 Data 2
3 20-24 24 61
23-27 36 91
at ignition after ignition
41-47 9O
Vernier lines (uncorrected)
Vernier lines after (corrected for 4400 error) Oxidizer Oxidizer tank tank after before
ignition bits/sec
66
6O
49
69-71
59-61
61
ignition ignition
73 50
66 35
7O 46
67 58
59 44
68 53
5. 1-13
temperature decrease of i0 degrees during the interval bounded by the initiation of the yaw turn and vernier engine ignition is probably reasonable. Maximum temperature decrease conditions are synonymous with total solar eclipse conditions. Working from this argument, it can be hypothesized that the temperature of vernier engine 3 would not decrease below 54°F as a result of cooling due to partial shadowing and off-axis solar illumination. Certain inferences can be made if the following accepted:
i)
two propositions
are
than burn
TCA 3 temperature rate of change (increase) was greater 0. 50 °F/rain during the interval bounded by the midcourse and acquisition of engine thermal data a_ 05:34. TCA 3 could experience 10°F during the interval midcourse yaw maneuver
z)
a temperature decrease no greater than bounded by the initiation of the preand initiation of vernier ignition.
Statement l suggests a TCA temperature lower tha_ 54°F at midcourse, while the second statement suggests that a temperature lower than 54°F cannot be achieved as a result of the partial solar eclipse condition. Hence, the temperature of engine 3 will cooled by a superficial process. decrease below 54_'1c only if the engine is
engine
A cooling process can be considered in which fuel flows into vernier 3 and expands in the combustion chamber. During the expansion,
energy is extracted from the thrust chamber body, causing a decrease in the thrust chamber temperature level. Absence of engine thermal data during the rnidcourse burn does not permit verification of the suggested cooling. However, an evaluation of other TCA ignition attempts, where engine thermal data is available, strongly indicates the aforementioned cooling. Burns 2 and 3 infer some superficial cooling for TCA 3; however, it must be kept in mind that the oxidizer line temperature sensor indicates oxidizer flow during burns 2 and 3. An investigation of other burns, 34 and 40 for example, infers that the observed 'I-CA 3 temperature perturbations (increases) are the result of fuel flowing. Burns 34 and 40 indicate that the engine 3 thrust chamber temperature approaches the fuel temperature and then cools very rapidly. A cross sectional view of a vernier engine (Figure 5. i-6) illustrates that propellant enters the barrel through a fuel inlet n_anifold located tom of the thrust chan_ber barrel. The fuel then travels through groove along the thrust chan_ber wall and enters the combustion through the injector head. at the bota spiral chamber
A comparison of the engine 3 temperature profiles for the midcourse (burn l) and third burn are presented in Figure 5. i-5. The thrust chamber barrel thermal response curves are similar for the midcourse and third burn attempts in the regions where data are available. A generalization to
5. 1-14
MOLYBDENUM THROAT SILICON INSERT NOZZLE EXTE_ Ae/A t = 86: I
f
,*-': t i
CARBIDE
I
5.09 Dia.
SILICON
CAR_
RING
/
ROKIDE
•540 Dia.
I
Assembly
VORTEX INJECTOR OXIDIZER
Figure
J
INLET
5. I-6. Thrust
FUEL
Chamber and
INLET
Injector
5. 1-15
all subsequent burns indicates that the thermal response (warmup following the ignition atten_pt) of the engine is sin_ilar for burns 2 through 40. In fact, overlays of the long duration ignition attempts indicate that engine thermal behavior is repetitive. Because of the similarity and repetitive nature of the data obtained during burns 2 through 40 and the similarity that exists between the warmup transients for burns 1 and 3, Jt can be inferred that the same phenomena occurred during midcourse and subsequent ignition atten_pts. One important exception to this is that the thermal data do not positively indicate oxidizer flow during midcourse but suggest limited oxidizer flow in the engine 3 line during subsequent firing attempts.
The second proposition, which states that: fuel flowed through the engine 3 fuel line during midcourse burn, is primarily inferred from the similarity in the thermal response exhibited by TCA 3 following burn 1 and the response due to subsequent burns. Again, arguments regarding TCA 3 premidcourse temperature drop and postmidcourse rate of change (identical to those of the preceeding discussion) are required. A careful examination of the flight data (Figure 5. I-7) indicates that fuel flowed during all firing attempts subsequent to the midcourse burn. The fuel flow phenomena is verified by the increase or decrease in engine 3 thrust chamber barrel temperature during the engine operational interval and the period immediately following engine operation. A comparison of engine 3 thrust chamber barrel data obtained during the Zl. 5-second burn (burn 40) and engine data obtained during similar engine operational periods at the Edwards Test Station are shown in Figure 5. i-8. Visual inspection of the data indicates that the thermal behavior of the thrust chamber fuel flow periods. Curve C of Figure the engine and fuel thermal parameters The correlation between the analytical 5. 1.2. 3 Vernier Burn Thermal assembly barrels are similar during 5. i-8 is an analytical curve based on for the n_axin:un_ flow condition. ':-" and actual data is excellent.
Inconsistency
An investigation of the thermal data obtained for all 41 ignition attempts indicates that the thrust chamber assembly thermal behavior (engines 1, Z, and 3) was inconsistent. The data indicate that the thermal behavior of vernier engines i and 2 was different fur burn intervals of the same duration. With the exception of the midcourse burn where it is questionable whether oxidizer flowed to the engine, the thern_al behavior of engine 3 appears to be consistent for all burns. This is best illustrated by
*Curve
C of
Figure
5. 1-8
is
based
on
semiempirical
formulae flow. No gravitational criticism. attempt field.
for
heat was made There-
transfer in pipes and assumes fully developed to correct the formulae for the effects of the fore, the analytical curve is subject to valid
5.
1-16
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5. I-7.
SC-2
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3
5. 1-17
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5. 1-7
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SG-Z
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3
5. 1-18
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5. 1-19
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Figure
5. I-7
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SC-Z
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3
5. I-20
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5. 1-7
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3
5.
1-21
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5. 1-22
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SC-2
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5. 1-23
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5. 1-24
observing Burn II to the
the thermal behavior of TCA 1 during the of the 0. 25-second series exhibited thermal burn attempts for that engine.
0.25-second burn attempts. characteristics similar Likewise, burn 27 of the
2. 00-second
2.0-second burn attempts appears to be longer parison of the thermal data for TCA 2 indicates
in duration. A similar comthat the temperature rise resulting it is is basis. 5. 1-9
resulting from burns 2 and 3 was approximately 20°F; the increase from burns 9, 15, 21, 27, and 33 was approximately 60°F. Hence, concluded that the thermal behavior of the thrust chamber assemblies inconsistent The within above themselves anomalous even engine when viewed on an individual
engine
behavior
is demonstrated
in Figure
where the thrust ignition is shown
chamber barrel temperature as a function of burn number.
increase resulting from engine The data presented in this
figure is merely a quantitative measure of the energy released by the engines during the 2. 0-second burns. Energy release is synonymous with temperature change. While Figure 5. i-9 does not give any qualitative information regarding TCA thermal behavior, it does indicate, for example, that the thermal behavior of TCA 2 during burns 2 and 3 was substantially different from the thermal behavior of the engine during burns 9, 15, 21, 27, and 33. This is further illustrated by observing the flight data presented in Figure 5. i-I0 and extrapolating back to the peak temperatures for all2.0-secondburns. Similarly, burn II of the 0.25-second burns and burn 27 of the 2. 0-
second burns appear to be anomalous for TCA i. The engine total temperature rise during burn Ii is very similar in magnitude to the temperature rise exhibited by TCA i during the 2. 0-second burns. An inspection of Figure 5. I-II indicates that the thrust chamber barrel (engine i) temperature change (AT) increased by more than a factor of 3 for burn 27 when compared with the barrel temperature change for other 2. 0-second burns. An investigation of the command signals indicates that the engine was commanded on and off for a nominal data (Figure 5. 1-12) 2. 0-second burn. indicates that the However, a review of the strain gage engine operated at the minimum thrust
level (-80 milliamperes) as commanded for 2. 0 seconds. Strain gage data indicate that the engine probably did not shut off after 2. 0 seconds, but continued to burn at midthrust level for an additional 2. 50 seconds. TCA I thermal behavior ted temperature during burn 27 is not anomalous, rise is indeed explainable. and the larger than expec-
The data presented in Figure 5. i-9 also rise of the engine i barrel continued to increase second firing. One explanation for this behavior extrapolation and probably process that more logical determines inference
indicate that the temperature with each successive 2. 0would be an error in the temperature. resistance decreased Another to the flow as a result
the engine peak is that the thermal wall gradually
of heat through the combustion chamber of the successive firing attempts.
resulting
It is interesting to note and compare the temperature from burns 2 and 3 with those for burns 9, 15, 21, of irreversible data further for TCA
increases 27, and 33.
The 2
data suggest that some type between burns 3 and 9. The sistent, levels of operation
change took place within engine suggest two discreet, but yet conthe 2. 00-second burn attempts.
2 during
5. 1-25
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5. l-Z9
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5. 1-30
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Figure
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5.
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SC-Z
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5.
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5.
I-II
(continued).
SC-2
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5. 1-35
Figure
5. l-II
(continued).
SG-2
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5. 1-36
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Figure
5. I-II
(continued).
SC-2
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5. 1-37
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5.
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SC-Z 5. 1-38
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Igililii_,_!ili:: llii_!iiil!llilif!!l':r ? IfilitlililIil_tl_ltllitliiilil!lliiitttt_lt!_ ttit711_itti!t!il14_ttlti11!fitlltilii_ ! i i t tifill!lillilillliittii! Iiiitii_t!l_I!lP+!i Iiliit!tlIti1ilt_Ii_ iii Ift!ttft_titlllttll Ilttttittitl_i]i_]l I!I!INII!ilIIili!Iltl ili i! 7Ii ! illi[: _ lltlltiiiil!iilti!ilt!li i !tii i _+lf!l t!1ilt!f71t_fltitit_11 lillilit[_liitlttilitiI H,l,:td,t._:t_.t lh,llt:l,IT,th_, f
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t'+i_:_l'.t!',,'1,_i_+i ;:l_t111flt _l}tlr:l_l_l:til '_] : :ll:l:!J.'.!!!:!..'_!". _l;I;:111i:;tlb:_l_:4
L,:li!!ili, iIl
iiit!!ili7 gil_I_ Iili711i711I'.i !lil:. 1 i!I,ttitititttf_tt F_+tlII t ': ':+_llhii;ii_'_lil;i
LL,.171 ...................... ';'.L'//}.7:
!iiliitltf!litil_tll_ _i i '.!!t!f4}!t!t!:tilIi _i!_tiit
!_;l;;l;ll_l;It
_
_ 1 }i41ii:i
i;il]l_tilil_ii'.tiiil
......... ,-,,-,r.
i:':!i!i:!i!!L_t
NN
Figure 5. i-ii SC-2 VPS Thermal Response, Leg 1 5. 1-39
Figure
5. I-ii
(continued).
SC-2
VPS
Thermal
Response,
Leg
I
5. 1-40
J
oo !
D.,
v
0
-w--I
or4
5. 1-41
The type of phenomenon observed on this engine is probably what would be expected if the engine lost some of its physical mass or if the thermal resistance of the engine to the flow of heat from the combustion chamber through the combustion chamber walls was lessened significantly. Another possible explanation of the burn inconsistencies is suggested examination of test data obtained during vernier engine firings at the Test Station. Simulation tests to reproduce the thermal results
from an Edwards
observed during the flight indicate that the oxidizer/fuel ratio may vary for firings of the san_e duration. This phenomenon results in the propellants releasing different amounts of energy for firings of the same duration. Since the engine barrel temperature rise is directly related to the oxidizer/fuel ratio, one would expect variations in the barrel temperature with variations in the oxidizer/rue] ratio. Figure 5. 1-13 presents a quantitative measure of the energy released or absorbed by the propellant lines during the Z. 0-second burns. (Energy release is synonymous with temperature change, LT). While 5. 1-13 does not give any qualitative inforxnation regarding vernier line thermal behavior, it does indicate, for exan_ple, that the thern_al behavior of vernier line 2 was very consistent during the 2. 0-second burns. That is, the line temperature change is relatively constant for all Z. 0-second burns. This suggests that the quantity of oxidizer flow during all these burns was the same. Figure 5.1-14 shows the oxidizer line temperature sensor thern_al response as a function of propellant flow rates. Figure 5. 1-14 is based on send-empirical formulae. 5. i. Z. 4 Tun_bling Mode Thermal Observations (Failure Review Support)
An SC-Z _'tumbling n_ode" thermal analysis was performed in an attempt to detern_ine the location of the sun vector with respect to the vehicle during the postn_idcourse tumbling period. The analysis was based on the temperature changes noted on the 75 temperature sensors as the spacecraft orientation changed from that of normal transit to ti_e tumbling mode. Most temperatures attained a new equilibrium condition in the tumbling mode; however, some temperatures continued to change until the spacecraft was lost. The following indicators : list of thermal responses covers the most significant
I)
A
thern_al
switch
opened
on
each
con_partment, which radiators. could only This
and
the
radiator there that lower were the
temperature no solar sun vector above the hemisphere.
dropped to energy incident
a level on the
exist if indicates
never intersected x-y plane and must
the spacecraft have been
from incident
a direction from the
z)
The the
lower spaceframe sun was incident
warmed from below-
up considerably, the X-Y plane.
indicating
that
5. I-42
f
tit
-HI
Z fif
O
UI
o
b_ H_
c_ 0
i*4-
!
_d
N-h
•
_
g'_
:t:rt U:
_::
-=.
0) Ld
M
52
'ix
?t=
:i-c
5. 1-43
0
0
:ii
:i
ii_
HI
r/l 0
®_
_
0
--..4
_
0 0
• "lJ
0 0
!;! iii;
_i_
711 _" ] !!i _!i_ ii:
a_ 4a_
:i
_i_ L ¸
•._
...,-I
ffl
'<
.7_:
i_,_
_4
7i11 _t+
_0
T,
5,
1-44
3)
The
main the
retro
nozzle,
crushable considerably, the spacecraft. lines incident
block
heat indicating
shield, that
and-AMR the sun was
temperatures striking
increased bottom of engine that the
4) 5) 6)
All vernier indicating The SDC equilibrium,
propellant sun was dropped that
increased from the
in temperature, bottom of spacecraft. and never completely reached shaded.
temperature indicating
it
considerably was perhaps
The solar panel the sun striking 34 degrees off this temperature.
dropped the back normal.
to a temperature which would exist with side of the panel at an angle of Intermittent shading could also produce
7)
The planar array temperature increased to a value which would exist with the sun incident at an angle of 78 degrees off normal. Intermittent shading could cause the same temperature result. All three shock absorbers dropped to a temperature cates only partial solar illumination. which indi-
s) 9)
The KPSM increased in temperature, indicating that it was perhaps being illuminated from the bottom, since the bottom has a higher solar absorptance than the top. The auxiliary minium case battery temperature dropped, but the polished aluwhich is isolated from the battery reached a very indicating the sides plate. tubes that the sun was illuminating or bottom but not significantly the on the
lo)
high temperature, compartment from top prime radiator
ll) 12)
The upper be shaded
spaceframe continuously.
in the
vicinity
of leg
1 appeared
to
The -&/SPP solar axis dropped in temperature, indicating that it was being illuminated from a direction not normal to the solar axis or rotation. During normal transit, the sun is normal to the axis of rotation. The outboard face of the compartment A canister was receiving the same amount of solar heating during tumbling as it did during the normal transit attitude since the temperature never changed. This indicates that the solar load was equivalent to a 70-degree off normal incidence. The compartment B canister temperature dropped considerably; however, it appears that it was getting a small amount of solar illumination based on the equilibrium temperature.
13)
perature
Tumbling changes
about many discussed
of several above, and
different axes any reference
could produce to the incident
the temangle of
5. 1-45
the sun on a given surface is merely intended to Rive an equivalent heating orientation. The tumbling rate of the vehicle was such that no fluctuations were observed on any of the temperature sensors; therefore, no conclusions were drawn as to whether or not the various components were tumbling in and out of the sun to produce the observed equilibriun_ temperatures whereever partial illumination was apparent. Based on the steady-state temperatures after midcourse, a bounding of the most probable locations of the sun vector has been presented in Figure 5. 1-15. The figure indicates that the sun is coming from the lower hen_isphere and fron_ that quadrant formed by the _X, -Y, and +Z axes. No definite conclusions can be reached as to how closely the tumbling axis coincides with the sun vector. There was no definite indication of any continual change in tumbling axis throughout the tun_bling period; however, some small temperature shifts were noted on sonde items which could have been caused by the successive engine firings rather than a slow transient in the tumble axis. Perhaps one of the strongest conclusions that can be drawn from the thermal study is that the sun had to be coming from the lower hemisphere due to tile temperatures observed on the compart1_ent radiators after the thermal switches opened. 5. I. 3 SUMMARY AND CONCLUSIONS 5. I. 3. i Thermal
Performance SunqiT_ary
Thermal performance of the Surveyor II spacecraft was highly satisfactory. Prior to the attempted midcourse maneuver, 40 of the 75 temperature sensors indicated temperatures within ±5°F of their predicted values. The largest deviation between actual and predicted temperatures was 19=F on the noncritical spacecraft structure. No temperatures were outside their predicted tenlperature ranges prior to the postn_idcourse tumbling mode.
The thermal performance was very nearly the same as that of SC-I with the exception of those specific areas where thermal finish changes were incorporated to improve thermal performance. Thern_al finish changes were incorporated on the auxiliary battery, Canopus hood, elevation axis motor housing of the A/SPP, ten_perature changes respective objectives. Mission 5. I-4. and the helium tank. Each of these changes produced in accordance with predictions, thereby satisfying their
Table
B predicted and Most subsystems
actual thermal that normally
data reach
are presented equilibrium prior 5. 1-2 and
in conditions
during the 1_idcourse
transit coast mode maneuver. Data
had reached equilibrium are also present in Table
to the attempted for that time the equilibrium
period immediately conditions attained
following the attempted midcourse during the tumbling mode.
for
5. 1-46
COMPARTMENT
Y +
COMPA;TMENT
(3" OO
t-..=
oo ,,,o ! t_ I .-,j u1
-X
Y
ENI
AUXILIARY
BATTERY
VECTOR
LOCATIONS
-Z
X
_x
f
30 deg
v
Figure
5. 1-15.
Probable
Location
of Sun
Vector
During
Tumbling
5.
1-47
<
8_ E_ 0
I--I
L_
8o
L)
o o
Q
Z
P_
e
L)
u
o u
0
rn oo
r_
<
c_
I
M
<
_ m "lJ " , , _ I_, _., !_, P., L) L) L.) L.) L) L) 0 L) L) _e Z_,<
5. 1-48
_
_
_o
Illl,lltprtti,O
oo_oo
_
_o_
o_
n_ i 009
o
_-__ _
._
m
0 L)
!
...........
g_
_!ooooooooo
F_
JJJ ii ii
_
_
_
5. _-49
Temperatures observed during the postmidcourse tumbling mode indicate that the sun vector intersected the spacecraft from below the X-Y plane and between the +X and -Y axes. A comparison of equilibrium temperature data between SC-I and SC-2 is presented in Table 5. i-5. 5. 1.3. 2 Vernier Burn Thermal Data Summary
Thermal data indicate that TCA 3 did not ignite during the midcourse burn. The results of thermal control investigations to date have not produced any positive conclusions regarding the seemingly anomalous behavior of vernier engines. At most, studies have shown that the thermal behavior exhibited by the vernier Specific inconsistencies I) Vernier a) engines was are tabulated l burn burn. II is different from inconsistent below: during the 41 firing attempts.
engine
TCA i thermal behavior during that for any other 0.20-second
b)
TCA I thermal behavior during burn 27 is different from that for any other 2. 00-second burn. However, this was not a thermal problem, since the strain gages indicate that the engine probably burned for 4. 5 seconds instead of the commanded 2. 0 seconds. engine 2
2)
Vernier
a)
TCA 2 thermal behavior during burns 2 and 3 is substantially different from that for burns 9, 15, 21, 27, and 34. Temperature rise of the engine barrel, resulting from the 2. 0second burn, was approximately 20°F for firings 2 and 3 and approximately 60°F for the later firings. TCA 2 appeared engine 3 show any maneuver. positive TCA indications 3 thermal attempts. of ignition behavior to burn consistently at two temperature levels.
b) 3)
Vernier
a)
Thermal during always
data do not the midcourse seems
consistent
for all 41 burn
b)
Thermal data tend to support arguments that fuel flowed at midcourse but that there was little or no oxidizer flow; however, the data are not conclusive. Thermal data indicate both oxidizer and fuel flow during burns subsequent engine to mid3 ignition course; on any however, of the 41 the data do not indicate burn attempts.
5. 1-50
TABLE 5. 1-5. COMPARISON OF STEADY-STATE TEMPERATURES IN MISSION A AND MISSION B, PREMIDCOURSE
Actual Steady-State Temperature,
°F Operation Allowable Limits
Flight Vehicle
Sensor and
Location
by
Subsystem
Mission
A
Mission
B
mechanisms A V-15 V-16 D-13 D-14 EP-8 EP-34 V-20 V-Z5 V-47 V-17 V-18 V-19 70 93 68 68 97 123 42 44 35 9Z -85 66 74 94 71 73 99 118 31 Z8 34
Compartment Upper tray Lower tray Transmitter Transmitter
A B
14010 125/o 210/o 21o/0
125140
Main battery Battery charge regulator Radiators No. 5 No. 8 No. 2 Thermal shell inside Thermal shell outside Thermal switch No. 5 inside Compartment Upper Lower tray tray B
185/0
150/-300 150/-300 150/-300
9Z
-8Z
lZO/O
69
150/-300
Boost regulator Radiators No. 4 No. 1 No. 5 Thermal shell outside Thermal switch No. 4 inside Wiring Auxiliary Auxiliary compartment Landing gear assembly harness battery battery
V-21 V-ZZ EP-13 V-24 V -45 V -46 V-Z3 V-26 V-29 EP-Z6
93 98 115 67 73 66 -70 88 88 35
99
103 128 70 84 70 -7Z 93
lZ5/o
125/0
18510 150/-300
150/-300
150/-300
1Z5/0
91
64*
125/0
130/20
V-48
-Z
130/30
Leg 2 Crushable block Shock absorber No. I No. 2 No. 3
V-31 V-44 V-30 V-32 V-33
83 -6Z 84 72 82
74 -48 76 73 8Z
160/-140 160/-140
lZ51-zo 1251-2o 1251-2o
5.1-51
TABLE
5. i-5.
(continued)
Actual Steady-State Temperature, °F Operation Allowable Limits
Flight
Sensor
Location
by
Subsystem
Mission
A
Mission
B
Antenna/solar panel positioner mechanism Solar panel drive Elevation axis drive Solar cell array Planar array A/SPP mast Spaceframe Upper Near Near Lower Under Under Retro Leg Leg Leg Propulsion Vernier chamber No. No. No. 1 2 3 tanks 1 2 3 P-15 P- 13 P- 16 P- 5 P-6 P- 14 engine thrust assembly P-7 P-10 P-ll and substructure M- I0 M-12 EP-I2 M-8 V-34 6O l 109 -50 -84 45 -17 Iii -50 -88 165/-225 165/-225 165/-200 280/-280 160/-140
spaceframe leg leg 1 2 V-27 V-35 60 -79 53 -81 160/-140 160/-140
spaceframe compartment compartment points V- 37 V- 38 V- 39 B A V-28 V-36 48 -27 42 -24 1601-140 1601-140
attach l 2 3
39
-36 44
44 -32 44
160/-140 160/-140 160/-140
59
72
59
54 84 63
i25/2o t4o/2o 13o/2o
Propellant Oxidizer Fuel 1 Oxidizer Fuel Z Oxidizer Fuel 3 Propellant Leg Leg Leg Helium I 2 3 tank
75/41 t
76/52t 77/24t 75/34 t 79/40t 76/53t
76/50tt 77/57tt 75/35tt 83/47tt 75/46 tt 75/53tt
lOO/O lOO/O lOO/15 lOO/15
100/15
loo/o
lines P- 8 P-4 P- 9 P- 17 23 21 21 to 29 to 26 to 26 18 20 20 72 to to to 28 27 27
lOO/O
100/0
lOO/O
100/10
6O
5. 1-52
TABLE 5. i-5.
(continued)
Actual Steady-State Temperature,
°F Operation Allowable Limits
Flight Main
Sensor retro case case
Location
by
Subsystem
Mission
A
Mission
B
Upper Lower Nozzle Flight
P-3 P- 12 P-22
73/67 74/46 -124
t ? t
72/73?t 76/59?? -i18
70/40 70/Z5
control control electronics I 6 FC-44 FC-45 FC-47 FC-46 FC-54 FC-55 FC-71 FC-48 FC-70 90 124 78 170.* 175.* 180.* 79 45 90 137 85 175** 175 ...... 174:',-';',-" 8?. 4O 165/0
Flight
Chassis Chassis Canopus Roll Pitch Yaw Roll gyro gyro gyro
board board sensor
190/0
130/-20 185/175 185/170 185/170
actuator tank
200/0
115/-10
Nitrogen Radars RADVS KPSM SDC
88
86
t60/-5o
R-8 R-9 R-10 R-13
12
11
100/-zz
140/-18 112/-42 ii0/-20
56
22 33
63
14 20
VS preamplifier A/VS preamplifier Altitude marking radar
Electronics Antenna Edge dish of dish
R-7 R-6 R -27
14 to 16 -12 -185
18 -14 -191
lzo/-5
135/-20 2001-300
Television TV TV TV * Not at ...... Corrected ? Launch ttLaunch 3 mirror 3 ECU 4 steady state. for bit rate + 63 hours. + 15 hours. TV-17 TV-16 T-3 -120 -134 -124 -120 -128 -103
18o/-5o 5o/-2o 65/-2o
error.
5. 1-53
5. I. 3. 3
Recommendations discussion
from
Line
Heater 5. i. 2.3
Anon_aly indicated that failure of the
The
in subsection
vernier line heater to cycle was caused by epoxy contamination. The following recommendations are believed to effectively correct the problem with minimum impact on current spacecraft launch schedules:
i)
Increase thermal dissipation capability of the vernier engine line heater, which will result in a reduction in the line heater duty cycle as defined by this investigation.
2
z)
Prohibit attachment of wire harnesses with high faces to the lines in solar thermal vacuum tests dation of test results. Perform a black light inspection at Hughes after (but before curing) on the unit level, and remove epoxy that could contaminate the lines. Perform shipping a final black light inspection to Cape Kennedy. on the
emittance surto avoid invali-
3)
epoxy application any excess
4)
spacecraft
before
5. 1.4
ANALYSIS
DISCUSSION A and B
5. 1.4. 1
Compartments
Compartment A interior temperatures (sensors D-13, D-14, EP-8, EP-34, V-15, and V-16) are shown in Figures 5. I-[6, 5. 1-17, 5. 1-18, 5. 1-22, 5. 1-54, and 5. 1-55; external temperatures (sensors V-18, V-20, V-25, and V-47) are shown in Figures 5. 1-56, 5. 1-57, 5. 1-61, and 5. 1-71. Compartment B interior temperatures (sensors EP-13, V-21, and V-22) are shown in Figures 5. 1-20, 5. 1-58, and 5. 1-59; external temperatures (sensors V-Z4, V-45, and V-46) are shown in Figures 5. 1-60, 5. 1-69, and 5. 1-70. The thern_al tunnel internal temperature (sensor V-29) is shown in Figure 5. 1-64. Compartment 5°F higher system than those temperatures of SC-I for during the compartments mission A and were approxiB at the same
mately
time in the mission. No anomalies were observe@ during the normal transit period, and all compartment system temperatures correlated well with predictions. The seasonal change in the solar constant between the SC-I and SC-2 missions was sufficient to cause a maximum temperature increase of 3°F in compartment B. Solar thermal vacuum test data accumulated prior to the flight indicated that a 15°F temperature differential would exist between the SC-I and SC-2 temperatures in compartment B at the same solar intensity. Flight data did not support this evidence and, that those temperature differences observed between thermal vacuum tests were related to test operations simulation rather than vehicle differences. consequently, suggests SC-I and SC-2 solar or environmental
5. 1-54
During the tumbling mode, the compartments appeared to be in an orientation such that the radiators were not receiving any solar illumination. One thermal switch on each compartment opened after the attempted midcourse correction. Compartment A thermal switch 8 and compartment B thermal switch 5 opened at approximately 12 and i0 hours (28. 5 and 26. 5H
mission thermal 5. 1.4. 2 time), respectively, switches appeared Auxiliary Battery profile within (sensor EP-26) (Figure 3 ° F of predictions. The after the attempted n_idcourse correction. The to open within specified temperature tolerances.
5. 1-51)
The auxiliary battery temperature prior to midcourse maneuver was
temperature had not reached a steady-state value at this time. After midcourse, the battery temperature dropped at a rate of 4°F/hour as a result of vehicle misorientation. When the auxiliary battery temperature reached 34°F (at approximately L + 23H9M), auxiliary battery mode was commanded on in order to utilize the auxiliary battery power before the battery became too cold to function properly. Auxiliary battery mode remained on for approximately 9 hours and 46 minutes, whereupon main battery mode was restored. The auxiliary battery reached 79°F during the operational period. The auxiliary battery remained off for the next II hours and 52. minutes and declined to 28°F. At this time, RADVS was commanded on, and the magnitude of the electrical load caused a switching to the auxiliary battery. Although the auxiliary battery was well below the desired temperature of 95 ± 15°F at the time of RADVS turn on, it functioned nominally until it was commanded off during RADVS operation. The that of SC-I temperature iary battery temperature nonoperational performance 5. 1.4.3 SC-2 auxiliary battery case thermal design was modified from in an effort to increase the auxiliary battery transit equilibrium by approximately 30°F. Comparison of Missions A and B auxiltemperature profiles reveals that the Mission B auxiliary battery was 64°F and steady-state was excellent. and Solar close to steady-state as opposed to 35°F during for Surveyor I. Overall, auxiliary battery
Antenna
Panel
Positioner
(A/SPP)
The A/SPP mechanisms, solar panel, and planar array temperatures prior to the midcourse maneuver were at equilibrium temperatures within 10°F or less of preflight predictions as determined by flight sensors EP-12, M-8, and M-10 (Figures 5. 1-19, 5. 1-31, and 5. 1-32). Following the midcourse maneuver, the solar panel, planar array, mast, elevation axis motor, and solar panel stepping motor stabilized at -42, 48, -58, 2, and 9°F, respectively, within approximately 6 hours or less following midcourse maneuver. Comments about the A/SPP temperatures during the tumbling mode are contained in subsection 5. I. 2.
5. 1-55
5. 1.4.4
Spaceframe
Spaceframe temperatures (sensors V-27 and V-28) are presented in Figures 5. 1-62 and 5. 1-63. Spaceframe steady-state temperatures during coast phase I were about the same as during SC-I flight and were from 0 to 12°F lower than the predicted temperatures. Tables flight data, temperatures. 5. 1.4. 5 the 5. I-2 and 5. I-3 show the comparison for flight, and between SC-I and SC-2 flight
predicted
temperatures
postmidcourse
Landing
Gear Gear
and
Crushable
Blocks
Landing The approximately predicted increased 75°F. organic leg
2 steady-state temperature 90 minutes after launch and
of 72°F (sensor was 18 degrees
V-31) warmer
occurred than the
value of 54°F (Figure 5. 1-65). during the normal transit phase This increase can be attributed white paint on the leg.
The leg temperature gradually and, at L + 15H, had risen to degradation of the
to continued
Although the solar intensity during Mission A, the l_,_ission ]B leg temperature ture is attributed to the initially nondegraded
Mission B was higher than during was lower. This lower temperawhite paint on the legs due to testing. than leg 2
the protective wrapping used on SC-2 during solar thermal vacuum Legs I and 3 are not instrumented in flight, but should be warmer due to the absence of any shadowing caused by the solar panel. Shock Absorbers
Shock absorbers i, 2, and 3 ran 8, 9, and 2°F cooler, respectively, than the predicted steady-state temperatures (sensor V-32) as indicated in Figure 5. 1-66. SC-I and SC-2 shock absorber temperatures were about the same, with the exception of shock absorber 2 which ran approximately 8°F cooler on SC-2 even though the solar intensity during the flight was greater than that of SC-I. This could be caused by either the lower leg plate temperature or merely variations in the thermal finish of the shock absorber. In any event, the deviation tolerance of :1:25 °F. Crushable Block is well within the temperature uncertainty
The
crushable
block
heat
shield
steady-state
temperature launch and 5. 1-68).
of -48°F agreed very
(sensor V-44) occurred well with the predicted 5. 1.4. 6 Thrust Chamber
approximately temperature Assemblies
6 hours after of -51°F (Figure (TCA)
Vernier initiation of the
engine thermal premidcourse
performance yaw maneuver,
was as expected. vernier engines
Prior to l, 2, and
3
5. 1-56
were within the predicted temperature range. Predicted TCAs l, 2, and 3 were 65, 80, and 70°F, respectively.
at the initiation of the premidcourse yaw maneuver were for TCAs tures for for these I, 2, and 3, respectively. TCAs i and 3 were If and engines. An extrapolation
temperatures
for
Actual temperatures 54, 88, and 63 °F
The steady-state equilibrium tempera7°F lower than the nominal predictions of the actual flight data indicates that of approxof 80°F.
TCA 2 would have imately 85°F, or
reached a steady-state equilibrium temperature 5°F higher than the nominal predicted temperature
Thermal effects of the gyro drift check on TCA 2 temperatures (sensor P-10) can be seen in Figure 5. 1-38. TCA 2 reached a peak temperature of 93°F during the gyro drift check which was initiated at L ÷ 06H54M24S. A positive temperature perturbation indicated that TCA 2 received increase maneuver thereby increased solar illumination during the gyro drift check. An in solar illumination is experienced by TCA 2 during a positive because the shadow line cast by the solar panel shifts inboard, exposing more of the TCA to the sun. Peak because (Vernier TCA temperatures are not available for the midcourse burn interval. yaw
vernier engine
engine telemetry data was data is not sampled during temperature 5. 1-35,
not sampled during this telemetry mode i. ) P-7, P-10, for vernier
TCA transit presented in Figures and 3, respectively. 5. 1.4. 7 Propellant
profiles (sensors 5. 1-38, and 5. 1-39
and P-ll) are engines i, 2,
Tanks 5. 1-41 and predictions. indicated
Fuel tank temperatures (sensors P-13 and P-14) (Figures 5. 1-42) up to midcourse maneuvers were within 3°F of preflight During midcourse maneuvers, the fuel tank temperature sensors
increases of 17, 21, and 20°F on tanks i, 2, and 3, respectively. These increases are attributable to: I) mixing of the fuel within the tank, subsequently breaking up the isothermal stratification within the tanks and increasing the conduction film coefficients between the 2) flowing of fuel through the standpipe assembly fluid and tank to which the well, and sensor is
attached. Thus, the fuel which was warmer than the standpipe outlet imparted a temperature increase to the flight sensor. Spacecraft tumbling after the attempted midcourse maneuver produced mount and blanket temperatures higher than those observed during transit. These higher temperatures prevented the tanks from cooling as they normally do during this phase of the mission.
(Figures
Oxidizer tank temperatures (sensors P-6, P-16, and P-17) 5. 1-34, 5. 1-44, and 5. 1-45) up to midcourse maneuvers
were
within at least 6°F of preflight predictions. During midcourse maneuvers, the oxidizer tank temperature sensors indicated increases of 25, 33, and 26°F on tank i, 2, and 3, respectively. Following midcourse maneuver, the fuel tank temperatures remained at higher levels (66 to 75 ° F).
5. 1-57
The same explanation formulated above for the fuel tank temperature profiles also applies to the oxidizer tanks. It has been suggested that oxidizer tank 3 may have been empty at launch. However, an analytical investigation of the thermal response of oxidizer tank 3 indicates the presence of propellant before and after the attempted midcourse correction. 5. 1.4. 8
Propellant vernier Lines propellant lines for engines I and 3 behaved properly
The
during the transit phase. Thermal data from P-4 (Figure 5. 1-33) indicates that engine 2 line cycled during the early stages of the transit mission as expected; however, line temperature cycling terminated at approximately L + 3H. The line then gradually decreased in temperature. Prior to the premidcourse yaw maneuver, vernier 2 line temperature was 14°F. (Subsection 5. I. 3. i provides further discussion on the vernier engine 7oxidizer line. ) Subsequent to the completion of the premidcourse yaw maneuver, the
propellant lines exhibited a positive temperature increase. Oxidizer line temperatures were 48_':-% _?', 47 and 51°F for engines l, 2, and 3, respectively, at the initiation of the midcourse burn. Vernier oxidizer lines I and 2 exhibited large temperature perturbations as a result of the warm propellants (65 + 5°F) flowing through the cooler propellant lines (30 to 48°F); however, the thermal sensor on the engine 3 oxidizer line showed only a slight negative perturbation during the midcourse burn. Propellant line temperatures (sensors P-4, P-8, and P-9) are presented in Figures 5. 1-33, 5. 1-36, and 5. 1-37 for vernier oxidizer lines i, 2-, and 3, respectively. An examination of the data presented in Figure 5. 1-33 indicates that
the vernier engine 7-propellant Figure 5. 1-33 also illustrated time until the heater remained progressed, line within temperature
line heaters commenced to cycle at L + 90M. an increase in the heater on time with mission on at L + 3H. As the transit mission to maintain the oxidizer 2 (19 to ?.6°F), and the line by the data. aluminum may have
the engine 2 line heater was unable the cyclic dead band of the thermostat level gradually decreased as shown
Thermal analyses indicate that the exterior surface of the foil heater blanket surrounding the propellant line and line heaters
been contaminated by a high emittance substance. The emittance of clean uncontaminated aluminum foil is 0. 04 ± 0. 01. Calculations indicate that the emittance magnitude of the larger exterior than surface of the blanket was probably an order of normal.
":-'Value uncorrected temperature regions
for
4400
bits/sec
error. to be
The
actual and
line
i and
3
are
estimated
(37-45)
(41-47)°F.
5. 1-58
5. I. 4.9
Helium
Tank
Helium tank thermal performance was as expected. Prior to initiation of the premidcourse yaw maneuver, the transit steady-state equilibrium temperature was 72°F, or 3°F lower than the nominal prediction of 75°F. The thermal finish design for the SC-I and SC-2 helium pressurization tanks differed in the quantity of 3M black velvet paint on the inboard face of the tank. The 38°F black band spans the entire circumference of the SC-2 helium tank. The black band covered approximately three-fourths of the circumference of the SC-I tank with the inboard face painted white. The SC-I helium tank stabilized at 59°F during the Mission A coast phase. The transit in Figure 5. 1. 4. 10 The temperature 5. 1-45. Main main Retro retro Engine temperature profiles (sensors P-12 and P-22) shown profile (sensor P-17) of the SC-3 helium tank is shown
in Figures 5. 1-40 and 5. 1-46 were exactly as predicted for the upper and lower motor case and within 2°F of preflight predictions for the nozzle prior to midcourse maneuver. Following midcourse maneuver, the upper retro case continued to cool at a slightly higher rate than during normal transit attitude. The lower retro case temperature slowly increased (25 °F/hr) to 72°F, and The the nozzle temperature case and experienced the retro a 177°F rise. increased
lower
retro
nozzle
temperatures
due to the solar load impinging on these areas during the postmidcourse period. The retro case temperature reacts very slowly to changes in heat input due to its very high mass, whereas the retro nozzle, which has a much lower mass, reacts quite rapidly to sudden environmental changes. 5. i. 4. ll Flight Control Electronics and Canopus Sensor
The flight control electronics chassis boards, gyros, and Canopus sensor internal temperatures are presented in Table 5. i-6. These temperature results are for the steady-state coast phase. The actual temperatures are within the predicted accuracy of ±20°F (or ±2°F for the gyro temperatures). TABLE 5. I-6. Items Electronics Electronics Canopus Roll Pitch Yaw gyro gyro gyro board board l 6 FLIGHT CONTROL Sensors FC-44 FC-45 FC -47 TEMPERATURE, Predicted 100 138 Actual 90 138± 2 °F
89
177 174 177
85
176 174 176
FC-46 FC-54 FC-55
5. 1-59
In the mission plots for FC-46, FC-54, and FC-55 (Figures 5. 1-24, 5. 1-27, and 5. 1-28), the apparent discrepancy between flight data and predictions is actually due to telemetry errors in these "high accuracy" temperature channels. At the If00 bits/sec data rate which prevailed before midcourse, this error is 8 to 10°F. Due to the possibility of fogging of the Canopus window, the Canopus hood paint pattern was changed to increase the window temperature. Effect of the change was indicated by the Canopus sensor temperature (V-47) which rose from 78°F in Mission A to 85°F in Mission B. It is expected that the Canopus window temperature increased much more than the temperature sensor since the sensor is located on the electronics inside the unit and is somewhat removed from the Canopus sensor. Therefore, the objective of the change was accomplished. 5. I. 4. 12 Roll
Roll Actuator
actuator
thermal
performance
was
as
expected.
Prior
to initia-
tion of the premidcourse yaw maneuver, the roll actuator reached a steadystate equilibrium temperature of 8Z°F. The SC-I roll actuator also stabilized at 82°F. The nominal predicted temperature for the roll actuator was 88°F. The transit temperature profile (sensor FC-71) of the roll actuator is shown in Figure 5. 1-30. 5. i. 4. 13 The Nitrogen nitrogen Tank tank steady-state temperature prior to midcourse was
40°F as compared to 45°F for Mission A, and was 12°F below the predicted value. The tank remained within its operational limits throughout the mission. It was found that the tank temperature (sensor FC-48) dropped rapidly to 10°F during midcourse due to gas expansion in the valve resulting from spacecraft tumbling (Figure 5. 1-26). 5. 1. 4. 14 The occurred ature is Attitude gas jet Gas 2 Jets steady-state 3 hours warmer temperature after than of 87°F (sensor 5. 1-29). vahe of FC-70) This 77 °F. temper-
approximately about I0 degrees
launch (Figure the predicted than gas Although
flight
Gas jet 3 probably ran warmer sensor on jet 3 to confirm this.
jet 2, although there was jets 2 and 3 are attached
no to to on
their respective legs in a similar manner, the jet on leg 3 is probably warmer because the temperature of that leg is expected to be higher due lack of solar panel shading. Jet l was expected to be the warmest based a landing gear solar thermal vacuum test. 5. i. 4. 15 As 5. 1-51, all 10°F within RADVS evidenced RADVS of in Table 5. 1-1 and Figures R-8, and were 5. R-9, 1-48, R-10, essentially 5. 1-49, and at 5. R-13) 1-50,
and were
temperatures
(sensors predictions
premidcourse
equilibrium.
5.
1-60
However,
after
the midcourse
correction
was
initiated
and
the
spacecraft
subsequently went into a tumbling mode, RADVS temperatures changed considerably. The KPSM and preamplifier components achieved a steadystate temperature at approximately 25 to 30 hours after launch. The signal data converter temperature was still decreasing at the time of activation, which was 44. 8 hours after launch. Although the KPSM and preamplifier temperature differed considerably from no_rtinal values at postmidcourse equilibrium, these units remained within operational limits. Only the signal data converter exceeded operational or survival temperature boundaries, as shown in Table 5. I-2. Postmidcourse equilibrium temperatures of the KPSM, doppler sensor, and altimeter sensor were 41, 48, and 95°F, respectively. The signal data converter temperature went below its lower operation limit of -18°F and lower survival limit of -50°F at approximately L + 21. OH and L + 25. 5H, respectively, and was -85°F at the time the RADVS system was energized. The RADVS system was energized for a 10. 2-minute period at E + 44. 8H, and all components came on as evidenced by the temperature increase shown in Figures 5. 1-48, 5. 1-49, 5. 1-50, and 5. 1-51. The component temperature sensors indicated the following:
l) 2) 3) 4)
KPSM temperature of 9. 53 ° F/rain.
(R-8)
increased
from
31 to 129°F
at a rate
Signal data converter temperature ÷3°F at a rate of 8.64°F/rain. Doppler increased velocity sensor preamplifier from 38 to 53°F at a rate
(R-9)
increased
from
-86
to
temperature of I. 4 °F/min.
(R-10)
Altimeter/velocity increased from
90
sensor preamplifier temperature to 104°F at a rate of i. 24°F/rain.
(R-13)
Although the signal data converter was 35°F below its survival temperature limit, the radar system apparently remained operable. However, possible degradation to the signal data converter was not ascertainable since the radar system was not exercised in a descent maneuver. All that can be said is that the unit responded normally to the RADVS on command as evidenced by its rapid Marking temperature Radar rise. (AMR) profile was as predicted and was within prior to the midcourse maneuver (approxiof sensors R-7 and R-27 are shown in the AMR than nomiplatform
5. I. 4. 16
Altitude
The AMR unit temperature 5°F of premidcourse predictions mately L + 16H). Mission plots
Figures 5. 1-47 and 5. 1-52. Following the midcourse maneuver, unit reached steady-state temperatures which were much higher nal values for a properly stal_ilized spacecraft. The electronics
5. 1-61
ECU heat sink and the two antenna sensors equilibrated at approximately 78, 99, and 155°F, respectively, whereas nominal flight temperatures should be -12, -16, and -185°F, respectively. However, the AMR unit remained within transit operational limits throughout the mission despite spacecraft disorientation. Command 0730 (emergency AMR signal) was sent at 265:09:34:17 GMT. Telemetry indicated that sensor R-7 (AMR ECU heat sink) went from 97°F at 09:34:28 to a full-scale reading at 09:34:29. Thus, it appears that the AMR was expelled from the retro nozzle within this time interval, thereby verifying retro ignition. The SC-2 AMR heater duty cycle was i. 65 hr/cycle, whereas the SC-I duty cycle was I. 16 hr/cycle. However, the SC-2 duty cycle was determined using only a few cycles due to the shortness of the mission. The longer SC-2 duty cycle can be explained by a warmer retro nozzle than SC-I. 5. i. 4. 17 Television TV System
Approach
The approach TV temperature prior to the midcourse maneuver was -108°F, within 8°F of the preflight prediction. Following midcourse, its temperature increased 176°F to a steady-state value of 68°F, indicating that a considerable amount of solar energy was impinging on the unit during the postmidcourse tumbling mode. The approach TV electronics temperature (T-3) is presented in Figure 5. 1-53. Survey TV
The survey TV electronics temperature was as predicted. Prior to midcourse, the electronics temperature was -132°F, 7°F below prediction. Following midcourse, the electronics temperature increased to -48°F, indicating an increased solar energy load on the unit. The hood and mirror assembly was within 9°F of the predicted
steady- state temperature (-if3 °F) prior to midcourse maneuver, although the rate of cooldown was actually much greater than predicted (see Figure 5. 1-73). Following midcourse, the assembly temperature increased to -59°F, thereby indicating an increase in solar illumination. 5.1.5 REFERENCE B.N. Taylor, 20 "Temperature Launch Date, Predictions " IDC 2221. for 19/77, Surveyor Mission B 1966. for
i.
September
[3 September
5. 1-62
5.
1.6
ACKNOWLEDGEMENTS This go to l) section was the following: The entire coordinated by H. E. Knudson. Other acknowledge-
ments
Hughes made
Surveyor in the
Thermal of the
Control Mission
Section B
for
condata. the
tributions
analysis
thermal
z)
Dr. J. M. F. Vickers and D. L. Ayers, JPL, vernier system temperature plots discussed review analysis (subsection 5. i. 2. 2).
who prepared in the failure
3)
Jerry Lewis, Hughes thermal control, for his extensive work in analyzing the vernier propulsion system thermal data in connection with the failure review analysis (subsection 5. i. 2. 2).
5. 1-63
!i!! !!: i
:;!i llil
!ii_ !i _! !i!;
Figure
5. 1 -16.
Transmitter
A
Figure
5. 1-17.
Transmitter
B
5. 1164
tH
iil i :_iiii_il
........
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lli:il Itili!
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Figure
5. 1-18.
Main
Battery
T_ 7i_iL : iii'_'_ i_:iiil.... . ilii _i......
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ill il i !!!
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iii_i_iiili
Figure
5. 1-19.
Solar
Cell
Array
5. 1-65
Titl
Figure 5. 1-20.
Boost
Regulator
........................
• +1,: +++, i+11,, , ....... ...... .........
ilil
+,+i ++
iii ]:+i ii++ ILl:If iiii i!i! !![i !ii!]!i!! :::: !if: [TEi i!i!!ii:i
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,r ....
ii!i ,_;_,.: iii:liii:li:iiiii_
::_: :_l:li!i: ii!ii!ii! :i!i!]iil ...... _• : !ii;; i;lilJ!!
.4 _2Z+,L+_@_ttTTfTI:IIIfI:
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:::
tTt?
.:,_,_
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lift
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.........
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_
I ..................
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:i i!il i if iii
4_: ...... ; ; ;; 771; I ii iiib
7
,4i; I[i
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[_ ............ i . • ; ; ,,i_ i i i _i_
{i i
i;._ ; ; ;;;
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5 lii]i];i i i :;;;Jli; z.. ; ; : :ii::: ;;; I .... ;i i: _
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7
,-,I.
,-o.
_l.
e.
ill
IL
Ill.
_.
am.
II.
ILt'l I _'B [ _T
l lIE
(1'_)
II.
_.
Figure
5.1-25.
Canopus
Sensor
5. 1-68
Figure
5.1-26.
Nitrogen
Gas
Tank
Figure
5. 1-27.
PitchGyro
5. 1-69
i!::
Figure
5. 1-28.
iili
Figure
5. 1-29.
Attitude
Gas
Jet
2
5. 1-70
Figure
5. 1-30.
Roll
Actuator
Figure
5. 1-31.
Planar
Array
5. 1-71
t:< t
':::::i:: ]
"::: '
!!r T_
hHIIilllIil it.
_7,T
Ili
..... i:
i:!: :!!
i!!!i
tH
III Iii H_
;111111 ii iii;i:;:;::
Ji
n,i i,,
ttl
tli 111
>, i+l iii
tii
lil
"eL I.> !!i al. ,a L
Figure
5. 1-70.
Compartment
B Radiator
5
i:::
Figure
5. 1-71.
Compartment
A Radiator
2
5. 1-91
Figure
5. 1-72..
Auxiliary
Battery
Compartment
Figure
5. 1-73.
Survey
Camera
3 Mirror
5. 1-92.
5.Z
ELECTRICAL
POWER
SUBSYSTEM
5.2.1
INTRODUCTION
The electrical power (EP) subsystem generates, stores, converts, and controls electrical energy for distribution to other spacecraft subsystems. There are two sources for this energy: l) storage batteries, and 2) radiant energy converted directly to electrical energy used for system loads or battery charging. During transit, the primary source of power is radiant energy via the solar panels. Figure 5.2-I shows associated equipment groupings. The performance of the EP subsystem during the SC-2 flight was nominal as compared to test data and simulation analysis predictions. Subsequently, specific comparisons will be made in the body of this subsection. Regarding the and regulator efficiencies loads, comparison to considered. total system, various loads, solar panel input power, are calculated from flight data. Analysis of specific prediction, and explanation of discrepancies will be
for
easy
In Table 5.2-I, major reference to mission
events are presented with I) time from launch plots (subsection 5.2.4.2) and 2) time in GMT
for reference to various list information, i.e., commands and engineering data reduction system (EDRS) processed data. In general, the divisions of Table 5.2-i correspond to flight phases of importance to the EP subsystem; consequently, it may not correspond to flight phases in other subsections. Basically, the flight region is divided into times corresponding to significant changes in electrical loads. The time sequence 16.84 to 44.79 hours after launch was not further subdivided due to insufficient data. Load changes corresponding to these flight phases are partially illustrated by the regulated current (EP-14) and more completely by the battery discharge current (EP-9).
5.2.2
ANOMALY
DESCRIPTION mission, flight. no anomalies
were
(
Even though the SC-? flight was not a nominal detected in the electrical power system during and seeming due to ground annotated for misrepresentations data processing guidance and
Lack of information plots after midcourse are possible, plots have been
in the of scanty clarification.
mission data.
Where
5.2-I
22v IUNREGULATED BUS
_" O0 O0
._
CONTROL REGULATOR FUGHT
29., FLIGHT
EPI2 EP24 EP25
r
EPIO
J
.._
TRIP CIRCUIT OVERLOAD
I,
I I I
P CONTROL
I t'_ I
EP1 I_ 29v NON-ESSENTIAL
J
PANEL SOLAR
OPTIMUM CHARGE REGULATOR
22v
UNREGULATE p
MAIN POWER SWITCH
II
._I
DCDC ONVERTER --_-_" BOOST
VOLTAGE DROP ----
•
29v
ESSENTIAL
j
J MAIN BATTERY
AU×,L,ARY EP_ JE,B
BATTERY CONTROL EP31 EP32
1[
I
REOOLAT,ON
EP7 R27
I
_
AUXILIARY BATTERY
I
I
J I
I
I
___ERI,
I 1
R29
EPI4
22v II REGULATED RETURN
EP°/
R17 EP11 RI8 EPi6
I
I t
BOOST REGULATOR $285 EP4
_s_ EP_!
22v
UNREGULATED
RETURN
BATTERY CHARGE REGULATOR
Figure
5. 2-1.
Electrical
Power
Schematic
5.2-2
TABLE 5.2-1.
EVENTS AND TIMES, ELECTRICAL Total flight time -- 45.035 hours
From hours Inc r e merit 0.273 Launch
POWER
Time, GMT (day:hr :min:s ec )
From To From
Time Launch,
To
Comments to sun acquisition high high power power on
263:12:32:00
263:12:48:23
0
0.273
(transmitter 263:12:48:23 263:13:16:33 263:18:30:46 263:19:22:05 264:04:36:44 264:04:54:20 263:13:16:33 Z63:18:30:46 263:19:22:05 264:04:36:44 264:04:54:20 264:05:00:41 0.273 0.742 5.979 6.834 16. 078 16.371 0.742 5.979 6.834 16.078 16.371 16.477 0.469 5.237 0.855 9.244 0. Z93 0.106 Transmitter Coast Transmitter Coast Transmitter Midcourse transmitter thrust phase 264:05:00:41 264:05:23:02 264:05:23:02 265:09:19:57 16.477 16.849 16.849 44.798 0.372 27.949 Transmitter Many engine transmitter P_ADVS power
higi_ power
high maneuver
power
high and FC power on high power
starts and high power
265:09:
19:57
265:09:30:09
44.798
44.968
0.170
power on mode cycling
265:09:30:09
265:09:34:17
44.968
45.035
0.067
Retro sequence End of mission
5.2-3
5.Z. 3 SUMMARY AND CONCLUSIONS 5.Z. 3. 1
SC-Z Summary of the comparison subsystem. of flight data for
Table 5.2-2 presents a summary to test data for the electrical power Conclusion
5. Z. 3.2
Operation of the electrical power subsystenl was non_inal throughout the spacecraft's flight. Information detailing the unregulated current change during midcourse vernier correction and emergency vernier ignitions are presented in Table 5.Z-3. The various values of vernier burn at midcourse are associated with the various techniques of analyzing the flight data. The most probable value of this current change is 1725 milliamperes, where the expected value of the change in current is about 1670 milliamperes. The various techniques used in calculating this change in current varied from averaged data, unn_anipulated nonaveraged data, and reduced nonaveraged data analyses of the midcourse velocity correction to averaging of leading and trailing edge current jun_ps for all long vernier burns. Further considerations as to the uncertainties associated with the determination of the vernier engine Energy solenoid remaining valve current are of continuing and n_ain concern. is shown in
in the
auxiliary
batteries
Figure 5.Z-2. At midcourse, coincide. After midcourse,
nominally predicted and flight data practically the spacecraft tumbled, and practically no
energy was available from the solar panels; hence, the spacecraft was totally dependent on the batteries for energy. Toward the end of the flight when available energy was low, the RADVS power was turned on. During the time of RADVS turnon, the batteries were switched through various modes of operation, as noted in Table 5. Z-6 and Figure 5. Z-18. The low unregulated bus voltage (Figure 5.2-23) during RADVS power supports the prediction that the batteries, especially the auxiliary battery, were nearly depleted of energy. The main battery was able to supply the current load alone until the end of RADVS power on. When RADVS power was turned off, the main battery provided energy to the end of flight which occurred shortly thereafter. 5.2.4 ANALYSIS
The analysis considers six areas: mission telemetry plots, power loads and sources budget, comparison of flight loads and flight acceptance test (FAT) loads, cyclic loads, vernier engine solenoid power, and power mode cycling. 5.2.4.1 Mission Telemetry Plots
Figures 5.2-2 through 5.2-9 are selected mission plots which are pertinent to the electrical power subsystem. They represent the averaging of the analog signals over a time period corresponding to 30 telemetry frames. Consequently, due to the scale of these plots and data averaging, they give excellent information for consideration of trends in data flow. Many annotations have been made on these plots related to commands and ground data processing.
5.2-4
TABLE
5.Z-2.
SUMMARY
OF
RESULTS,
ELEC
TRICAL
POWER
Predicted, Item Boost OCR Solar OCR Battery Total Selected regulator efficiency panel output output energy energy energy loads 58.0 4- 3.4 watts 63.8 used used energy efficiency From 77.5 80 ,1400 1120 4578 5698 Flight percent percent + 80 + 64 4- 200 4- 225 w-hr w-hr w-hr w-hr 4770 5910 Data 75 75 percent percent 1440 1140 + + or
Specification, FAT (minimum) (minimum) w-hr w-hr 192 192 w-hr w-hr
Transmitter (average Transmitter (average FC power value) on/off, value) phase value)
B
high
voltage
watts
A high
voltage
55.2
4- 10.3
watts
63.2
watts
regulated
47.6
4- 3.6 watts
49.87
watts
(average FC thrust
power
on 31.6 9.2 + ± 8.7 0.9 4- 12.3 watts watts watts 33.65 I 0.34 550 watts watts watts
Regulated Unregulated RADVS Vernier power burns data data data burns on, unregulated
534.5
Midcourse-averaged Midcourse-unaveraged Midcour Average Vernier AMR Gyro se-unaveraged of line heater heater many 3 heater
39.2
4- 10 36.2
watts
36.7 36.7
watts watts watts watts milliamperes) milliamperes) watts
watts 33.6 watts watts watts 2.g 5.1 watts watts watts
28.4
to 42.0 1.9 4.5
36.7 36.7 (100 (230 11.0
10. 5 watts
5.2-5
iil _iit
I:_:I::H
iiiti_2il
i
Figure
5, 2-Z.
Z2-Volt
Unregulated
Bus
Figure
5. 2-3,
Unregulated
Output
Current
5.2-6
i;:il
1!i_;! ii;
i_l il i
!: ..... !-+' :i::
':
:_
----7
!Lil
_!i! !
i!L i:;: i_
!i: :!i i l:i i[i i
i!it!i
!i[i :i:!
![i! Li!!!I
.........:7
_t_r
i.iir1_r i.ii_l,i
Fi: !i:
:_iiiii:
_,,_ _ ........Iti!
t:_i1 :l: i;l l:=q ili. 'i!L:
Q4.
iii_ !iii
1:[i _:::
it:: :I:H
i'i
:
_.
Figure
5.2--4.
Boost
Regulator
Differential
Current
l: i ,,_.:
I;': 1i !i;
il i. i:i !i
i!!:Tii!t!!i
:;illi
!_;! i! !:
i!i!/iF
I:I"Fil i_
' ii !_ !
_!iilU!!i!
,o
Figure
5. 2-5.
Battery
Discharge
Current
5.2-7
Figure
5. 2-6.
Solar
Cell
Array
Volt:age
I
!
Figure
5. 2-7.
Solar
Cell
Array
Current
5.2-8
Figure
5.2-8.
Regulated
Output
Current
r- .....
TZTN:;t
!;_,',i:::lii; i11 il
'N t'_t •
r "t'tT:
, ...... _i :i!i:i
::
! !!_!!_di
: 1 i;I
:.tii21_
i
iiiilii!i+i!!_
!]ii]]i]i _!&
:l:t
:NO DA" A : !'.! :: ;
:::-
BAl-r
MI_,E
ON-
iil II!! !:
:i iii:
" :J 1_:1:[
..... :!i!!
i;!i i
:::i'd::q
ii,
Figure
5. 2-9.
Auxiliary
Battery
Voltage
5.2-9
5.Z.4.2
Power Energy
Loads Used
and
Sources
Budget
Figure 5. Z-l 0 presents the battery energy remaining as a function of time. Table 5.2.-3 gives the battery energy used during flight in approximately each mode of battery usage. Predicted battery energy remaining results from an updating of a portion of the SC-Z nonqinal mission energy prediction (Reference i). The energies used from flight data and predicted loads are almost identical until aIter midcourse. The flight data lacks the bumps representing transmitter high voltage on and n_idcourse maneuver which are correctly represented in the predict plots since the plot points are at large intervals (Table 5. Z-3). Reference 1 predicts that, at end of midcourse, total energy used is 1940 w-hr, and that energy out of the optimum charge regulator (OGR) is II40 w-hr. solar panel energy input of liE0 w-hr at end of midcourse. Power Figures Data 5.2-ii through 5. Z-17 present various are data): power parameters directly as This compares very favorably to a and a total energy usage of I_30 w-hr
calculated from EDRS flight data. The parameters from the following telemetry channels (averaged i) Z) 3) OCR Solar Boost efficiency panel =((EP-Z -I." EP-16)/(EP-10 I-" EP-ll = ((EP-I
calculated
_:`EP-II))
-':`I00
power
= EP-10 efficiency
regulator
-':`EP-[4)/((EP-7
+ EP-14)
-':_ EP-Z))``:`- i00 4) Shunt unbalance current = (EP-9 -(EP-4 5) 6) 7) Total loads = (EP-9 power power 5.2-3. + EP-16 + EP-16 + EP-14 _ EP-17) + EP-7) -
+ EP-17)
-':-" EP-2
Regulated Unregulated TABLE
= EP-I = EP-Z
-':_ EP-14 .I." EP-4 ENERGY USED Ma in Auxiliary Battery Energy Used, w-hr
BATTERY
Time From Launch, hours 0to 16.3 Battery Mode Main battery 16.3 to g3. l Main battery 23.1 to 32.9 Auxiliary battery (0317) Main battery#
Solar Panel Energy,
w-hr
Battery Energy Used,
w-hr
Battery Energy Used, w-hr 810
llZO
810
828
828
1ZZO
488
73Z
3Z.9
to 45.0
17Z0 Total 4578 power 5.2-10 is on.
1480 3606
240 972
-",-'Power
mode
cycling
while
RADVS
bO
°,-I
C
-i-I
E
:>..
E4
©
I
_4
_4
b.O
,,-i
[..q
5.2-11
Figure
5. 2-Ii.
Total
Loads
!I:i _II !!il
i!::iii!i_ !!!i!iii
_r;]i:;;iliiill _:i:i
"Ii_ "11
[i!7{!;i
i[:11117
iilj£_
JLi
Figure
5. 2-12.
Optimum
Charge
Regulator
Efficiency
5.2-12
Figure
5. 2-13.
Solar
Panel
Power
fillii!_
" T i
i/
:If l:i! :H ]:_I
J:_ iill
rl i:;l
i!i!ii I
!![i :T:TIH..'I !TiiltTHll
:ii:i:: I
77i__ i!l::
!iiiii!i_!iiit!i_f!
:;l:i:itii II: IP, ::iii4:_<,:I ;_: r,:
h
I:_,_: m_:iq!tliiq!
-4. -_. I. 8. II. IL il. II. Si. Silll _l_ I _1 II4[ Ifl_l
Figure
5. Z-14.
Boost
Regulator
Efficiency
5.2-13
_i,il
:i_ !ii
_i
!z
:i
:!i::ii:
i',Hl:l'.tlii',lll !!i!itiiiit,!H_iNtli]iJt!;ltl_lt[i!Lii; :::1: !/i:: til i/::i! !
_illt!t!ii!!!ll [::l!iiJiiiit!il:tt::ttiiHNt!4itliql_i:iti{!4 i'_,itili;/i!! ii_ti!i_
!i
I;HN !HIl:Lfi;
• H:_,!lili:Iii ilil;ittititt::!ttt!_:ti!!::ti_ti!t_!iit_!:: i_!i i_i!1!i:_ i_i !
!!_ !!! ii_i_ii i i_! ;_ i':i Htlii:iii! ':illi':!
_;;I::_T_I_,! 1[;!il ii!!i! ili!!
:i: :i
I:: "i ,: !1 ji ] P if! :I:I ]7 ]7 _7!U
:
j _
i till it _ ! 1 :i I : I ! :il! ti : +! :,,: i::, :_i_!l!il];|i]]_7
:iii :41: 7T-
!!!Ilil ljt;iP, l ,i
i
-3.
!
-L
+:i
ih_[!r+_7;
i:
iilll
i ili",:4:.ili[];i
_!:li!:i
:;]::!:!!:
-IL
a.
I.
_
iI
i_
li
il.
l_-
ill
it.
Figure
5. 2-15.
Shunt
Unbalance
Current
Figure
5. Z-16.
Unregulated.
Power
5. 2-14
,/
,...,
iii;
li i-ll ;:Ir
:;i
J"
.....
:: : :;::
::i q
-:i
_ii i!
;;
bli
ir ......... li
'_
:tl i:1; Iqr
t:III'xMTF =,H: H, V( ._0
:ti,,,.....
iii i
LAUNCH
" iii
i,i.
_:!i
:i1
r:_ :i; !i
:; :!
,_tlFtt:_l! ?/: ..... ...............
.... _-=U ' ':
i; !'
i
Figure
5. 2-17.
Power
Consumed
and
Loads
5.2-15
Figure 5.2-11 shows the total loads for the electrical power subsystem for the entire SC-Z flight. Total energy used during the flight can be estimated from this plot, and this estimate is recorded in Table 5.2-3. Figure 5.Z-IZ is a plot of the OCR efficiency. The average efficiency appears to be 80 percent. Figure 5.2-13 is a plot of solar panel power. This power is received for 15.9 hours of the SC-2 flight (16.2 - 0.3). This represents an energy input of approximately I120 w-hr (average solar panel power of 88.0 watts * OCR efficiency of 80 percent* 15.9 hours). After midcourse, the spacecraft tumbled, and no significant energy was received from the solar panel. Figure 5.Z-14 is a mission plot of boost regulator efficiency which is relatively constant at 77.5 percent. After midcourse, the telemetry data is sparse. Out of mode and no data, as well as bad data conditions, exist. Yet, after midcourse, the low data rate telemetry provides data for computation of the boost regulator efficiency which agrees with the 77.5 percent efficiency before midcour se. Figure 5.Z-15 shows the shunt unbalance current through midcourse. The current is generally biased at about +0.5Z ampere. This includes the EP-17 input. Figure 5.2-16 is a mission plot of the unregulated power. Transfer to spacecraft internal power is shown vividly at hour zero. Figure 5.Z-17 shows total power consumed, as well as the sum of the regulated and unregulated loads through the midcourse maneuver. Transmitter high voltage on conditions, vernier ignition at midcourse, and thrust phase power can be observed. Prelaunch power is also plotted.
Comments During l:l without on Load Sharing on condition, load sharing was assumed to auxiliary battery mode on, where the diode
-_
be
high current mode the diode. During
was between the main battery to be 3:Z (auxiliary to main). These energy 5.2-3. assumptions remaining
and unregulated bus, load sharing was assumed This is the same as for SC-I (Reference Z). construction of the plot of the calculation of the values bus voltage (EP-Z) at
battery Table
are reflected in the in Figure 5.Z-10 and from
in
It is estimated
the unregulated
the end of RADVS power on that there was less than i00 w-hr of energy remaining in the main battery 5 minutes before the end of the flight. This is reasonably close to the indicated remaining energy in Figure 5.Z-10 for the main battery. Tolerance on the remaining battery energy is of the same order of magnitude as the estimate of the remaining energy.
5.2-16
5.2.4.3
Comparison Comparison
of Flight
Loads
and
FAT
Loads and FAT-measured loads and large test results
of telemetry-measured
(Reference 3) will be made for selected units, various heaters, current drains. Specification values (Reference 4) and special (Reference 5) will also be used in comparison. Selected Equipment Loads
Results of comparing loads are presented in Table are as follows:
flight and test specification selected equipment 5.Z-4. The loads and equipments considered
I)
Transmitter Table 5.Z-4. Reference 3.
High FAT
Voltage On/Off. Data are data for the transmitters values are somewhat
presented is taken than the
in from the FAT regulated
Flight
lower with
values; however, flight power data
the tolerances associated bracket the FAT values. and Off. are well
2)
Flight Control Power On commands 0300 and 0311 (Reference 4) limits. Flight Control 07?.7 commands P.ADVS Power
The load changes due within specification
to
3)
Thrust Phase Power On. is within specification. On. Command 0637 applies
The
first of the
many
power
to the
P.ADVS.
The power consumed is close to that expected. Figure 5.2-18 (EP-17, radar and squib current) shows the current profile. The average value of EP-17 was about 28 amperes. It should be noted that P,ADVS power on occurs near the end of the SC-2 flight where completely 5.2.4.4 Cyclic Gyro Loads Heater contains compared gyro heater effects. to the altitude energy remaining exhausted. in the batteries was almost
The
The periodic loading that occurs in EP-4 gyro heaters have a short on-off cycle when and vernier line heaters. was examined. Figure ignition. The which compares Line Heaters
marking radar (AMR) nonaveraged telemetry
A graph of frame-by-frame 5.2-19 (EP-4)contains such
data prior to the midcourse approximately 0.5 ampere, AMP` and Vernier
average gyro heater load is favorably to the FAT data.
effects
Figure 5.2-20 are averaged
is an EDP.S plot of EP-4 at 20 rain/in. out in this plot. The cyclic load effects are apparent. A trace of vernier above EP-4 in order to show how
Gyro heater of the AMP,
and
vernier line 3 heaters (P-9) has been placed
line 3 temperature the middle frequency
5.2-17
TABLE
5.2-4.
SELECTED
EQUIPMENT
LOADS
Current, Command GM'I Command(s)* (day:hr :rain: se_) Flight Specific ation/'I est Flight Specification/Test 'l irne, milliamperes Power, watts
T ransmitte high R only R R R R R R voltage 0105 on) 0106 0110 0106 0107 0103 0107
r on/off (filament HV off on off on off on 263:18:28:59 263:18:30:46 263:19:22:13 264:04:36:44 264:05:23:02 264:07:19:16 265:02:42:21
(Reference
3)
(Reference
3)
170 1860 2140 1860 2000 1830 1990
± • ± i ± ± ±
10 40 40 180 330 300 200 2180 (Reference 4)
t
2200
4.9 53.9
± ±
0.3 1 1.2
63
.8
2200
61.7
+
1.2
63.8
53.9 2200
+
5.2
56.9 52.8
± +
9.5 8.6
63.8
57.7
±
5.8
63.2
Flight power R 0300
control on 264: 12:05:57 1590 ± 100
(Reference
4)
1720
46.2
±
2.9
49.87
Flight power R Flight thrust power R 0727 0311
control off 264:13:47:16 control phase on 264:04:54:20 1090 440 ± i 30 40 1640 ±
(Reference
4)
(Reference
4)
20
1720
49.1 4)
±
1.0
49.87 (Reference 4)
(Reference
1160 470
31.6±8.7
33.65 10.34
U
0727
9.2±0.9
RADVS U 0637
power
on 265:09:19:57 28130 (average) 4- 500
(Reference 29000
4) 534.5 + 12.3
(Reference 0.550
3)
Vernier
burns** 1865 1725 1350 to 1600 ± 509
(Reference 264:05:0_:02 264:05:00:02 264:05:00:02 t670
5) 39.2 m 10
(Reference 36.7
5)
U U U
0721 0721 0721
(EDRS) (SSP)
1670 1670
36.2 Z8.4 to 33.6
36.7 36.7 36.7
U
0721
(average)
Table Figure
5.25.2-19
5,
1998
(average)
1670
42.0
*R
=
regulated; values
U
= unregulated, result from different techniques.
*"'Different
5.2-18
J
) i
11 ilfi
;i
+
ii !i
]
J
...............
END
OF
VERNIE
]
I
_L
Figure
5.2-18.
Unregulated
Output
Current,
Midc
I
:::::::::
J
J
:i
Unregulated
Output
Current,
5.2-19
_
:
,i,
1!111:11 !I:
[[i
iiiJ
!lil!l_Ji!
i!i!ili!_i_:IItNH _IIIIIiIIi
t_Itt_ilr_
I_BSB
_i i__i
,inl,i
i
T r_
2Jl
i
ourse
III_NIIN_IIINI ;_t_:
_HH_
t14;_:
!!!!!_!![:!I!Y!!!!
HHH
t]ltig]t hllilillli]IH]il
l i i
.st
]
.PRi::CEDING p:_<_
.
oscillation in EP-4 is associated with the vernier line heater. Only the AMK and vernier line 3 heaters are cyclic at this time. The vernier line 3 heater uses approximately 92 milliamperes, and the AMR heater draws about Zl2 milliamperes. This agrees very favorably with test data, indicating that vernier line heater 3 should draw about 100 milliamperes and that the AMR heater should draw mate flight history 5.2.4.5 Vernier about 230 milliamperes. of which heaters are on, Solenoid Valve Figure off, or 5.2-21 cycling. shows an approxi-
Engine
Current
As part of the vernier engine failure study, a careful attempt was made to determine, by observing the change in the unregulated current telemetry (EP-4) and the battery discharge current telemetry (EP-9), the actual current drain required when the engines were turned on. This determination is clouded and made somewhat uncertain by the presence of the following interferring Other AMR, data on these same telemetry during channels: this time (gyro,
i)
cyclic heater loads changing and vernier line heaters}
z)
Noise-like effects of the roll actuator on the analog-to-digital converter Other undefined noise on these
saturated
signal
waveform
3)
channels.
These data can be examined in several different ways, each of which gives a different result that varies between a low of 1.36 amperes to a high of 2.38 amperes. These values, summarized in Table 5.2-4, were obtained as described below.
1}
Change in averaged 1°87 ± 0.51 amperes
EP-4
current
level
at midcourse
--
The average value (averaged over 30 samples or 7.5 seconds} before engine ignition was subtracted from the similar averaged value during engine ignition. This value has a large uncertainty due to the presence of the various cyclic heater loads, as well as the roll actuator effect.
z}
Change in unaveraged 1.73 amperes
EP-4
current
level
at midcourse
--
Obtained by comparing the unaveraged value of EP-4 before and after engine ignition when the gyro heaters are in the same condition, i.e., all off or one on. This value is only subject to the uncertainty of the roll actuator effect (see Figure 5.2-18).
3)
Change 1.35 to Obtained
in unaveraged 1.6 amperes by deleting
EP-4
current
level
at midcourse
--
gyro
heater
loads
from
the unaveraged
value
of EP-4. the leading
The higher value edge difference,
of current whereas
change is associated with the smaller value of current
5.2-21
.r4
>
.c
©
U 0 U_ U_
<
4--}
0
c_
I rxl
5.
2-22
0 0
,-_ 0
I
I'M
©
5.2-23
change is at termination initial and final values
of midcourse burn. The difference indicates that these are not the best
in
values, e.g., perhaps another load turned on at the end of midcourse. The various possibilities associated with the difference constitute 4) Average trailing in _nitial and final changes a continuing investigation. change edge of unaveraged changes EP-4 -2.0 in current values
and
EP-9
leading
and
current
amperes
Table 5.2-5, is a summary of all vernier firings with data for the long burns (9.85, 2.0, and 20 seconds). Figure 5.2-22 is a frequency distribution of the current change (AI) values. The data in this figure indicate that the most probable value of 2xI is 2.00 amperes. Figure 5.2-23 shows a plot of AI versus burn number. This scatter plot places most of the AIs between 1.7 and 2. I amperes, with no particular trend in the data. range in the values is partially due to the effects of the actuator, as well as other load effects. 5.2.4.6 Power RADVS Mode Power Cycling On The roll
Near the end of the SC-2 flight, RADVS power was turned on (265:09:19:57 to 265:09:30:09). During this time interval, the auxiliary battery control (ABC) was cycled through various modes of operation. Table 5.2-6 is a summary of this power mode cycling. Figure 5.2-18 (EP-17) and Figure 5.2-24 (EP-2) supplement this table. Interestingly, at the end of RADVS power main battery was carrying the electrical load. the discussion of load sharing, the main battery as little as 100 w-hr of energy are avaiIable. on (265:09:30:09), only As previously indicated is almost discharged-the under perhaps
In general, Figure 5.2-24 (EP-2) expresses the expected changes due to the various power modes. Especially noted are the initial automatic battery transfer at 265:09:22:16 when RADVS power was coming on and the attempt to switch to main battery mode (265:09:22:16) without disabling the battery transfer logic. Removal an increase of the of isolation about 0.5 diodes volt in (265:09:2.4:24) EP-2. and (265:09:27:27)
caused
5.2-24
TABLE
5.2-5.
VERNIER
BURNS
AI (EP-4), milliamperes Time, GMT Number 1 2 3 4 5 6 7 8 9 lO 11 12 13 14 15 16 17 18 19 2o 21 22 23 24 25 26 27 28 29 3o 31 32 33 34 35 36 37 38 39 40 {day:hr:min:sec) 264:05:00:02 264:07:28:25 264:07:50:03 264:19:44:59 264:20:07:05 264:20:35:20 264:20:55:06 264:21: 15:12 264:Z3:33:23 265:01:00:34 265:01:05:42 265:01:09:23 265:01:14:41 265:01:19:46 265:01:28: 265:02:01: 265:02:08: 265:02: 265:02: 265:02:26: 265:02:39: 265:03:17: 265:03:23: 265:03:29: 265:03:34: 265:03:39: 265:03:47: 265:04:17: 265:04:23: 265:04:29: 265:04:35: 265:04:41: 265:04:56: 265:05:43: 265:07:45: 265:07:46: 265:07:47: 265:07:48: 265:07:49: 265:08:05: 11 19 11 13: 34 19: 37 O6 14 24 53 O7 33 O7 56 31 53 51 34 2O 1Z 19 00 12 15 18 25 12 Burn Time, seconds 9.85 2.0 2.0 0.25 0.25 0.25 0.25 0.25 2.0 0.25 0.25 0.25 0.25 0.25 2.0 0.25 0.25 0.25 0.25 0.25 2.0 0.25 0.25 0.25 0.25 0.25 2.0 0.25 0.25 0.25 0.25 0.25 2.0 2.0 0.25 0.25 0.25 0.25 0.25 20.0 Bits/sec 4400 ii00 ii00 137.5 137.5 137.5 137.5 137.5 If00 137.5 137.5 137.5 137.5 137.5 Ii00 137.5 137.5 137.5 137.5 137.5 II00 137.5 137.5 137.5 137.5 137.5 ll00 137.5 137.5 137.5 137.5 137.5 If00 1100 II00 1100 1100 1100 1100 1100 2168 1718 1836 2109 1758 2109 2051 2071 2109 1797 2373 2373 2051 2325 Leading Edge (L) 1610 2480 2099 2373 Trailing Edge (T) 1310 2266 2080 2373
AI (EP-9), milliamperes Leading Edge (h) 1560 2369 1538 1758 Trailing Edge (T) 1610 2710 1709 2099
1929
2051
1758
2099
1929
1758
(7495)
2002
1855 1660
2344
(-681z)
1636
2099
5. 2-25
o
E
Z
0
o
L)
o
©
I
o
-_=
5. 2-26
LOAD
Figure
5. 2-23.
Radar
arid Squib
Current
(RADVS
Power
On)
Figure
5. 2-24.
22-Volt
Unregulated
Bus
(RADVS
Power
On)
C
;
_
POWER
OFF
_ _
,t
_,!f_
"i¸
_" _ :
TABLE
5.2-6.
POWER
MODE
CYCLING
Reference Command GMq (day:hr :rain: s ec) Command Title and Tin_e, Cotangents Figure Otherwise 5.2-18 (Apply Unless Indicated) to Figures 5.2-18 5.2-23, seconds an(]
0727 265:09:19:06
FC
thrust
phase
power
on.
Voltage to 20.4 battery
drop: volts mode drop: volts battery 5.2-23 current
dc
20.7 in
main
0637 265:09:19:57
RADVS
powe
r
on.
Voltage to main Figure RADVS steps. 19.8
20.37 dc only. shows drain initial in initially in
59
265:09:20:13
Automatic auxiliary
transfer battery to by have main battery immediately transfer
to mode. RADVS battery. mode by due to load
78
0320 265:09:22:16
Restore tenable logic.
main battery
battery transfer
mode,
Attempt carried Auxiliary restored automatic low voltage.
188
0322 265:09:23:46
High
current
mode
on.
Little voltage. iliary
change Already battery rise: volts without in parallel battery. drop: volts with
in
unregulated in aux-
288
mode. 18.96 in main isolation with 326
032O 265:09:24:24
Restore enable logic.
main battery
battery transfer
mode,
Voltage to 19.4 battery diodes auxiliary
dc
0317 265:09:24:54
Auxiliary on.
battery
mode
Voltage to 18.96 battery in parallel battery.
19.4 dc isolation with auxiliary in main diodes
357
03Z3 265:09:25:29
High
current
mode
off.
Voltage (18.88 auxiliary
dropping volts dc). battery dropping volts change rise: volts without in parallel battery. no apparent de). due
due Still mode. due No to
to in
load
389
0322 265:09:27:09
High
current
mode
on.
Voltage (18.76 ticular
to par-
load
489
0322. 509
0320 265:09:27:27
Restore mode, transfer
main enable logic.
battery battery
Voltage to 19.22 battery diodes auxiliary
18.73 dc in
main
isolation with
0321 Z65:09:27:43 0323 265:09:28:{)I
Disable logic. High
battery
transfer
Switch, on
affect
525
unregulated drop: volts without has entire battery in via to
current. 19.19 in main isolation load. not allowed automatic battery 542
current
mode
off.
Voltage to 17.5 battery diodes; Auxiliary to swilch transfer mode.
dc
auxiliary
0630 265:09:30:09
RADVS
power
off.
Voltage to 20.0 battery
rise: volts mode. dc
17. in
18 main
672
5.2-29
5.2.5
REFERENCES J.R. Oelschlaeger, Hughes Aircraft "SC-Z Company, Nominal Mission Energy Prediction," IDC ZZ92/71, 19 September 1966. Final 68Z2ZR, Report," October Volume 1966. II, Section 5.3,
l,
,
"Surveyor I Flight Performance Hughes Aircraft Company, SSD
•
"Surveyor A-ZI SC-Z Flight Acceptance Test, Final Test Phase Report, j'Volume If, Section Company, SSD 69190R, July 1966. J. Mundy, "System Specification Functional Event for Engineering No. Z395Z3, Revision C, l July Power and Payload," 1965.
Solar Thermal 4. 0, Hughes
Vacuum Aircraft
,
Thermal Hughes
Dissipations by Aircraft Company,
o
J.E. Mundy, "Results of Special Tests Performed on Investigate Vernier Engine Solenoid Value Current," Company, IDC 2Z94.2/50, 5 December 1966.
SC-4 to Hughes Aircraft
5.2.6
ACKNOWLEDGEMENTS W. McIntyre J. Berger L.M. B ronstein T.H. Mansfield S.F. McCormick J.E. S.A. M undy Volansky Technical coordinator and writer
Signal processing Midcour se data Vernier burn data Loads analysis Midcourse data Signal processing
5. 2-30
5. 3 RF DATA LINK SUBSYSTEM
5. 3. I
INTRODUCTION of
This section contains a summary and analysis of the performance the data link subsystem during Surveyor Mission B.
The data link subsystem consists of the transmitters, transponders, receivers, command decoders, and antennas. It is the function of this subsystem to: l) provide engineering data transmission from the spacecraft at bit rates compatible with specific mission phases, 2) provide analog data, such as that from television and strain gages, at signal levels high enough for proper discrimination, 3) provide phase coherent two-way doppler for tracking and orbit determination, and 4) provide command reception capability throughout the mission to allow for complete control of the spacecraft from the ground. A simplified block diagram of the communications subsystem is shown in Figure 5. 3-I. The pertinent are as follows: Unit
Receiver A Receiver B Transmitter Transmitter Command
subsystem units on the spacecraft during the mission Part Number
231900-3 231900-3 263220-4 263220-4 unit 232000-5 link subsystem index, etc.,
Serial
Number 15 16 15 II 3 parameters are not meas-
A B
decoder
such
as
Unlike most subsystems, individual data losses, threshold sensitivity, modulation
ured or individually determined from mission data. The composite effect of these parameters on the performance is measured as received signal power at the spacecraft and the tracking station (DSII _) and as telemetry and command error rates. Consequently, it is impossible to compare individual link parameters to specified performance criteria. The best that can be done is to compare measured signal levels to predicted levels, and telemetry quality and command capability to predicted capabilities. To further cloud the analysis, omnidirectional antenna gain is a major contributor to the
5.3-i
_,
°_ IIiI_ w_
I
!
A
*M_
AN_I_a_A NUIC_
_vl_#
p_o.lw
_
low
i.Iw,oe
n_
I
pM s_Jt
t__
ipiP_ |iiii ii_ oI_41DJ,
Figure
5.3-I.
Communications
Subsystem
Block
Diagram
5.3-2
uncertainty are difficult
in received to achieve
signal and,
in
levels. most
Accurate omni-gain cases, deviations from
measurements predictions
can
most likely be attributed to antenna gain uncertainty. Because of the problems outlined above, analysis of the data link subsystem performance will, in general, be a qualitative analysis of the performance of the entire subsystem rather than a quantitative assessment of the performance of the individual subsystem parameters. Equally as important as subsystem performance evaluation in this analysis is the qualitative assessment of the premission and real-time prediction techniques used during the mission, since future missions must rely on these techniques as guidelines during the real-time operation. In general, the RF data link single exception was the performance Consequently, the actual and predicted All other subsystem units performed subsystem of receiver performances very close performed as expected. B, which was degraded. were not in agreement. to the nominal predictions. telemetered, data is corretest data, this approThe
DSIF, lated preflight section priate
The data contained in this report consist of spacecraft and mission event time data. Where meaningful, the to and compared with equipment specifications, previous predictions, and in-flight analysis contains the following discussions subsection notation: Anomaly discussion caused Summary contains Discussion (subsection of the degraded receiver by the tumbling spacecraft. and Conclusions a summary of relative Performance following predictions. which are shown
Specifically, with the
5. 3.
Z) -- This B, as well
subsection as the
RF
contains effects
a
(subsection subsystem to
5. 3. performance
3)and
This with postflight 5. 3. 4)--
subsection conclusions analysis. This
and
recommendations Subsystem contains
performance (subsection
the
Analysis items:
"subsection
1) z) 3)
General vector
discussion of data, relative to omni-gain of subsystem
equations contours.
used,
and
path
of the
earth
Discussion phases.
performance
during
specific
mission
Discussion of pertinent subsystem function of time from launch.
telemetry
signals
plotted
as
a
The major mission event times relative to the RF data link subsystem are tabulated in Tables 5. 3-1 and 5. 3-2. Table 5. 3-1 contains telemetry mode and bit rate, primary tracking station number, and station automatic gain controI (AGC) values as a function of time for the pretumbling and posttumbling phases. Table 5. 3-2 ration as a function of time for some cases, the times in these contains a tabulation of the subsystem configuthe pretumbling and post-turnbling phases. In tables are accurate only to the nearest minute.
5.3-3
TABLE
5. 3-i.
TELEME
TRY
MODE
SUMMARY
day:hr:nlin:
o T, I
sec 12:55:07 13:04:59 13:15:00 13:17:08 i3:26:29
Mode
[
Rate
i os,F
Station Pretumbling
DSIF
AGC, Phase Liftoff
dbm
Comments
263:12:31:59.824
5 5 5
550 550 550 51 51 5l
Low
modulation
index
-118. 90. 90.0
0 0
InJlial l,*,o
acquisition nay lock (SCM)
(SAA)
5 5 1 1 4 2 3 5
550 550 550 1100 1100 1100 1100 ll00
High Low
power power
] 1
Preparation transit I phase
for
-t11.6
13:29:26 13:3Z:51 13:34:50 13:37:37 13:39:24 i6:38:38
51 51 51 51 51 51 -137.1
>4orlrtal
bit
rate
selection
5 5
137. 137,5
5
E%i; tale
reduction
for
D5S-72
track
16:51:35 17:45:02
72 51
-146.2
I,SS I)S5-51
71
in in
two-way two-way
lock lock
5 5 5 4 Z i 5
137.5 137.5 1100 1100 riO0 1100 1100
17:52:02 18:01:26
51 51 51
-138.5 -i_5.6 l_it r_te increase for DSS-51 track
18:09:41 i8:13:25 i8:20:15 18:24:35 18:30:46
51
5 5 5 5 5
1100 1100 1100 17. 17.2 Z
51 51 51 51
-114.8 -112.0 -132.9
High
power
-
pre-Canopus
19:21:00 19:22:05 2t:50:06 22:02:00
}liRh
power
I
star
lock
I
i!lt rdte reduction for DSS-72 track 8 D,qS-7?. J)SS DSS }%it 5[ [I rate in set rise increase for DSS-II track two-way lock
72
-139,
22:18:00 22:50:50 23:i2:10 17.2 1100 It [l ii -148.0 -138. 2
Z3:21:40 23:24:18
1100 1100
I ransmitte
r off
]
DSS-
1 l having
11 11 1l [ rar;srmitter [ x_(,-way lock on / transmitter trouble
23:29:32 23:40:31 23:44:45 23:47:46 264:01:23:40
1100 it00 tl00 ll00 1100
11 11 blal't 11 -138.1 - 138.4 !Qarl En< receiver roduction receive r B B of test test power
01:40:07 02:14:00 02:54:44 02:59:37 03:02:28
1100 1100 i100 1100 1100
DSS,-]
1
having
transmitter
Irouble
5.3-4
Table 5. 3-I (Continued)
GMT, da y:hr:min:sec 03:04:08 03:05:41 03:07:42 03:13:18 04:14:00 Mode 1 5 Bit Rate 1100 1100 I)SIF Station II 11 DSIF AGC, dbm Comments
Gy 5 4 2 1 1 1 1 1
ro
I 100 1100 1100 II00 1t00 1100 4400 4400 4400
l 1 11 11 1l 11
Gyro
speed
check
04:15:51 04:18:10 04:36:43 04:37:53 04:51:50 05:00:02
I 1 11 11 11 -123. -123. 3 3
High Bit End
power rate
increase
premidcourse for :nidcourse
prenlidcourse thrust execution
Midcourse
Post-tumbling
Phase
264:05:00:00 05:20:50 05:23:16
1 1 1 1
4400 550 550 137. 5
11 11
-123. =-130.
3 0
Start
nonstandard
phase
11 11 11
-[35 -143
to to
-140 -144
Low
power
05:g9:20 05:31:45
2 2 5 5 5 5
137.5 137.5 137.5 137. 137. 13"7. 5 5 5
05:34:39 05:48:51 05:58:33 07:05:43
11 i1 4;'
-142
to
-153 Prior to 5 x station bit Iransler error rate to 42
-142
to
-152
4.
10 -3
4Z 42 4Z
-142
Io
-151 0
Spin High
period, power
1.
2
seconds
07:19:16 07:21:08 07:a2:Z0
= -ld0.
5 1 1
11 O0 1100 550
42. 4Z
2-second
thrusting
07:29:53 07:30:37
5 5 5 5
55(1 137. 137. 137.5 5 5
42 42
-117
to
-126.
5
07:34:04 07:35:06 07:46:36 07:47:31
4Z 42 42
-142 =-120
to
-151
Low lIigh
power power
5 1 I 5 5
1100 1100 550 550 137.5
07:47:58 07:51:37 07:53:02 07:57:54 07:58:04
42 42 42 4Z
2-second
thrusting
5 4 4 2 5 4
137.5 137. 137.5 137. 137.5 137.5 5 5
42 42
Low
power
10:21:05 10:25:00
42 42
-142
to
-152
Spin
period,
1. 06
seconds
10:29:58 i0:38:23 11:41:34
42 4Z
-147
n_ean
5.3-5
Table 5. 3-i (Continue d)
GMT, day:hr:min: 11:52:14 sec Mode Z Bit Rate I_7. 5 DSIF Station 42 1)SIN AGC, dbm Co n t nle nt s
IZ:04:£8 12:15:30
5 5
137. 137.
5 5
42 42 -145 to -160
[ ransn_itting a_!te nna A
on
omnidirectional
1Z:
15:30
5
137.
5
42
-145
to
-155
i rarlsl_litting a n'_e nna i_
on
omnidirectional
13:22:14 13:37:08
4 2
1 37. 137.
5 6
42 42
13:41:25 15:30:56 15:36:02 15:47:32
5 [ 1 1
1t7. 137. 1100 137.
5 5
42 51 51 =-126
thgh
power
5
51
15:49:00 15:50:22 16:00:00 18:00:00 19:35:18
i 5 5 5 1
137. 137. 137. 137. 137.
5 5 5 5 5
51 51 51 51 51 -144 to -148
l_) v,,
power
-144.
5
to
-147.
5 0, Z second thrusting
19:50:06 Z0:02:12
5 1
137. 137.
5 5
51 5i !). g second thrusting
20:09:50 20:28:32 20:37;28 20:46:24
5 l 5 1
i 37. 137. 137. 137.
5 5 5 5
51 51 5t 51 0, 2 second thrusting
i). 2
second
thrusting
20:56:27 21:10:51 21:16:53 22:06;30 22:30:00
5 1 v 5 5
137. 137, 137. 137. 137.
5 5 _ 5 5
51 51 51 61 61 -147. 0 I_o-way lock ',) Z second thrusting
22:55:50 23:19:58 23:ZZ:17 23:23:25 23:28:08
5 5 5 1 l
137. 137. 4400 4400 1 t00
5 5
51 11 il 11 11
t_i High
lransn_itter power
off;
51
two-way
lock
g - aecund
thrusting
23:34:24 Z3:38:46 23:40:17
5 4 5
11 O0 1100 1100
11 11 I l
23:40:4Z 23:43:31 265:00:59:21 01:01:12 01:04:38 01:06:16
5 5 1 5 1 5
137. 137. 137. 137. I37. 137,
5 5 5 5 5 5
11 11 11 11 11 11 Luw I?, Z power second thrusting
0.
2-second
thrusting
5.3-6
Table
5. 3-1
(Continued)
GM2, day:hr:min:sec 01:08:Z3 01:09:54 01:13:44 Mode l 5 1
Bit Rate 137. 137. 137. 5 5 5
DSIF Station 11 il 1l 11 DSIF AGC, dbm Comments
0.
2-second
thrusting
0.
2-second
thrusting
01:15:11 01:18:21 01:20:20 01:2t:36
5 l 5 5
137. 137. 137.5 137.5
5 5
11 11 11 11 ii ll 11 11
0.
2-second
thrusting
ttigh
power
01:24:22 01:25:37
5 1
1100 1100
Z-second
thrusting
01:29:09 01:29:38 01:30:24 01:39:53 01:44:36
5 5 5 4 5
1100 137. 137. 137. 137. 5 5 5 5
Low
power
11 11 11
01:59:51 02:01:58 02:06:53 02:08:40 02:12:38
i 5 1 5 1
137. 137. 137. 137. 137.
5 5 5 5 5
0.
2
second
thrusting
11 11 11 11 11 0. 2-second thrusting
0.
2
second
thrusting
02:14:01 02:18:34 02:20:07
5 I 5
137.5 137.5 137. 5
11 11 11 11
-143
to
-152
0.
Z
second
thrusting
02:g4:g7 02:26:35 02:34:00 02:35:36 02:36:2Z 02:40:08 02:41:16 02:42:21 03:12:40 03:17:56
1 5 5 5 1 5 5 5 1 5
137. 137. 137. 1100 1100 1100 137.5 137.5 137. 137.
5 5 5
0.
2-second
thrusting
It 11 11 11 It
High
power
2.
0-second
thrusting
It 5 5 11 11 11 ll 11
Low 0.
power thrusting
2-second
03:Z2:39 03:24:23
i 5
137, 137.
5 5
0.
Z-second
thrusting
03:Z8:06 03:29:36 03:33:48
1 5 1
137. 137. 137.5
5 5
0.
2
second
thrusting
11 11 11
0.
Z-second
thrusting
03:35:02 03:38:21 03:39:31 03:43:46
5 1 5 5
137.5 137. 137. 137. 5 5 5
11 11 ll
0.
2
second
thrusling
High
power
5.3-7
Table
5. 3-I (Continued)
GMT, day:hr:min: 03:44:44 sec Mode 5
Bit Rate 1100
DSIF Station II DSIF AGC, dbm Comments
03:45:51 03:48:50 03:50:19 04:16:14 04:17:56
1 5 5 i 5
1100 1100 137. 137. 137, 5 5 5
11 11
2.
0-second
thrusting
11
L_
puwer
1 1 11
O.
2-second
thrusting
04:Z2:4Z 04:24:12 04:28:Z6 04:30:11 04:34:38
i 5 1 5 I
137, 137, 137. 137. 137.
5 5 5 5 5
i 1 11
l).
2- s,,_
ond
thrusting
1 1 i1
l).
2- _pcond
thrusting
11 11
I).
2- _econd
thrusting
04:35:52 04:40:22 04:41:37 04:52:36
5 I 5 5
137. 137. 137, 137.
5 5 5 5
11 11
0
2-second
thrusting
11 11
High
powe
r
04:53:28 04:54:26 04:56:49 04:57:55
5 1 5 5
1100 1100 1100 137, 5
11 11 11
2.
0
second
thrusting
04:58:19 05:32:53
5 5
137. 137.
5 5
11 11 11 11 -t
i.ov, } l_gh 2.
p_,,_e powe ond
r r thrusting
05:35:52 05:45:20 05:46:34 05:47:10
6 5 5 5
1100 1100 137. 137. 5 5
O-_e<
11 ii
I_) v,. lmwer High power
06:32:45 06:34:38 06:50:40 06:53:54 06:58:41 06:59:12 07:31:00
5 5 5 5 5 5 5
137. ii00 137. 1100 137. 137.5 137.5
5
42 42
5
4Z 42
5
42 42 42 42
Lov, _-150, 0
power
07:41:49 07:42:50 07:43:26 07:50:02
5 5 i 5
137.5 II00 1100 1100
High
power
42 42 42
l;ive
I).
Z-second
thrusts
07:50:54 07:51:17 07;54:21
5 5 5
137. 1_7, 137.
5 5 5
4Z 4Z 42 42
Lcv, -147 to -152 [
power
08:00:52 08:02:11
5 5
137,5 1100
l{igh
power
42
5.3-8
Table
5. 3-i
(Continued)
GMT, day:hr:znin:sec 08:02:28 08:07:56 08:09:02 08:09:19 09:11:50 09:12:34 blode 6 5 5 5 5
Bit Rate It00 II00 137, 137. 137.5 5 5
DSIF Station DSIV AGC, dbm CoI_ln_ents
42 42 42
21,B-second
thrusting
42 42 42 42
Low tligh
power power
5 2
1100 II00
09:32:19 09:34:17
Z 2
llO0 I1O0
42
Emergency
AMR
command
09:35:00
42
-lg3
to
-128
Abrupt
loss
of
signal
5.3-9
o
C
.%
_
o C_ •<
:
Cl
o o _ UCl
_
<
m
<
z
0
I-.I
o
o
0
H
,<
,.-1 0
L)
.._
Z 0 O
_ 0
U U
i-,
.<
,-1 0
I o
m
,-1
"H<
0
t_ tr_ _
o_
-_
_
_o:
oo
.-I. _m
5.3-10
.r-4
o %)
v
I
_4
5.3-11
_D O
!
_9
.4
c_
5.3-12
5. 3. 2
ANOMALY Degraded
DESCRIPTION Receiver B Performance
5. 3. 2. 1
The only subsystem anomaly observed during the mission was the threshold degradation of receiver B which was most apparent during the first 16 hours of flight. A comprehensive review of test data by systems engineering taken at AFETR (Reference 1) revealed that receiver B had similar problems prior to the prelaunch countdown, apparently masked by RF air link variations. Postflight analysis of the flight data and postflight tests on other spacecraft receivers led to the final conclusion that receiver had become degraded, probably prior to the countdown. For completeness, a brief history of events relating to this anomaly will be given prior to an analysis of the pertinent flight data. After the gantry was removed was had during the countdown, receiver at receiver A countdown, at AFETR, consulting the B AGC A. a
B
indicated that the signal level Since no change of this nature possibie anomaly was suspected. spacecraft/performance/analysis/command, analysis change, team since led to it was
about 25 db below the level been noted during Mission Discussion among analysts and spacecraft that when a multipath the gantry effect was
the conclusion first reported
had caused removed.
After launch and initial spacecraft acquisition at Johannesburg, the signal level at receiver B was still 18 db below the level at receiver A. Since the spacecraft roll attitude was unknown prior to Canopus acquisition, the difference in signal level was not immediately considered a problem. However, during this pre-Canopus acquisition period, a comparison of spacecraft receiver signal levels, DSIF signal levels, and corresponding omnidirectional gains indicated that no earth vector position could be found which satisfied the observed conditions. Six hours after data taken during launch, Canopus the 360-degree acquisition was roll was compared initiated. Receiver B to that from antenna
AGC
gain patterns. They agreed relatively well with the expected variations. However, the absolute values were about 16 db below the expected values. This data indicated that the antenna patterns were correct, and that either the receiver AGC characteristics had changed or a loss of 16 db existed between the diplexer and receiver B. In order to investigate the anomaly and to determine the two-way (transponder mode) capability for the midcourse maneuver, a special threshold test was run at 01:g4 GMT. The DSIF transmitter power was lowered in 2-db steps, and AGC telemetry from receivers A and B was recorded. The point of observed receiver/decoder indexing was also noted. This test indicated that: l) expected gain variations for the proposed midcourse maneuver were less than 24 db and, hence, the maneuver could be made in the transponder mode, level changes 2) receiver A closely agreed AGC with calibration data was nearly correct known changes in DSIF transmitter (signal power),
5. 3-13
and 3) receiver B AGC calibration was not correct, showing excessive changes in receiver signal levels. (A change of Z db at the transmitter caused a 3- or 4-db change at receiver B.) The fact that the observed receiver/decoder index could have been caused by either receiver did not allow a direct assessment of whether receiver B was degraded or had merely shifted AGC calibration. A planned postmidcourse test to determine if receiver B was degraded was eliminated after the mission became nonstandard. Data from the special threshold test run 13 hours after launch is tabulated in Table 5. 3-3. Figure 5. 3-2 shows the receiver A in-flight calibration data (separate curves for index caused by receiver A or B) compared with preflight calibration data for temperatures of 75 and 125 °F. The flight temperature was close to 90 °F, indicating that the proper calibration curve should lie between the latter two curves. As mentioned before, the receiver that caused the index during this test was not known, but by assuming each receiver in turn and then comparing the AGC curve generated with the preflight curves, a reasonable conclusion can be reached. First, assume receiver A caused the index. From prelaunch test data, that index point was -122 dbm. This can be used to tie down the relative test data from Table 5. 3-3 to the absolute dbm scale of Figure 5. 3-2. The curve thus generated lies outside either of the preflight curves and would require a further assumption that receiver A had a 3- to 4-db error in its AGC calibration. Next, assume receiver B caused the index. In this case, the telemetered signal level for receiver A at the start of the threshold test can be used as an absolute value. Thus, the second curve of Figure 5. 3-Z was constructed. This curve lies between the two preflight curves and, in fact, indicates crossovers very near those shown in the preflight data. A deviation from the preflight curves does exist at levels below -i14 dbm, but the overall close agreement with preflight data leads to the conclusion that receiver B did cause the index, and was therefore degraded. calibration curve was generated. with the preflight calibration
This data
Based on that assumption, a revised data is shown in Figure 5. 3-3, compared for 75 and 125 °F.
Special tests were subsequently run by systems engineering on a spacecraft receiver to determine if a failure mode could be found which would shift the receiver AOC characteristics and degrade the threshold performance as had been observed. The tests indicated that such a failure could be duplicated by simulating a loss in gain in either the A6 or A8 modules. Figure 5. 3-4 shows the special test data taken when simulating losses of 3- and 6-db in the A6 or A8 modules. These modules have caused problems in the past and, in fact, receiver degradation due to them was noted on SC-1 during solar thermal vacuum tests and on SC-3 during vibration tests. In addition, these special tests revealed that there was no obvious failure that would just shift the AGC without also causing degraded performance (Reference I).
5.3-14
TABLE GMT, day (hr:min:sec) Start 01:37:03 01:39:48 01:42:04 01:44:21 0 1:47:15 01:49:23 01:51:28 01:53:30 01:56:12 01:57:22. 02:00:27 02:04:33
5.
3-3. 264
SUMMARY
OF
RE(;EIVER
THRESHOLD
TEST
DATA
DSS-11 Transmitter Attenuation, db 0 -2: -4 -6 -8 10 12 14 16 18 20 22 24;,'-"
Receiver 207 224 242 259 278 299 318 336 355 3?5 388 401 410
A,
BCD
Receiver 215 234 255 279 301 321 338 353 363 371 376
B,
BCD
test
379 381
_.-'Decoder
index
indicated
from
spacecraft
telemetry.
Based on the evaluation of flight data and the special receiver test, it is concluded that receiver B was degraded and probably had become degraded before launch. On the strength of these conclusions, all telemetered AGC flight data in this report are analyzed using the calibration curves contained in Figures 5. 3-2 (assuming B-caused index) and 5. 3-3. If the mission had proceeded successfully beyond midcourse, this would not have been catastrophic. Extrapolation of flight data, 16-db degradation in receiver B threshold, would still have in a positive command margin of l to 2 db at lunar distances.
anomaly assuming resulted
5. 3-15
o
L)
0
o
_o
-,-4
©
0
;>
cD
I
Lf_
"-
5.3-16
0
4--)
o
L_
,,"4
©
E
0
<
>
CD
_4
I e_
M
73^_
7
9
5. 3-17
_D ¢,)
_J3
0
¢.) °,-_ ,
5.3-18
5. 3. 2. 2
RF
and
Data
Link
Problems
Associated
With
the
Tumbling
Spacecraft
The anomaly that caused the spacecraft to tumble during midcourse thrust and eventually resulted in mission failure was not in any way due to the RF subsystem. However, once the failure occurred, the performance of the data link was substandard. Under the circumstances, the RF subsystem relative performed as expected, to a normal mission. though the resulting link was substandard
The link degradation due to tumbling resulted telemetry bit rates and an increased bit error rate. 137. 5 bits/sec was available for low power and If00 transmitter operation.
in lower allowable A maximum bit rate of bits/sec for high power
The
telemetry
quality
was
apparently
degraded
by
two
separate
effects.
First, telemetry signal to noise ratio (SNR) was changing as the spacecraft tumbled, resulting in below threshold SNRs during some periods after midcourse. Reported DSIF signal levels during high power operation were cycling between -117 to -127 dbm right after midcourse and between -127 to -132 dbm near the end of the mission. The nominal threshold signal level for if00 bits/sec was -138 dbm, indicating that the levels were well above threshold. For low power operation, however, the reported DSIF signal levels were cycling between -135 to -140 dbm after midcourse and between -145 to -155 dbm later on in the mission. The nominal threshold for 137. 5 bits/sec was part a result -152 dbm, indicating, in this of below threshold SNRs. case, that bad telemetry was in
The second degradation effect was less obvious, causing bad data during periods of high power operation or during periods of low power operation when the reported signal levels were above threshold. During these periods, word errors occurred in a periodic manner at the spacecraft tumble rate. (A more detailed discussion of data quality can be found in Section 5. 4, signal processing. ) Correct telemetry discrimination and decommutation require that the DSIF receiver be phase coherent, or phase locked, to the spacecraft transmitted carrier. Momentary loss of phase lock will, in general, result in transients in the data stream or short periods of bad data. Phase lock is maintained as long as errors in the tracking loop remain within +90 degrees. The tumbling spacecraft resulted in excessive tracking loop errors which could have caused periodic bad data.
The
primary
loop
errors
are
phase
jitter due
to noise
and
error
caused by the sinusoidal carrier modulation resulting from the spinning omnidirectional antenna. Figure 5.3-5 shows the primary loop errors as a function of a single omnidirectional rotation. At the top of the figure, the omnidirectional antenna is shown in four positions relative to the DSIF station. The typical omnidirectional antenna gain pattern is shown at each position with the and G 4. relative gain in the direction of the DSIF station as G I, GZ, G3,
5.3-19
#J
®
@
1
0 tt , I!, i
_." _, _i.'_i,
f_
Figure 5. 3-5. Carrier Tracking Loop Error
5. 3-20
The The
gain goes through a maximum phase jitter due to receiver position. The expression
and noise
minimum is shown
value on every rotation. as a function of omnidirecto noise is given by
tional
for the RMS
jitter due
ejitter whe r e N S = = receiver received noise carrier
{RMS)
:
V_
(1)
power power related to transmitter antenna gain,
Since received Equation I can
carrier power is directly be expressed as
ejitter where G(t) K The phase directional antenna = = omnidirectional lumped constant
(t)
=
KG(t)
(a)
antenna gain link parameters spin is also shown as spinning motion of the the W c transmitted W c s A sin W st carrier a function of omniomnidirectional given by
error due to spacecraft antenna position. The a doppler shift of
causes
AW D
-
where W W A = = = carrier spin frequency frequency (rad/sec} (rad/sec) of the omnidirectional head in the
c s
maximum direction velocity of
displacement of the station light error
c The carrier
=
tracking tracked by
loop
resulting lock loop
from
the by
existence
of this
modula-
tion being
the phase
is given
AWW esteady state B Z o
ZWZs ZBZ(1
{Ws S/N::= (Wst) +\Bo/
4
]
5.3-21
where AW Bo = = = maximum frequency shift due to spin factors for threshold and
loop natural frequency ratio of limiter voltage actual signal level
at threshold suppression
Since the RMS jitter due to noise will ride on top of the it is clear that at position 4 (Figure 5. 3-5) the peak loop error and, at this point, the loop could momentarily lose lock. Figure starting 5.3-6 has plots of DSIF receiver AGC at 7 seconds after midcourse (05:00:19
sine error, is maximum
error
and dynamic GMT). As
phase can be
noted,
the time of the decrease in receiver AGC corresponds to the time when large noise spikes occurred in the dynamic phase error. Also, the noise spikes occurred on the negative peak of the sine wave, as predicted. Figure 5. 3-7 shows the dynamic phase error at DSIF-42 during retro ignition. The loop error shows the effect described above and, in fact, right after ignition loss of lock can be seen on almost every negative peak of the sine wave. (Loss of lock occurs when the peak goes to the outer limit of the grid. ) The loop bandwidth (Bo) at DSIF-4Z had been modified prior to this time to accommodate the tumbling spacecraft and was approximately two times wider than the other DSIF station bandwidths. These data clearly show that the proposed problem did exist, even with a wider loop bandwidth. It is thus concluded that this mechanism also caused periodic bad data throughout the tumbling phase of the mission. 5. 3. 3 Summary and Conclusions
Table 5. 3-4 contains a summary of the measurable performance parameters compared with applicable requirements and premission predictions. Most subsystem parameters are not directly measurable, and those that are measurable are difficult to summarize due to time variability. Received signal level, for example, is a function of time and spacecraft attitude. The summary for these parameters reflects wide tolerances, with corresponding wide variations in actual performance, in cases when the earth vector was in the omnidirectional antenna null. Performance and predictions outside the null are much more closely bounded. More detailed information The analysis: is found following in the subsections can be dealing drawn with each mission phase.
conclusions
as a result
of the foregoing
i)
RF subsystem receiver B. experienced
performed as expected with the exception In most cases, close to nominal performance in both the up- and downlinks.
of was
5. 3-22
oo
=
Figure 5. 3-6.
o5" ,',_o ,'/G
Automatic
d,,'*l 7"
Gain Control and Dynamlc Phase Error at Midcourse
DSIF
5. 3-23
v
O3 Go
_fl
1
2
t/_illlll/111
I
3
IJl 1_tt f_Ji 1 1 1 !
!
J_J i
9 10
I
_ v_,,,'v,vrtl
I
4
0
1
5 IGNITION 6 7 8
,/"_ /\ I1,IX /q /_ tl /_ I1 R t_, /!, / fl J,,/tg ,[I l/tf 1 1
I0 I 12
A
Jl
_"l_l ,J ,_ 1d ,,t 1 1 I
13 14 15
IIJ 1
16 17 18 19 20
1
20 21 22
11
23
1
24 25
"_
_
_
']"
I
26
I
27
l
28
I
29 30
T
3O 31
1
32 33 34
1
35 36 37 38 LOSS OF PHASE LOCK
NOTE: VERTICAL LINES ARE TIMING TICKS LOCATED ONE SECOND APART. NUMBERS ARE INCLUDED ONLY TO INDICATE SEQUENCE AND HAVE NO TIME MEANING.
Figure
5. 3-7.
DSIF-42
Dynamic
Phase
Error
Variations
During
Retro
Ignition
5. 3-24
TABLE
5. 3-4.
PERFORMANCE
PARAMETER
SUMMARY
Parameter
Predicted
Value
Requirement
Actual
Performance
Transmitter at acquisition
frequency
2295.
001694
mc
2295
rnc
•
23
kc
2294. one-way
999779
mc
(5
seconds
after
acquisition)
Receiver at acquisition
B
frequency
2113.
309168
mc
2113.
31
mc
±
21
kc
2113. acquisition)
318944
mc
(at
two-way
Receiver levels phases
A during
signal coast
_'ime
variable Presome value
>-
114
dbm
,'_
Level nominal
between and
Z and -_ -95 dbna
4
db
above
predictions. dicts are nominal ± 12 db.
Receiver levels maneuver
A during
signal** star
Time predictions. dicts
variable Preare some value
> - 114
dbm*
Level db about
between nominal
417. and
0
and
- 13.
0 dbm
> -116
nominal • 10 db.
Receiver levels phases during
B
signal coast
Time
variable Pre some value -
>-
114
dbm,::
Level above
between nominal
+1. and
0
and
-3.
6 dbm
db
predictions. dicta are nominal • 7db.
-_ - 107
Receiver levels maneuver during
B
signaiV,r_ star
Time
variable Pre some value db. -
>-
114
dhm
',_
Level nominal
between and >
6
and
-7 dbm
db
about
predictions. dicts are nominal • 10
- 112
DSIF during
signal coast
levels phases
Time
variable Presome value db.
>-136.7 (carrier (17.2 threshold)
dbm power) bits/sec
Level of 1100 nominal
bet_veen and
+0. >
5 -139
and dbm
-Z.
5 at
db
predictions. dicts are nominal ±8
bits/sec
DSIF during maneuver
signal star
ievels_:*
Time
variable Presome value db.
None
Level nominal
between and
+4 > - 150
and dbn_
- 13
db
of
predictions. dicts are nominal • i0
DSIF during maneuver
signal midcour
levels se
Time
variable Presome value
>-135.4 (carrier (at 4400 high power)
dbm power) bits/see
Level nominal -
between and
+1 > -124
and dbm
-3.
0
db
of
predictions. dicts are nominal •3db.
Transmitter power Transmitter power output output
A
high
40.
6
_0. -0.
3 05
dbm
> 39.
6
dbn_
No
data
A
low
ZI.
09
+0. 21 -i, 19
dbm
> 19.
1
dbna
No
data
Transmitter power Transmitter power output output
B
high
40.
6 +0. I dhm -0.1
39.
6
dbm
Output
bet_,een
40.
6
and
40.
0
dbm
B
low
21.
1
+1.2 -0.2
dbm
> 19.
1
dbm
Output
between
19.
8
and
19.
2
dbm
Phase bandwidth
jitter
12
cps
<
36
degrees
<36
degrees
(3_)
No
data
Phase bandwidth phase)
jitter
152 (thrust
cps
<22
degrees
<22
degrees
(3rr)
Jittcr prior
< to
4.
0
degr,'es
(3 thrust
_)
midcourse
Command
reject
rate
<1/2000
_- 1/2000 level >
at 114
signal dbm
No at
rejected signal levels
commands _ -95
in dbm
125
sent
Telemetry rate
hit
error
<3/1000
-3 5 3/1000 SNR __ 11 at db input Minimuna input SNR BF.R 10 = • 2, 0. 8 7 x db 10 at
'::Threshold above _:=X:The in gain star -114
value dbrn maneuver exists.
applies at any caused
to one
command time. the earth
threshold
and,
as
such,
only
requires
one
of
the
two
receivers
to
be
w_ctor
to
pass
through
deep
antenna
nulls
where
the
greatest
uncertainty
5.3-Z5
w
2)
Performance of receiver B was not as predicted. The telemetered AGC was grossly in error, requiring a complete in-flight recalibration. Postmission analysis of pertinent data and special tests indicate that the receiver was degraded by approximately 16 db. Although operational problems would have resulted, this degradation would not have aborted the mission had it continued to the terminal descent phase.
3)
New omnidirectional antenna pattern measurement data, taken on the JPL range, was quite accurate in the regions viewed during the mission. Very good agreement was noted where gain levels were above -10 db, with lesser but still surprisingly good agreement at -15 to -20 db. Omnidirectional antenna A uplink patterns (Zll3 mc) were noted to be less in agreement with measured measured data. This was expected with a dipole angle which since the patterns was different than were that of
SC-2. It is concluded that of the patterns to positional antennas.
these data tolerances
demonstrate the sensitivity of the omnidirectional
4)
RF subsystem premission predictions and real-time analysis techniques used during Mission B \vere relatively accurate and, in most cases, were conservative. IRF link performance was good during the tumbling phase of the
S)
mission for both telemetry and command links. Data quality was substandard relative to a normal mission, but still adequate. The fact that a two-way (transponder) link was maintained with a degraded receiver oscillation on the system performed. The I) following (receiver B) and \vith l-second signal level is a measure of how doppler well the
recommendations
are
made:
Both Missions A and B had problems with receiver AGC telemetry. Considering this, it is strongly recommended
that a
system calibration be made during S'YV tests and that all applicable prelaunch tests run at AFETR clearly check for AGC changes. This information is not only required for postmission analysis, or partial but also failures may help flag any impending receiver leading to degraded performance. should be placed on failures
2)
Temperature
transducers
the transmitter
and
receiver modules that contain the respective VCXOs. It is very difficult, if not impossible, to correlate unit temperature data to any single presently telemetered temp_rature. The ability to check prelaunch frequency reports and to update DSIF tracking predictions is severely lessened because of this lack of correlation between frequencies and telemetered temperature data.
5. 3-Z6
3)
Recovery of DSIF station data for use in postmission analysis is not being done correctly. Although much data was received for SC-Z, there was an almost complete lack of all the calibration data needed to translate oscillograph deflections back to physical parameters (i.e., dbm) at the DSIF station. Many pieces of calibration information were provided, but never enough to determine the final curve in absolute engineering units.
5. 3.4
SUBSYSTEM General
PERFORMANCE Discussion
ANALYSIS
5. 3.4. I
mission included
Before specific phases are discussed, a general treatment will be undertaken. Information applicable to all mission in this subsection. Subsystem Most Parameters estimates of performance are based on
of the phases
is
quantitative
received
signal levels which, in turn, are determined from individual link parameters. Those parameters used in the performance predictions and the subsystem analyses are tabulated in Table 5.3-5. Equations using these data are derived here; parameters discussed in later portions can be evaluated from these data. Tables 5.3-5 and 5.3-6 consist of measured data taken from flight acceptance (FAT), solar thermal vacuum (STV), and command and datahandling console not available. (CDC) tests or specification values where measurements were
Computations In this
Used reference is made The to received equations signal used are levels listed and below
subsection,
quantities computed from these and will not be derived again: I) Spacecraft transmitter
levels.
high
power
output
is
Pxmtr(dbm)
= 10 log
(Ptm
x 10 3 ) + L
where
Pxmtr P
= transmitter = telemetered
power
(dbm)
= Phig h (watts)
tm
power
output
L
: loss from determined
transmitter to power monitor _- 1.5 from pre-STV hardline calibration
db (value data)
5. 3-27
TABLE 5.3-5.
UPLINK PARAMETERS FROM FAT, STV, AND CDC TESTS
Description Value
Transmitting RF power
system
(DSIF) +0.5 70. 0 -0.0 dbm
Antenna SAA SCM Circuit SAA SCM Receiving Circuit
gain ZO.O± Z.O db -0.5) db
51. 0 (+I. O, loss
-0.5 ± 0.0 db -0.4± 0. i db system loss A B -3.2 -3.7 tracking noise Z40 SNR ± 24 Hz loop ± 0.3 + 0.3 db db (SC-2)
Receiver Receiver
Uplink
carrier
]Equivalent Bandwidth Threshold
12 db
Uplink
channel SNR 9 db
Threshold
System
noise 2700°K
Temperature Equivalent Bandwidth Data/subcarrier index Subcarrier index noise (predetection) modulation
13430 7.2
Hz
/ carrier
modulation
1.6±
0.16
5. 3-Z8
TABLE
5.
3-6.
DOWNLINK STV, AND
PARAMETERS CDC TESTS
FROM
FAT,
Description
Value
Transmitting RF power
system
(SC-2)
Transmitter (low Transmitter (low Transmitter (high Transmitter (high Planar Circuit power) array loss power) power) power)
A 21.09 B 21. A 40.6 B 40.6 gain (+0. I, -0. i) dbm db (+0.3, -0.05) dbm 1 (+1.2, -0.2) dbm (+0.21, -1. 19) dbm
27.0+0.5
Transmitter Omnidirectional Transmitter Omnidirectional Transmitter Omnidirectional Transmitter Omnidirectional Planar Carrier Receiving Antenna SAA SCM array frequency system gain (acquisition (85-foot
A antenna B antenna A antenna B antenna B -2. 7 (+0.2, -1. O) db B -Z.8 (+0.2, -I.0) db A -1.8 (+O.Z, -I.0) db A -Z. 0 (+O.Z, -i.0) db
-2. Z (+0.0, 2295 (DSIF) MHz
-0. 3) db
aid antenna)
antenna)
21.0+ 53.0
1.0 (+I.0,
db -0.5) db
5.3-Z9
Table
5. 3-6 (continued) Description Circuit SAA SCM Effective Maser Parametric Amplifier Lunar {Johannesburg SAA antenna) 3Z0 + 50°K II0 + B5°K noise temperature 55 + 10°K loss -0.5 -0.18 + 0.0db + 0.05 db Value
temperature channel
Carrier
Equivalent maneuvers Equivalent coast mode Threshold
noise bandwidth (at threshold) noise bandwidth (at threshold) SNR
for
15Z Hz
for
12 Hz
Acquisition Maneuvers Coast SCO descriptions Equivalent bandwidth, predetection noise Hz + 10 percent mode
9.0 14.0+ 11.4
db 1.0 db db
4400 bits/sec 1100 bits/sec 550 bits/sec 137.5 bits/sec 17.2 bits/sec Strain gage Strain gage Strain gage Reject/enable Gyro speed 1 2 3
4770 1190 644 158.5 25.1 Z81 524 464 377 874
5. 3-30
Table
5.3-6
(continued)
Description
Value
SCO
center
frequencies,
KHz 33.0 7. 35 3.90 0.96 0.56 1.70 3.00 5.40 2.3 5.4 ratio for
4400 bits/sec Ii00 bits/sec 550 bits/sec 137.5 bits/sec 17.2 bits/sec Strain gage Strain gage Strain gage Reject/enable Gyro speed Threshold telemetry I 2 3
signal-to-noise data, ±I. 0 db
4400 bits/sec II00 bits/sec 550 bits/sec 137.5 bits/sec 17.2 bits/sec Strain gage Strain gage Strain gage Reject/enable Gyro speed SCO modulation 4400 bits/sec I 2 3
I0.0 I0.0 I0.0 10.0 I0.0 7.0 7.0 7.0 10.0 10.0 indices, ±i0 percent 1 6 0 935 0 3 1 15 1 45 1 45 0.615 0.615 0.61 0.655 1.600
1100 bits/sec 550 bits/sec (acquisition} 550 bits/sec 137.5 bits/sec 17.2 bits/sec l 2 3
Strain gage Strain gage Strain gage Reject/enable Gyro speed
5.3-31
z)
Spacecraft
transmitter
low
power
output
is
Plow
: Phigh
PDSIF
H
+ PDSIF
L
(dbm)
where
Plow = transmitter low power output
Phigh
: telemetered
transmitter
high
power
output
PDSIF
H
= DSIF
received
signal
level
at high
power
P DSIF L
= DSIF
received
signal
level
at low
power
3)
Spacecraft
omnidirectional
antenna
gain
(uplink)
is
PR G R = PT OT (_) 2L
where G R : received omnidirectional antenna gain (uplink gain)
PR
= received
signal
level
(determined
from
spacecraft
AGC)
PT
= DSIF
nominal
transmitter
power
G T k R L
= DSIF
nominal of
antenna uplink
gain signal of computation and DSIF losses downlink parameters
= wavelength : slant range
at time
: nominal
spacecraft
(Note: For downlink gain, appropriate are inserted in a similar equation.)
5.3-32
4)
Signal-to-noise
ratio (SNR) for any subcarrier PS
PN
is
SNR where
MPR
K Tef f BWsc
PS
= signal
power
in
predetection
noise
bandwidth
PN M
= total
noise
power
in
predetection
noise
bandwidth
= carrier based carrier
to on
subcarrier subcarrier
modulation oscillation
loss adjustment modulation index
constant on the
PR K
= received : Boltzmann's
carrier constant
power
reported
by
the
DSIF
Tel f
= DSIF
system
temperature
reported
by
the
DSIF
BW sc When using
= subcarrier
equivalent
predetection
noise
bandwidth
these
equations,
attention
must
be
given
to
the
desired in flight, paramagainst experience. as for the
accuracy of the answer. Since several parameters not measurable spacecraft telemetry, and DSIF station reports are used, computed eters have potentially large errors. Their validity is thus weighed similar test data and/or is judged quite subjectively based on past These equations are not used so much for their numerical results total picture of subsystem problems or computation but subtle errors will not. Bit Error Rate performance errors will tend generated. to be Any uncovered gross in this
subsystem analysis,
Calculations
One subsystem parameter of interest is telemetry bit error rate (BER). This parameter serves as an example of the problems encountered when attempting to evaluate postmission data. BER is required to be less than 3 x 10-3 at input SNR ratios of 10 4- 1 db. (A change effective with SC-3 will allow only 9 + 1 db for a BER of 3 x 10-3.) BER cannot be measured in flight, but the word error rate can. Therefore, real-time used, assuming a bad parity word represented a single additional assumption that the data used were representative, observed BER was computed (see Table 5. 3-7). bit printer error. the data With worst were the
5. 3-33
TABLE Time, hr:rnin:sec
5.3-7.
BIT
ERROR
RATE
DATA
SUMMARY
FOR
DAY
265
Number
of Bits 3256
Parity
Errors 2
BER 0.6 x 10 -3
03:29:59 03:35:42
to 1
03: 35:55 to | 03:47:28
5192
4
0.8
x 10 -3
03:48:06 03:56:34
to }
4224
0.7
x 10 -3
03:57:10 04:04:47
to
}
4224
14 25 8888 -3 - 2.8 x I0
04:15:51 04:05:24
to }
4664
11
shown
The below DSIF System (DSIF-
SNR at this time of the from Equation 3: AGC/ll00 bits/sec at
observed
high
BER
was
computed
as
04:07:45 = 44.7°K
= -138.7 = 16.5 db
dbm
noise temperature 11 pretrack) constant = 1190 Hz
Boltzmann's Bandwidth Noise power
= -198.6 ± 10 percent (+0.41,
dbm/deg/cps = 30.75 -0.46) dbm (+0.41, -0.46) db
= -151.35 loss -2.01 -4.56 loss power = -2.
Modulation Carrier Subcarrier modulation Subcarrier SNR
(+0.40, (+0.62, 55 (+0.
-0.46)
db
-0. 73) db 22, -0. 27) -0.27) = 10. db dbm 10 + 0.68 db
= -141.25 power
(+0.22, - noise power
= subcarrier
5.3-34
The tolerance on this computation is only approximate and is probably greater. Based on the SNR requirement of i0 + I db, the measured parameter (BER) meets the specification. However, it is not clear that the new requirement of a SNR of 9 ± i. 0 would have also been met. Omnidirectional Antenna Gain Maps
level
In order to better visualize and interpret the significance of the signal data, traces of the earth vector on the omnidirectional antenna gain
contour maps are presented. Figures 5. 3-8 and 5. 3-9 show the antenna upand downlinks. Since signal level variations are, for the most part, the result of increasing range (i.e., more space loss) and changing omnidirectional gain, these plots allow visualization of the expected signal level changes for comparison with plots of uplink and downlink signal levels versus time.
5. 3.4.2
Mission
Phase
I:
Prelaunch
to Spacecraft
Acquisition
During the prelaunch phase, subsystem performance is assessed during the launch pad systems readiness test (SRT) and prelaunch countdown test. Next to assuring normal system performance prior to launch, the most important subsystem data taken during this phase are transmitter and receiver frequency data. Frequency data are used to predict the frequencies at initial acquisition and are transmitted from the Cape prior to launch. The DSIF, in turn, uses these data to tune the DSIF receiver for one-way lock and the DSIF transmitter for eventual t_vo-way lock. The prelaunch frequency data for the transmitter and receiver are
plotted in Figure 5. 3-10. Also, the measured frequencies, as well as the predicted frequencies at acquisition, are noted. These frequencies tended to decrease with time, with the notable exception of the receiver best-lock frequency in the L-10 report. Since a temperature increase always causes a frequency decrease, and since the temperature in the compartment was increasing, the data were considered reasonable with the exception of the receiver frequency at L-10. The temperature directly affecting the frequency is not actually measured, since the telemetered sensor is in the thermal tray and not at the voltage controlled crystal oscillator. Relative temperature versus frequency information is thus considered to be most reliable. (See recommendation ment, the receiver prediction the transmitter prediction The predicted 2 in subsection 5. 3. 3.) Based on frequency was taken from the L-20 frequency from the L-10 report. were thus: -- 2295. : 2113. 001694 309168 MHz MHz this report judgand
frequencies
Transmitter (one-way) Receiver (two-way)
5. 3-35
_2
3_<)i
W
VI
4_
i
' +11
!
-_..4
0
4_
<
0
.r4
F
o
°T'4
0
0a I-4
°,-4 1._
O
_q
._-.4
! • / i
J
j
(
3
_d
I e¢3
_d
iI
_0
i
/ /
i
7
i,i
_
i
°_
............
'T--TI
! g
'
t
.LJ
5.3-36
¢)
<
0
_D
. ,-N
qD
_
m
0
fM
"2> v_
_
• _-.I
_
.r-I
£
0 ..z
°_-,I
o
v
of)
I
5. 3-37
13-,
0
<
,--4
0
,o
%
<
'_ 0
0
u
v
0
I
_4
o_-i
5. 3-38
D
<
o
O
°_-I °_-_
e
D
©
N
D
<
c *d ®
-'_ .,_
2
"_;N _'_ 50
_ _
f
O
m
O
/
I
M
N
'I°
5.3-39
×
Figure
5. 3-I0.
Prelaunch
Frequency
Data
and
Actual
Frequency
at Acquisition
5. 3-40
The actual were:
frequencies
at initial acquisition
(as
shown
on
Figure
5. 3-I0)
Transmitter (one-way) Receiver (two-way) The difference between predicted and
= 2294. = 2113. actual = 1915 = 9776
999779 318944
MHz MHz
was: Hz Hz frequency was Had this frereduced to 3. 0 KH z.
Transmitter Receiver It should be noted to the L-10 report been used, the
closer quency
that the actual receiver acquisition frequency, which was discarded. in prediction would have been
error
Table 5. 3-8 is a summary of the significant events during the initial RF acquisition at Johannesburg. The signal levels at receivers A and ]3 during acquisition are shown in Figures 5. 3-11 and 5. 3-12. One-way acquisition was accomplished in about 12 seconds from first RF contact, and two-way lock was accomplished in 10 minutes. Problems with the antenna drive, coupled with a low receiver best-lock prediction frequency, caused 5- to 6-minute delay in the two-way acquisition as compared with an optimum acquisition. Figure 5. 3-ii shows that receiver A was captured a in
the AFC mode right after DSIF transmitter turned on. Antenna drive problems are also clearly shown as signal level variations in the receiver passband. Figure 5. 3-12 indicates that the signal was in the passband of receiver B at turnon (the receiver has a 13 KHz passband), but because the turnon frequency was low, the doppler shift caused the signal to go out of the passband without locking up. DSIF transmitter tuning resulted in the signal slewing back into the receiver passband. Receiver phase lock is shown occurring about 3 minutes and 12 seconds after initial transmitter turnon. The spacecraft high power transmitter to high turnoff was turned off 32 minutes The and maxi-
12 seconds after being commanded mum allowable time to accomplish
power by the Centaur. is 1 hour.
5. 3-41
TABLE 5. 3-8.
Event Transmitter power on Spacecraft heard by B high
ACQUISITION EVENTS
Comments Spacecraft power by commanded Centaur. 5 seconds to high
GMT (Day 263), hr:min: sec 12:44:21
signal DSIF
first
12:54:55
Initial contact to predicted
prior
first visibility.
DSIF acquires spacecraft in one-way mode DSIF switch from
12:55:07
One-way utes and DSIF with
acquisition in 23 min7 seconds from launch. to maintain on SCM. contact
12:57:10
unable
acquisition to 85-foot
antenna (SAA) dish (SCM) from SCM 12:57:50
spacecraft
DSIF switch to SAA DSIF switch to SC M DSIF switch to SAA DSIF switch to SC M
from
SAA
12:58:15
DSIF with
unable
to maintain on SCM.
contact
spacecraft
from
SCM
12:58:20
from
SAA
13:00:00
DSIF now able to track spacecraft on SCM. Signal level at ground receiver increased 26 db due to increased gain.
DSIF
transmitter
turned
on
13:01:36 13:01:46 (From telemetry) Receiver A in AFC capture mode. Receiver B not phase locked. DSIF receiver dropped lock, indicating phase receiver B. phase lock on
Signal in passband of both spacecraft receivers
Phase
lock
receiver
B
13: 04:53
DSIF acquires spacecraft in two-way mode
13:04:58
DSIF reacquired downlink, indicating complete two-way acquisition in 32 minutes and 58 seconds from launch.
DSIF confirms two-way phase DSIF switch to SAA DSIF switch to SCM
good lock SCM
13: 04:59
from
13:05:10
DSIF with
unable
to maintain on SCM.
contact
spacecraft
from
SAA
13:06:20
DSIF now on SCM. receiver Spacecraft 32 minutes
able to track spacecraft Signal level at ground increased 27. 8 db. was and in high power for 12 seconds for
Transmitter power off
B
high
13:16"33
initial acquisition phase (a maximum time of I hour is allowed). 5. 3-42
2
Figure 5. 3-Ii. Signal Level at R_ceiver A During Acquisition
Figure 5. 3-1Z. Signal Level at Receiver B During Acquisition
5. 3-43
5.3.4.2
Mission The coast
Phase phases
Two:
Coast of the following: -Period from initial spacecraft the -Z axis
consist
l)
Pre-Canopus acquisition spacecraft is pointed
acquisition
until Canopus acquisition, during which time attitude is random in roll and the spacecraft toward the sun. - Period maneuvers. from Canopus acquisition
z)
Premidcourse midc ourse
until the
A normal However, tumbling
mission would contain a third coast or postmidcourse phase. since the mission became nonstandard, the postmidcourse or phase will be treated separately.
Figures 5.3-13, 5.3-14, and5.3-15 are plots of DSIF, Receiver A, and Receiver B signal levels from launch to the midcourse maneuver. The premission predicted signal level after Canopus acquisition is shown on each of these figures. Since the spacecraft attitude is random in roll prior to Canopus acquisition, no premission predictions are made for this period. After Canopus was acquired, with the following predictions: i) 2) 3) DSIF levels A B agreed levels levels the signal levels came into close agreement
to within were were +2 +I
+0.5 to +4 to +3
to -2.5 db db above above
db
of predicts.
Receiver Receiver
predictions. predictions. show and traces of the earth omnidirectional
vector
Referring to Figures 5.3-8b and 5.3-9, which relative to omnidirectional antenna B downlink
antennas A andB uplink gain contours, it can be noted that changes in signal levels during the pre-Canopusacquisition phase andrightat Canopus acquisition are in complete agreement with the antenna gain contour maps. The antenna gains during the pre-Canopus phase were B as follows =>-i approximately: to -4 db going to -I db
l)
Omnidirectional antenna at Canopus acquisition Omnidirectional antenna at Canopus acquisition Omnidirectional antenna Canopus acquisition 5.3-14 during
downlink
2) 3)
A
uplink
=>-2
to -l db
going
to -i0
db
B
uplink
=>-5
to -3 db
going
to -l db
at
specific
Figures events
and 5.3-15 show the coast phases. the station
signal level variations The special receiver transfer from
caused by threshold test
(see subsection are of particular
5.3.2.1) and interest.
JohannesburgtoGoldstone
5.3-44
5. 3-45
Figure
5. 3-14.
Receiver
A
Automatic
C,ain
Control
i,! :,1!!ii
ii!!:i:11;!::: i_ii
_,,+]
_ iii
-l.
Figure
5. 3-15.
Receiver
13 Automati(
C,ain
Control
5. 3-46
5.3.4.3
Mission
Phase
Three:
Canopus
Acquisition
Maneuver
At approximately L + " hours, initiated. Two complete rolls about make a star map adequate to identify roll were required to finally acquire
the star acquisition maneuver was the Z axis were required in order to Canopus. An additional 240 degrees the star.
of
Real-time analysis indicated that the roll maneuver would take the earth vector through deep antenna nulls, thus requiring that the data link be in one-way (nontransponding) mode. Also, analysis indicated that the downlink telemetry threshold could be exceeded during a portion of the roll maneuver if only omnidirectional antenna B was used for transmission. Omnidirectional antenna gains of -30.0 +10 db were predicted during this maneuver. At 18:30:46 GMT, transmitter B was commanded to high power. Transponder B was turned off at 18:33:01 GMT, and DSS-51 reacquired the spacecraft in the NBVCXOmode. Star mapping was initiated at 18:37:34 GMTwith the spacecraft transrnittiondata in mode 5. The initial roll on omnidirectional antenna B produced downlink signal variations of approximately 40 dbwhich agreed with the premaneuver predictions. Spacecraft datawere sustained throughout the maneuver but were sufficiently noisy that another roll on omnidirectional antenna A was initiated at 18:54:45 GMT. A complete star map was obtained from the two rolls. Spacecraft-received signal levels during the roll maneuver indicated deviations of approximately 34 db on receiver A and 30 db on receiver B. This again agreed with premaneuver predictions. However, it was at this point that the 20-db bias in the receiver B absolute signal level was detected. This anomaly was discussed earlier in subsection 5.3.2.1. Omnidirectional antenna B was again selected at 19:06:37 GMT and the spacecraft allowed to roll until Canopus was acquired. It was necessary to manually lock on to Canopus, and this step was initiated at 19:11:57 GMT. At 19:14:21 GMT, transponder B was turned on, and DSS-51 acquired the spacecraft in two-way lock at 19:15:39 GMT. Transmitter B high power was commanded off at 19:22:05 GMT, which resulted in 51 minutes and 19 seconds of high power operation for star acquisition. The DSS-51-received signal level for low power operation was -132.9 dbrn, a 20.9-db decrease from high to low power operation. A nominal 1100 bits/sec telemetry margin of +5.0 db existed at this point. Figure 5.3-16 is a plot of the DSIF signal level during the period of the star maneuver, with significant event times noted. Figures 5.3-17 and 5.3-18 are expanded plots of the same data taken from station reports. The equivalent omnidirectional antenna gain is also shown. Since the resolution of the expanded data is relatively poor, comparative antenna gains are shown only at selected points. Signal level variations agree well with the antenna gain valves, giving a high degree of confidence in the antenna patterns.
5. 3-47
68189-2-151
il:!
-,"4
I
t:m u3
!
,S
i
I
r¢3
_4
(1)
-,--t
..%
5. 3-48
r/?
_)
r--oo H
4-1
0
5. 3-49
0
If
r_
0
II
5. 3-50
Figure 5.3-19 shows the signal level at receiver A during the entire star maneuver. These data, taken from spacecraft telemetry, have much better resolution than the DSIF data. Telemetry points are plotted for every degree of spacecraft motion, showing equivalent omnidirectional antenna gain for a complete spacecraft revolution. Relatively good agreement existed between the omnidirectional gain values and the signal level except in the primary antenna null. This disagreement is hypothesized to be caused by the omnidirectional antenna dipole angle difference between the SC-2 omnidirectional antenna and the omnidirectional antenna pattern data. Although expected, this null shift demonstrates the sensitivity of the antenna patterns to omnidirectional antenna positional tolerances. In future missions, external spacecraft configuration changes will result in omnidirectional antenna pattern changes, especially at the nulls; therefore, operation in or near these nulls signal levels. may cause poor correlation between actual and predicted
Figure 5.3.-20 shows the signal level at receiver B during a portion of the roll maneuver. As in the case of receiver A automatic gain control data, this figure contains predicted omnidirectional antenna gain valves over a complete roll period. There is good agreementbetween the omnidirectional antenna pattern and automatic gain control data in the gain region above -8 db and to a lesser extent down in the null. As expected, the pattern in the null agrees much better than did that for omnidirectional antenna A. 5.3.4.4 Mission Phase Four: Midcourse Maneuvers
Roll-yaw was selected from four possibilities as the midcourse maneuver and was optimum for the communications link. Real-time predicted the following variations in nominal omnidirectional antenna during the maneuver: i) 2) 3) Omnidirectional Omnidirectional Omnidirectional antenna antenna antenna B A B downlink: uplink: uplink: -6 -l -2 > < G G < 1.6 db db db
analysis gain
> -20 > -15
> G
Predicted minimum margins were 16.0 db for 4400 bits/sec telemetry, 8.0 db on receiver A, and 13.0 db on receiver B command links. Two-way (transponder) mode was recommended as a result of the special threshold test run at 01:36 GMT.
At 04:36:43 GMT, the spacecraft at 04:37:54 GMT the 4400 bits/sec data level was -123.3 dbm prior to maneuver. was initiated, and at 04:48:06 midcourse maneuver ended reading -123.3 the maneuver,
was commanded to high power, and rate was selected. DSS-II signal At 04:44:00 GMT, the roll maneuver The prelevel
GMT the yaw maneuver was initiated. at 04:51:57 GMT with the DSS-II signal
dbm and having indicated approximately as predicted. The maneuver performed
a 2-db variation during is mapped on the
5. 3-51
specific gain patterns as shown in Figures 5.3-8 and 5.3-9. Since the maneuvers were performed in telemetry mode i, no receiver AGC data are available to check against the premission prediction. However, the command link was maintained with no receiver indexing, indicating above-threshold oper ati on.
Figure 5.3-21 shows the DSIF receiver signal level variations during the maneuvers. The large amount of noise in the data (a data processing problem?) allows only limited analysis. Data points taken from this plot are compared with relatively 5.3.4.5 Mission course) to predicted omnidirectional good correlation. Phase Five: Nonstandard gain levels in Figure 5.3-22,
Spacecraft
Tumbling
(Postmid-
At 05:00:02 GMT, the vernier engines were fired for the midcourse correction. Because vernier engine 3 apparently did not fire, the spacecraft became unstable and began to tumble. Figure 5.3-23 shows a plot of the DSIF-receiver signal levels for the period just prior to engine 70 seconds after ignition. Large signal level variations began continued throughout the rest of the mission. burn here through and
Ground link AGC was used in the postmission tumbling dynamics analysis (see Section 4.7) to determine spacecraft motion during the first 20 seconds after ignition. Assumed earth vector paths on the omnidirectional antenna B downlink contour map generated plots of signal level versus time for comparison to actual DSIF signal levels. Table 5.3-8 contains a summary of the primary spacecraft tumbling
periods throughout good data were not and dynamic phase During power
the remainder of the mission. Except available, the period was determined error. tumbling, the telemetry bits/sec for low power
in cases where from station AGC
for high
spacecraft and 137.5
bit rate was If00 bits/sec transmitter operation. Twothe dbm remainder of the (high power) or
way transpondor lock was also maintained mission. The downlink signal level was -135 to -155 dbm (low power). 5.3.4.6 Mission Data Plots
throughout -i17 to -135
Subsystem telemetry signals are shown in Figures 5.3-24 through 5.3-28. {See also Figures 5.3-14 and 5.3-15 in the coast phase discussion for AGC signals.) The plots were terminated at midcourse because spacecraft tumbling thereafter caused many bad data points (from double bit errors). Postmidcourse plots generally were mis-scaled and unreadable, but some the more usable are found in Figures 5.3-29 and 5.3-30. Comments are omitted since most of the sudden variations are due to telemetry processing problem s. of
5. 3-52
1
j
•
t,
1
]
. MI_S flail ltli IIIII451
la,
Figure
5. 3-20.
Receiver
B
Automatic
Gain
Control
(Telen
E_UT
5. 3-53_
FRAME
/
letry)
During
Star
Map
letry)
During
Star
Map
FOLDOUT. FR, ,M_. '_ k
68189-2-156
_>
<_
0
,r-+
_Q 0
>
>
,+-i
>
+-4 rq I
M
I
I
I
.+'=1
7_,'t._7 7t_'Jt,_IS" J/gZ]?
5.
3-55
Figure
5. 3-22.
Predicted DSIF Signal Levels Roll and Yaw Maneuvers
During
Midcourse
5. 3-56
Cr_
I
b_
I
12.5
SEC
_ - 140.0DBM
_'050000 THRUST GMT EXECUTION
-I 23.3DBM
CHART
SPEED 2MM/SEC
Figure
5. 3-23.
DSIF-II Automatic Initial Spacecraft
Gain Control Tumbling
Variations
for
5. 3-57
TABLE 5, 3-8.
Time Operation day:
SPACECRAFT
of Commands, hr: rnin: 05:00:13 sec
ROTATION PERIODS
Time hr: Interval n, in', sec Used, Period, seconds
Midcour
se terminate
264:
05:00:30-38
O, 81
(first firing) Thrust Rate power mode on off 05:00:4i 05:03:48 05:03:44-45 05:04:48-51 05:11:45-48 05:14:46-48 O. 883 O. 89 1.08 1.18
Inhibit gas jet amplifier (end rate mode) End engine (2 seconds) 3rd engine (2 seconds) 4th through firing 9th engine (2 seconds) lOth engine lbth engine 22nd engine through firing* through firing':-" through firing 21st firing
05:14:29
07:28:25
No
data
firing
07:50:03
No
data
8th
engine
19:44:59
19:44:58-59 20:07:04-05 23: 5c): 5q- 00:00:01
l.Z (resolution 0. 99 (wave shape 0.91 s vary) poor}
firing
23:33:23
15th
265:
01:28:11
02:01:15-18
02:39:14
03:17:21-23
O. 86
26th '>:_ firing
03:39:07
03:47:53-55
0.82
27th engine (2 seconds) 28th through engine firing 33rd engine (2 seconds) 34th engine (Z seconds)
03:47:56
04:17:28-29
0.74
3and '>',_ firing
04:41:20
04:56:09-12
0.73
04:56:12
05:43:16-18
0. 705
firing
05:43:19
05:59:58-59
0.68
35th through 39th engine firing* g_ 40th (20 Retro engine seconds) firing, firing
07:49:25
No
data
08:05:12
09:34:09-28
O, 44
delayed
09:34:28
09:34:54-59
0.52
*Group ':"#Group
consists consists
of of
five five
0. 2-second 0. E-second
burns burns.
followed
by
one
2. 0-second
burn.
5.3-58
m
H_
:_t _t l
-t_tt
4_H
Hi:
fL
!q!
D
,st++.+
_ 7 r/J
r
!lil• li !li!
trlt
:_F!' ; il !_
;HI
tEIi : : Ii
:ti
H
i!iillsi!it!/i!i l !il il iii, [!: i iiitli_
_,: :i i:i: i]
il!l sl
_:1 il
;ii
h
kiljkt:i
fillllil
ILJ H_i
rl:l i tll
:111
:l:t
2:t11', !i_iiji
1,
ilii i;I _
[i:l i!_i
u.
i ii;i;,ii i
5. 3-24. Receiver A Automatic Frequency Control
Figure
:ll
il:
:i!
• i
Figure
5. 3-25.
Receiver
B
Static
Phase
Error
B
5. 3-59
Figure 5. 3-26.
Receiver B Automatic
Frequency Control
•it, !ii! -_
_!!
:!
_'1i!!i _i .....
:!!i
!.!!iiii!
i i:
iif_ iHi
:ii!i
i
i
__ _' ilii
_::i
!D!
....2LI
:i!!
iii
ili, .... ,,,_
4;;
....
I;
i! ;U
::: ! ::: .....
1"t:
i:1
'::
!![: !!:H :i!i
!! T!: :: t?,
! !! re.
i
Figure
5. 3-2-7.
Transmitter
A
Temperature,
Premidcourse
5.3-60
li !lillilllli l l]!li i _
I!I I
' _iiilii!!l_l
i :ilqqi!I I
:Iii !:' !
11
'ii:ii:i_ii
! I I
H
?
ii
_, ........
!!!
Figure
5. 3-28.
Transmitter
B Temperature,
Premidcourse
5. 3-61
kill,
il.
Figure
5. 3-29.
Transmitter
A
and
B
Temperatures,
Postmidcour
se
m_!!iii*
.... _
_
'
_
'ii
_:
!l I ?,
_',!
_ _
I iiiiHU tit!tf!t!it fttt I!I!! ilii] :_t:::i_I',THTI {ti ]!IU":!!ilii i:,tiLLi]i ,,_,, ttiiltti it:tilttti ftftltttitftfl!H!
I _I)rtHiltt! fill !{!!!!t_',i_J!i iiliHIHIE _t777l!!ri17;ii:I7i
I!_i! l_i titit):::iiti_j _ I,_ tl[tftit{tfJH it!ilttt fit!
,_: _,f,_:_::7;:ilii i: )
III_'It "(" ,_l 5. 3-30. " Figure Receiver Receiver A B Automatic Static Phase Frequency Error Control and
5. 3-6g
For the premidcourse plots, performance, and no unexplainable of each figure and the more significant
all data variations events
indicated were follows:
normal noted.
subsystem A brief summary
Receiver A Automatic Frequency Control (Figure 5. 3-24)-Receiver A was in the automatic frequency control mode throughout transit. These data represent the DSIF transmitter frequency offset from the automatic frequency control center frequency during the transit to doppler shift rate is noted at acquisition. Steps station transfer because the stations retuned their phase. A large error due in the data occurred at transmitters. Due to the
high impedance of this signal, several predicted signal processing effects are apparent. Steps occurred in the data at high power turnon due to return line drop caused by the additional current in the ground return lines during high power operation. Spikes occurred during engineering interrogations of mode 4 due to step change in commutator unbalance current. Receiver B Static Phase Error (Figure 5. 3-25)-Receiver B was used for transponding through most of the mission. These data thus represent the DSIF transmitter frequency offset from the receiver phase lock center frequency. Since these data are analogous to the automatic frequency control data discussed above, the comments apply equally well to these data. It should be noted, however, that this signal is not as sensitive to signal processing effects. Receiver B Automatic Frequency Control (Figure 5.3-28)-Since the receiver was phase locked during the majority of the transit phase, this telemetry signal was not a valid signal. Unlike the static phase error signal, which has a 0-volt output when not being selected, the automatic frequency control telemetry does vary with frequency changes even when the receiver is not in the automatic frequency control mode. However, the telemetry is not valid and is essentially meaningless. Transmitter Traveling-Wave Tube Temperatures (Figures 5.3-27 and 5.3-28 for premidcourse period and Figure 5.3-29 for postmidcourse data) -- These data represent the traveling-wave tube temperatures used
for
high power transmitter operation. Figure 5.3-29 contains telemetry glitches caused by the spacecraft tumbling, as discussed earlier, and shows temperature variations during the short high power transmitter operation times. Receiver A Automatic Frequency Control and Receiver B Static Phase Error (Figure 5.3-30)This plot is similar to figures 5.3-25 and 5.3-26 except for the period after midcourse when telemetry processing glitches occur. This information is included, however, to show the effects of tumbling-induced doppler shifts.
5. 3-63
5.3.5
REFERENCES W.C. Collier, "The Change Flight In Receiver of SC-2," B AGC Characteristics Aircraft Company,
le
Prior to and During IDC 2294.2/44.
I]ughes
2.
"Surveyor Mission B Space Hughes Aircraft Company,
a
Flight Operations Report," Volume SSD 64257R, November 1966. Final Report," October 1966. Volume 2, Hughes
l,
"Surveyor Aircraft
I Flight Company,
Performance SSD 68189R,
1
"Surveyor Mission and Performance SSD 64241R.
B - Telecommunications Subsystem Summary," Hughes Aircraft Company,
Prediction
5.3.6
ACKNOWLEDGEMENTS Frank Joe K. Rickman, coauthor time team Mike by the spacecraft/performance/ (Vic Amsteader, Bill Mitchel, Weiner, and Mike Williams) were coordinator and coauthor
Votaw,
Notes and records made in real analysis/command telecommunications John Steineck, Vern Story, Joe Votaw, of great use in postmission analysis.
5. 3-64
5.4
SIGNAL
PROCESSING
5.4.1
INTRODUCTION The I) Z) 3) 4) 5) signal processing signal engineering signal processing data rate processor auxiliary subsystem processor signal is composed (ESP) processor (AESP) of the following units:
Engineering Auxiliary Central Signal Low
(CSP) (SPA}
auxiliary
(LDRA) with a total of 6 operational 5 digital bit rates, modulation data
These
units
contain
Z electronic
commutators
modes, g analog-to-digital 17 subcarrier oscillators
converters that have available for transmission of pulse coded
and continuous real-time data, 9 summing circuits for the measurement of electrical subsystem performed normally throughout
amplifiers, and signal conditioning currents and temperatures. The the mission. telemetry included
can
be
A summary of test and flight values for signal processing found in Table 5.4-I. Values for the SC-I flight have been A complete 5.Z (RF data
for comparison. found in Section
mode, bit rate, and configuration log can be link) and will not be repeated here. All signal the full-mission details of each plots correc-
processing corrections throughout this report tion will be discussed
made to telemetry signals on are given in Table 5.4-2. The in Subsection 5.4.4.
5.4.2
ANOMA No
LIES anomalies were attributed to signal processing in the SC-Z flight.
5.4-I
(-_ ,.c_ _ a_O
,r,4
o _
o o
d
I
f'd
m
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oo oo
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_o
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I I I
t_I
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oo
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0 0,,1
I
0 ,,DO0
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I
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oo
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L_L_ L_ L"-
cr,=.=4 ,-=4 I I
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{:7',
I
u"l o o
c_
I
c,,.)
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I
M
N _a <
ma
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4-) ,-"4 4.1
X
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8"
u > >
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r./?
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I
0_
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0 0
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I
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m
5.4-2
TABLE
5.4-2.
IN-FLIGHT CALIBRATION MADE ON SC-2 MISSION
CORRECTIONS PLOTS
Correction Reference voltage M-3, P-l, Unbalance current
Telemetry M-4, P-2 M-6,
Signals M-7,
D-7, D-8, EP-I, EP-2, EP-3, EP-5, EP-10, EP-Z3, EP-30, FC-4, FC-3Z, FC-53, P-l, R-29 EP-4, EP-II, EP-17, EP-24, EP-6, EP-7, EP-9, EP-14, EP-16, EP-21, EP-22, EP-25
Current
calibration
5.4.3 5.4.3.1
SUMMARY Signal
AND Processing
RECOMMENDATIONS Performance Summary
mission. On-board current possible has been 5.4.3.2
The signal processing subsystem performed properly throughout the All telemetry channels gave proper indications in all modes used. calibration signals (reference voltage, unbalance current, and calibration) were used for telemetry accuracy improvement. A method to correct developed. Recommendations The following recommendations are made: be modified data. At is not in certain temperatures for 4400-bits/sec errors
l)
Initial processing of the DSIF magnetic tapes should to record (on digital tapes) all teIemetry bit stream present, data are discarded whenever decommutator lock. A previous recommendation to replace current corrections with values constant for the entire reconsidered. Loss of accuracy may result. An investigation should be current correction. Means outputs and significant. individual
z)
calibration signal flight should be
3)
initiated on the use of the unbalance must be provided to handle capacitive factors if these are shown to be
switch
5.4-3
5.4.4 SIGNAL PROCESSING ANALYSIS 5.4.4.1 Unbalance Current Corrections In each telemetry commutator, transistor switches connect each analog output voltage (representing a spacecraft voltage, current, or temperature) with a common commutator line connected to the input of one of two analog-to-digital converters. A bootstrap unloader circuit is connected to this common line to reduce the stray capacitance, equalize the load impedance, and provide bias currents for the commutator and master switches. Since these bias currents are not exactly equal, a difference or unbalance current exists. The telemetry circuit being sampled must supply this current, causing an error in the measured voltage proportional to the output impedance of the circuit. mutator The unbalance current for a specific telemetry channel in each com(S-5 for ESP and S-7 for AESP) is measured in telemetry modes
Z, 4, and 5. Figure 5.4-i shows S-7 up to midcourse. A warmup effect can be noted in that each time the AESP commutator is turned on, the initial value of unbalance current is up to 0.Z microampere lower than the value assumed after a few minutes of operation. Although no plot of S-5 has been included, typical values have already been given in Table 5.4-I. The change in unbalance current from mode i to mode 4, due to internal ESP load changes, is readily apparent. This effect also occurred on SC-I, and is part of the spacecraft signature list. The final report for SC-I made a number of recommendations for unbalance current corrections in automatic processing. Many of these have been accepted (for instance, delete corrections for temperature and capacitor-output channels), but as yet no means has been provided for individual selection switch corrections. A limit (+I0 microamperes) has been put in the correction processing to prevent wildly inaccurate "corrections", based on bad data values of unbalance current, from being made. 5.4.4.2 Potentiometer Reference Voltage Corrections
The nominally 4.85 reference voltage is supplied by either the ESP or AESP units to the landing gear and solar panel position potentiometers, to the propulsion pressure transducers, and to the secondary sun sensors. This reference voltage, derived from the 29-volt nonessential bus, varies due to load and input supply voltage changes. The ESP voltage is telemetered in modes Z and 4, and can be used to correct the affected signals whose calibrations are based on a reference voltage of exactly 4.85 volts. Since the AESP voltage is never telemetered, it must necessarily be obtained through computation.
5.4-4
Figure 5.4-i.
AESP
Commutator
Unbalance
Current
5.4-5
The mechanism position signals do not normally change in flight after initial deployment, since they are mechanically held. Therefore, any apparent difference in a given signal reading from the ESP commutator to the AESP can be due only to a corresponding change in commut=tor-supplied reference voltage. Based on this assumption, Table 5.4-3 was prepared to show both prelaunch and coast phase I calculations of the AESP reference voltage. Due to the granularity of the signal values used in the calculations, it seems reasonable to use a median, rather than a r_ean, val_e. Thus, the calculated
AESP (4.94 5.4.4.3 reference volts) at L Current voltage was + 15H.;:= Calibration i006 bcd (4.92 volts) at prelaunch and 1010 bcd
Signals
Current measurements are accomplished by measuring the voltage drop across a low resistance shunt which is in s_ri_s with the power line being monitored. This measurement is in the range of 0 to i00 millivolts. Since this voltage is not referenced to ground and is not scaled to the 0- to 5-volt telemetry input level range, it is necessary to amplify it with a differential amplifier. The nominal gain of this amplifier is 50, but its actual gain linearity and stability are not specified to a tight tolerance. To determine the current amplifier parameters and thereby increase the accuracy of current measurements, three calibration signals (with 0.2 percent stability) are amplified and telemetered in each commutator. These signals can thus be used by postmission processing for a continual in-flight calibration of the current amplifier. Telemetry plots of these calibration signals show that the gain of the and AESP current amplifiers was reasonably constant over the mission. SC-2, a new system of "calibrating" these signals in percent, not telem-
ESP For
etry volts, was used. The zero point on the scale is set at the unit flight acceptance test (FAT) measured value. The change in voltage of a given signal is divided by 5 volts (full scale) to convert to percent. Thus, it can be said that the current calibration signals, in gener=_l, have increased by 0.6 percent since unit FAT (see Table 5.4-4). This percentage change is not passed on to the current signal measurements, however, since the in-flight calibration process removes this effect c_,rr_pletely. Only if the every-frame correction were replaced by a constant correction (as has been recommended) would this variation be passed on directly as an error to the current measurements. The range of variation is 0.2 percent for AESP and 0.3 percent for ESP. error on a 35 ampere For the current latter, shunt. this would mean a 100-milliampere
':"Itust m
be
noted
that this value
was
not calculated
in time
to be
used
in
processing the mission plots. Thus, the pressure plots of P-l and P-Z in Section 5.6 are generally too high, being based on an assumed 4.85-volt reference. Correct values occur at commutator assessments, when the ESP commutator (which has a telemetered reference voltage) was used.
5.4-6
TABLE 5.4-3.
CALCULATION
OF AESP REFERENCE VOLTAGE
AESP Reference Voltage, BCD
GMT, day:hr:min Z63:11:05
Calculation Mode Signal M-3 M-4 M-7 Telemetry Value, BCD 910 380 338 9O4 X i000 X I000 X I000 910 904 38O 378 338 - 336 X:Vre f (AES
of P)
Z63:11:09
M-3
1006.6
M-4
378
1005.3
M-7
336
1006.0
S-I reference voltage
I000
263:11:15
M-4
377 X 996 380 : _77
S-I
996
1003.6
264:02:40
5
M-3 M-4 M-7
675 384 5O8 X 1002 X 1002 X 100Z 675 - 668 384 381 508 504
264:03:02
4
M-3
668
1012.5
M-4
381
I009.9
M-7
504
1009.9
S-I
1002
5.4-7
TABLE
5.4-4.
SUMMARY OF SIGNAL
CURRENT DATA
CALIBRATION
Flight Signal EP-18 Function E SP, 90% 0.52 0.36
Data, Remarks Mode Mode Mode Mode 4 2 I i at 4400 bits/sec
percent 0.62 - 0.54 - 0.43 0.12
EP-19
ESP,
50%
0.6
Constant
EP-Z0
ESP,
10%
0.64 0.60 0.32
Mode Modes Mode
4 Z and i bits/sec
I at 4400
EP-27
AESP,
90%
0.8
EP-38
AESP,
50%
Prelaunch Near midcourse
EP-29
AESP,
10%
Prelaunch Launch
and
after
midcourse
to midcourse
5.4-8
The ESP current calibration flight data are presented in Figure 5.4-2. EP-19 was not shown since it was relatively constant. It can be seen that the signals vary not only from mode to mode (this was known previously and is part of the spacecraft signature) but also are changed considerably at 4400 bits/sec, an effect not previously reported. In Figure 5.4-3, signal variation over a 7-minute period surrounding midcourse is shown. The data involved here have been averaged such that one point represents a 7.5-second (30 frame) interval, but an examination of unprocessed data in that same interval shows only slightly more variation than the 3 to 4 bcd shown in Figure 5.4-3. Investigations thus far have not shown any direct correlation between frame-to-frame variations in current calibration signals and changes in other current calibrations. But mode- and bit rate-dependence and long term changes in amplifier gain do exist, and thus an on-going calibration should be retained to avoid the errors associated with using constant factors for the entire flight. 5.4.4.4
Temperature The errors Measurement in temperature Errors at 4400 Bits/Sec at 4400 Data Rate result
measurements
bits/sec
from insufficient settling time for the constant current source used to convert resistance (which is proportional to temperature) into telemetry voltage. The output capacitor on the constant current source, when unloaded, charges to about 6.8 volts. At the highest data rate, this capacitor does not have time to restabilize at the lower voltage (typically 2.5 to 3.0 volts) before the particular data channel is sampled. At if00 bits/sec, there is four times as much time for settling, and no inaccuracy apparently exists. In the SC-I Final Performance Report, a detailed discussion determined which temperature measurements would be most in error. A table was also presented which listed the range of true values for a given telemetered value. It is the intention of this analysis to carry the investigation one step further: to present a means of reclaiming the true temperature values from 4400 bits/sec data. Application of this technique is of particular importance to SC-2 analysis, since the midcourse failure occurred when the bit rate was 4400, and since temperatures are an important factor. Thus, the two vernier line sensors (P-4 and P-8) will be used as an example for the reconstruction. The first and only prerequisite perature signal is that it be changing the better. The vernier line signals 5.4-4), since after the yaw maneuver to proper interpretation of a temunidirectionally with time, the faster satisfy this requirement (see Figure they were warming due to solar
radiation. It is necessary to revise the temperature correction table to stress the fact that the only values transmitted are those listed in the table: the intermediate values never occur (see Table 5. 4-5). The transmitted value uniquely limits the true value within the stated range (plus uncertainties associated with this uncontrolled design process). results can be qualitatively explained by studying the digitization the These process.
5.4-9
Figure
5.4-2.
Mode
and Signals
Bit
Rate (EP-18
Dependence and EP-20)
of
Current
Calibration
5.4-10
1 1
J
_
H
it, _r
H
f_
1
/
t_
1 1
a) 90 Millivolts (EP-18)
;lit t
NFHIIHI !!!!]!!!]!
I!LI
7_
ill I,
:_!i!H!!
,/_ !,
lillll!_l
_++++vfi ÷ i
iINtNtN
-,_u.
b) Figure < iJ 5.4-3. Current
10
Millivolts Signal
(EP-20) Variation Over 7 Minutes at Mi
Calibration
_._
5.4-11 I: F_ Fi_tll I/
-li
:ourse
FO_
FI_IVlF,
Line
i
b) Figure 5.4-4. Vernier
Line
2 at Midcourse
Line
Temperatures
5.4-13
TABLE
5. 4-5. RELATION TEMPERATURE
OF TRANSMITTED MEASUREMENTS
AT
AND TRUE 4400 BITS/SEC
VALUES
OF
Transmitted Value, 1023 1022 1020 1016 1008 I000 992 976 968 96O 944 936 928 912 904 896 864 848 84O 832 816 8O8 8OO 784 776 768 736 720 712 704 688 BCD
Range of Corresponding True Values, BCD 997-1023 992996 986991 977985 964976 961963 941960 932940 929931 902928 9O0901 898899 878897 869877 867868 830866 817829 808816 8O67797787767577487467026986896876626608O7 805 779 777 775 756 747 745 701 697 688 686 661
Fransmitted Value, BCD 680 672 656 648 64O 608 592 584 580 576 56O 552 548 544 532 528 520 516 512 480 472 468 466 464 460 457 456 452 450 448 440
Range of Corresponding True Values, BCD 657-659 640-659 631-639 628-630 598-627 589-597 573-588 569-572 568 549-567 544-548 540-543 558-539 525-537 523-524 515-522 510-514 5O8-5O9 474-507 468-473 465-467 463 -464 462 456-461 455 454 450-453 448-449 447 436-446 435
5.4-14
Here, a series of yes/no decisions are made, starting with the question of whether the measured voltage is greater than 2.5 volts (half-scale). Each succeeding decision involves an incremental voltage half that of the preceding, until the final tenth step refers to only 5 millivolts. If the measured voltage is higher than its final settled value at some decision point in time, the wrong decision may be made, and then there is no way that the following decisions can correct for this, since
1> [ (1)(0.5)+ (0.5)(0.5).... ]
N 1> _ i=l The preceding discussion and Table 5. 4-5 apply only to temperature measurements for which the current source output is initially charged to its positive maximum, which is the case for P-4 and P-8 in mode I. The correction process itself is extremely simple, granted that the above is valid. For example, assume that the telemetered value of a signal has been 512, and then changes to 516. At the time of change, the true value of the signal must be 507-8 bcd, since that is the borderline between the two transmitted value states. In the example (Figure 5.4-5), the corrected curves were of the vernier constructed line signals from a series of (0.5) i for any finite N
points at the measured state transition times. line temperatures start to cool, thus violating
After the midcourse burn, the the requirement of unidirec-
tional change. From this point on, the curve reconstruction becomes highly speculative, and no great faith should be put in the curve shapes or peak values. At 05:01:11, data are available at if00 bits/sec, which requires no correction. The most difficulty was experienced in fitting a reasonable curve to the vernier line 2 data. This signal remained at 576 bcd through 05:00:55, requiring that the true value should have been above 548 bcd at least up to that time. But this situation requires a sudden drop of over 6 bcd (5.4°F) in the following 16 seconds, which is hard to explain in physical terms. In conclusion, the reconstruction method give accurate values up to the 548 bcd level, but is subject to the analyst's judgment. _.4.4.5 Telemetry Data quality Bit Stream before Characteristics correction was excellent, with a presented is believed to the remainder of the curve
the midcourse
very low word error rate even though the data rate was generally maintained at II00 bits/sec. After midcourse, the situation changed drastically. On almost every mission plot in this report, many extremely spurious values can be seen on that part of the curve after midcourse. It is therefore of great interest to present a possible explanation for the rapid telemetry data quality that began when the spacecraft started (see Reference I). deterioration to tumble of
5.4-15
a)
Line
1
i:ti:
:_ i::i[}!:!'i:!i!_:_, ¸
P:: I 1:1
!
b) Figure 5.4-5. Comparison Vernier Line
Line
2 a_d Uncorrected
of Corrected Temperatures
5, 4-16
current)
Investigation of an unexplainable value of EP-9 (battery discharge during burn 27 led to discovery of a general mechanism that could
explain most, if not all, errors that occurred in telemetered data throughout the period when the spacecraft was tumbling. A definite pattern was established which indicated that the telemetry quality, i.e., bit error rate, was not constant in time but varying as the spacecraft tumbled. For most of the tumble period, the data were good, but once a cycle, the data became very bad. Consequently, the average word error rate was low (at ii00 bits/sec) which would normally lead to the conclusion that errors of more than i bit per word were unlikely. This, however, is an erroneous conclusion since the data quality was varying with time. Analysis of a frame-by-frame 16-second period surrounding the dump of the telemetered data during time of engine ignition (265:03:47:59}
a
revealed that very few parity errors were occurring per frame of data. Some frames had no parity error, and others had, at most, four parity errors per frame. Initially, this low word error rate led to the tentative conclusion that the data point of interest (EP-9) was valid and not a result of 2 bit errors. However, it was noted that, in the case where multiple parity errors occurred in a single frame, the errors were grouped together rather than distributed throughout the frame. This observation led to a more detailed study Table which 5.4-6 clearly contains showed that the errors were a tabulation of the position, occurring in a cyclic manner. in 16 frames of data sur-
rounding in frame and hence
the time of ignition, of the noted word errors. (Ignition took place 0. ) The asterisk in the table denotes suspected double bit errors, no parity errors. The number of words between the observed
word errors is also tabulated. .As can be noted, the errors were occurring approximately every 83 words before engine firing, and the period began to change right after firing. In the last two frames of data, the errors appeared to be occurring approximately every 74 words. .A change in rate corresponding to the time of engine firing is clearly indicated. The words containing double parity flagged data assuming bit errors were periodicity. A located double by extrapolating bit error was
the bad
detected by noting the values of the same words in surrounding frames and comparing them to the suspected value. The binary representations of the words were compared and, if two bits were different, it was concluded that a double bit error occurred. EP-9 of interest: (word 72) had the following binary values in the three frames
Frame
-l
1
0
0
0
0
I
I
0
0
0
0 +I
i l
i 0
0 0
I 1
0 0
0 1
i 1
0 0
i 1
I 1
(suspect
data)
BIT
2
BIT
6
5.4-17
TABLE 5.4-6.
WORD ERROR POSITION SUMMARY ARCUND BURN
Word Position of Error 46 28 II 94 82 83 83 83 82 83 82 83-84 83-82 82 8O 8O 8O 76 73-74 70-71 76-78 73-74 74-75 74 Number Between of Words Errors
?7
Relative Frame Number -8-`',' -7 -6
- 5-`:' - 4-'_ -3 -Z -i
77 59 42 24 7/8 9O
0 '_"
72 52 32 12 88
I ::" 2 3*
4 5 6 ':"
61/63 31/32
7/811i 81/8z
7 8 _'"
56 30
-':'Frame containing
suspected
words
with
Z bit word
errors.
5.4-18
Since the current is expected to change at the time of engine firing, the word in the preceding frame (-1) was not used for comparison. The word in the frame following engine firing (+1) is thus considered a most likely value. As can be noted, the suspect data in frame 0 shows probable bit errors in bits 2 and 6 relative to the word value in frame + 1. For further temperature change during evidence channels a single that double bit were investigated frame of data. errors to Since were indeed occurring, find a telemetered temperthe actual temperature 82)
several ature
could not change significantly in 1 second, it was felt that any large change would definitely be a telemetry problem. Temperature channel P-9 (word was found to have changed 22 degrees in frame +6. The BCD value in the surrounding frames of data was 574 counts, with a change to 550 counts occurring in frame 6. The binary representation of the two BCD values is as follows: BCD BCD 574 550 1 1 0 0 0 0 0 0 1 1 1 0 1 0 1 1 1 1 0 0
BIT
6
BIT
7
A difference of bit error indeed agreed precisely Fable 5. 4-6. Examination from 0.82 DSIF second
two bits was found, leading to the occurred in the word. Also, the with the predicted position based
conclusion that a double word position of P-9 on the data contained
in
of
the
spacecraft data, second
tumbling
rate, that
which
was
determined period bound was the
automatic at 03:47:54 firing bits/sec. being in
gain
control and 0.74
indicated at 04:17:28.
the tumble These times
time of engine mode 1 at 1100 ing in one word The data
(03:47:59). This transmitted 5.4-6
The telemetry mode contains every 0. 0i show about 83
mode during 100 words per second. words between
this period was frame, result-
Table
word
errors, the
or 0.83 second, prior These data correlate time of interest. It were caused by the
to engine firing and almost exactly with is therefore concluded tumbling spacecraft.
0.74 second after engine firing. the spacecraft tumble period at that the observed word errors
In retrospect, this result is not too surprising since the RF link was experiencing considerable variations due to the tumbling spacecraft. Signal levels at the ground receiver were varying due to spacecraft omnidirectional antenna gain variations, and the DSIF carrier tracking loop was experiencing large errors due to the frequency variations resulting from the transmitting omnidirectional antenna spinning in space. It is not clear at this time if the bad data were caused by low signal-to-noise ratios, by a momentary loss of carrier phase lock, or by a combination of both effects. There is no doubt, however, that the periodic bad data were caused by the effects of the tumbling spacecraft on the RF link. 5.4-19
REFERENCES
F. K. Number
Rickman, "Discussion 27," IDC 2292/140,
of Behavior l December
of EP-9 1966.
during
Burn
5.4.6
ACKNOWLEDGMENTS
This section was coordinated to F. K. Rickman for the analysis spacecraft tumbling rate.
by R. H. Leuschner. Recognition is due of the correlation between word errors and
5.4-20
5. 5
FLIGHT
CONTROL
5. 5. l
INTRODUCTION The principal requirements control, accurate angular of the Surveyor flight control maneuvers, precision velocity system correcare
attitude
tions, and soft lunar landing. In order to accomplish control system utilizes such hardware as gyros, gas liquid fuel engines, optical sensors, timing devices, tion sensing 5. 5. 1. 1 mechanisms. Control
these functions, the jets, solid fuel engine, radars, and accelera-
Attitude
Attitude control is accomplished by two basic types of active control systems. During coast phase, a bang-bang type of attitude gas jet system is employed which utilizes a novel technique of artificial rate feedback for loop stabilization and, during periods of large moment disturbances such as the main retro phase, the throttle-controlled vernier engine system is used. The error signals required for controlling the propulsion systems are derived from optical sensors or rate integrating gyros which are mounted on the spacecraft in such a way as to provide a three-axis coordinate system. During coast phase, where the gas jet system is used, two modes of operation are available. One choice is celestial referencing, using the sun and Canopus, and the second is self-contained inertial referencing (gyros). The first mode is used to establish accurate attitude, and the second mode is generally instance 5. 5. 1. 2 used occurs Angular when momentary inertial reference during an attitude maneuver. Maneuvers is desired; such an
The vers which given time jet system. 5. 5. 1.3
rate integrating gyros are also used for accurate angular maneuare accomplished by precessing the gyros at precise rates for intervals and slaving the spacecraft to the gyros through the gas
Velocity Midcourse a system
Correction velocity consisting correction of three capability vernier of exact engines, magnitudes a precision is timer, pro-
vided
by
5.5-I
and an accurate acceleration sensing device. The difference between the commanded acceleration level and the output from the accelerometer provides the error signal that commands the vernier engines to the required thrust levels. The constant acceleration and variable time concept used by the Surveyor flight control system provides the flexibility of choosing velocity corrections from 0 to 50 m/sec. 5. 5. I. 4
Soft Landing
Surveyor's soft landing capability is provided by a sophisticated technique utilizing radars for computing velocities and range. • The range information is then used by an on-board computer to provide veloclty commands to the vernier engine system according to an approximate, constant acceleration, VZ/R function. The velocity information is used by the vernier engine-attitude control loop to produce a near-gravity turn descent the spacecraft thrust axis to the true velocity vector. The velocity tion is also used, along with velocity commands, to generate error for the velocity control loop. by caging informasignals
In order to provide low velocity for the soft landing phase, approach velocity is decreased by a solid fuel rocket engine during the initial portion of terminal descent. The spacecraft attitude during this phase is inertially stabilized by the gyro-vernier engine control system. 5. 5. i. 5 Mission Performance
Surveyor II successfully performed all comrr_anded maneuvers from launch to midcourse vernier engine firing, including Centaur separation, sun acquisition, star acquisition, coast mode, and premidcourse maneuver (Table 5. 5-i). Failure of the leg 3 vernier engine to fire at the midcourse command caused immediate loss of attitude control and spacecraft tumble. Attitude control was not regained during any of 40 postmidcourse vernier engine firing attempts.
5. 5. i. 6
Analysis In order to properly items was prepared evaluate the spacecraft performance, (see subsection 5. 5. I. 7). The items a list of are cate-
analysis
gorized under major mission phases (such as launch through separation, coast phase, and midcourse correction) for easier identification and performance evaluation. A time and events log is presented in Table 5. 5-I, and a summary of results is given in subsection 5. 5.3. In subsection 5. 5. 2, a table of anomalies is presented along with a brief description of each anomaly. Subsection 5. 5.3 also contains the conclusions and recommendations of the investigation, and subsection 5. 5.4 contains the analysis effort.
5.5-2
TABLE
5. 5-1.
TIME
AND
EVENTS
LOG
Item Launch Separation Start of sun (2-inch motion) by M-9) minus illuminated lockon sun (cruise (manual (start) (stop) (manual on inertial cruise lockon) and roll (positive) on) roll (plus yaw)
Command
DSIF GMT, day:hr :min:sec 263:12:31:59.8 12:44:27. 12:45:18.3 12:47:41. 12:48:13. 3 0 4
Mission Time, hr:min:sec
(indicated acquisition sun sun sensor
00:12:27. 00:13:18. 00:15:41. 00:16:13. 06:05:34 06:37:38 06:39:57 06:54:24 09:03:22 09:07:23 14:35:43 14:37:05 14:38:06 14:38:31 14:41:07 16:12:00 16:16:05 16:21:38 16:22:20 16:28:01 16:28:02 16:31:48 16:42:29 18:56:25 19:18:03 35:01:23 36:56:11 38:07:14 39:15:96 40:24:12 41:11:19 43:33:12 45:02:17
6 5 5 2
Acquisition Primary Start of
sensor mapping of roll
star
0714 0704 0716
18:37:34 19:09:38 19:11:57 19:26:24 21:35:22 21:39:23 264:03:07:43 03:09:05 03:10:06 03:10:31 03:13:07 04:44:00 04:48:05 04:53:38 04:54:20 05:00:01 05:00:02 05:03:48 05:14:29 07:28:25 07:50:03 23:33:23 265:01:28:11 02:39:14 03:47:56 04:56:12 05:43:19 08:05:12 09:34:17
Termination Canopus Gyro Gyro Canopus Gyro speed Next Next Next Gyro Start Start Unlock Thrust Midcourse speed drift drift
mode lockon)
acquisition check check lockon check gyro gyro gyro check off
mode mode on
on
0700 0704 0716 0221 0222 0222 0222 0223
premidcourse premidcourse roll _ actuator on
sun plus and
and yaw
roll (110.
(plus
75.3
degrees)
0714 0713 0605 0727
5 degrees) helium
pressurize
power velocity
correction
3617 0721
Rate Inhibit
mode gas
on jets burn 1 2 3 4 5 6 7 (midcourse (midcourse (midcourse (midcourse (midcourse (midcourse (midcourse burn burn thrust thrust thrust thrust thrust thrust thrust (high (high levels) levels) levels) levels) levels) levels) levels) thrust) thrust)
0701 0707 0721 0721 0721 0721 0721 0721 0721 0721 0721 0730
2. 0-second
Approximate Approximate Emergency
2. 0-second 20-second AMR signal
command
5.5-3
5. 5. I. 7
Analysis
Items
--Flight
Control
System
The following list constitutes the postflight performance analysis effort for the flight control system. The degree to which the individual items were investigated depended on the impact of that parameter on the overall flight control performance assessment. 1) Prelaunch a) b) 2) Temperatures Nitrogen through weight (nitrogen pressure telemetry calibration)
Launch a)
Centaur
separation
Centaur Rate
separation stabilization verification
Separation rate magnitudes Time to stabilize Total angular Nitrogen gas b) c) 3) Sun a) b) c) d) e) f) g) Rate mode latch excursion utilization reset anomaly sensor
Response acquisition Automatic Maneuvers Roll Yaw Acquisition Response Nitrogen
of Canopus
sun
acquisition
verification
time of Canopus gas sensor
utilization
5. 5-4
4)
Canopus (star) acquisition a) b) Acquisition Star maps Star intensities (predicted/observed) Effect of pitch/yaw limit cycle on map Mean roll rate Dynamic telemetry calibration Other stars identified c) Sensor performance Field of view setting Sensor effective gain Lockon characteristics d) Acquisition maneuver verification
Roll control system performance Manual lockon required Average roll rate Nitrogen gas utilization 5) Coast phase attitude control a) Limit cycle Inertial mode
Frequency Amplitude Optical mode Frequency Amplitude b) Attitude control errors Noi se Tracking c) d) Gyro drift Gas jets Nitrogen gas utilization Thrust level
5.5-5
6)
Pren_idcourse attitude maneuver a) b) c) d) Timing accuracy Maneuver rates Attitude maneuver error Nitrogen gas utilization
7)
Midcourse velocity correction a) b) c) Detailed description of spacecraft motion (in terms flight control variables) Roll actuator performance Gas jets Reduction in tumble rate Nitrogen gas utilization d) Vernier engine transients attitude transients vernier predicted from spacecraft of
8) 9) I0)
Postmidcourse Retro firing Postmission a) b) c) d)
engine firings
tests and analyses and thrust command telemetry characteristics
Gyro error
ZZ-volt thrust phase bus current during midcourse Dynamic versus static calibration mapping telemetry signal Computer simulations of Canopus sensor
l I)
Total nitrogen gas utilization
5.5-6
5. 5. Z
ANOMALY
DESCRIPTION
The described heading). 5. 5. 2. 1
flight control anomalies that occurred during the mission are briefly below (anomaly details are presented under the appropriate
Rate
Mode
Latch
Reset
During
Launch
At a time corresponding approximately to the generation of the legs extend signals, the flight control programmer logic was reset from the normal rate n_ode to inertial mode. This condition remained until electrical separation from Centaur, at which time 30 seconds. When data were restored, returned to the rate mode, and no other remainder of the mission. in subsection 5. 5.4. 2. 5. 5. 2. 2 As Canopus in the Lockon SC-I A detailed data were lost for approximately the programmer logic circuit had anomalies were observed for the discussion of this anomaly is presented
Signal
Failure it was necessary to operate the Canopus
mission,
lockon circuits manually because the sensor did not generate a lockon signal. The failure was not completely unexpected because the sensor gain was increased intentionally by approximately 20 percent, based on calibration data, to compensate discussed further for possible sensor window in subsection 5. 5.4. 5. fogging. The anomaly is
5. 5. 2. 3
Midcourse
Velocity
Correction
Failure
The midcourse velocity correction attempt was characterized by vernier engine 3 failure to ignite and subsequent tumbling of the spacecraft which resulted in saturation of the telemetered gyro error signals in a minus pitch, plus yaw, and minus roll direction. The approximate tumble rate at vernier engine shutoff was 448 deg/sec. Ignition failure was confirmed by telemetered strain gage and engine temperature data. Subsection 5. 5.4. 8 contains a description of this anomaly. 5. 5. 2.4 Late Shutoff of Leg 1 Vernier Engine During Postmidcourse Burn for
Review of flight control data, in conjunction with strain gage data, the postmidcourse vernier engine firings shows that on burn 27 the leg 1 engine continued burning after the commanded termination. This phenomenon is of interest due to the possibility ticularly should such burning occur is discussed in subsection 5.5.4.8. of losing control during terminal of the spacecraft, pardescent. This anomaly
5. 5.3
SUMMARY
AND
CONCLUSIONS
5. 5.3.
l An
Performance SC-2 flight
Summary control performance summary is presented in Table
5.5-2.
5. 5-7
TABLE
5. 5-Z.
FLIGHT
CONTROL
RESULTS
Controlling Specification
Specification Value R_sults Comments
Prelaunch R,,II 1:htch Ya_ 4.5 pounds 4 5 pc-raids 17Z I70 !72 3°K 2 _ F Time was 12:32
Proper
gyro
temperature
control
_)'_
Verification
of
N 2
loading
Centaur Time than
separation required 0. I deg/sec of angular rate at Z24g10E (3.32. I) to null rates to less 224510E (3. 3 3. 3) 0.052 pound drift N 2 usage Design 224510E (4.3.15) 0.0012 0 _ _ _--_ m 0
o
_o_
n_ ml U ml_O
-
_
_
I
__°_n
0 _ 0
-o ooO
0
ooo o
_
_
o
o c_
,A
,.D
,if
_o_
z_ 0
I-.-I
_" 0 _.,_ m _
_
_
r_
I.--I
_
>
_'__
m oO
,F-I
_d
1---
C_
C)
O_ nZ_m o
.el
o
u
_
u
_
___
o_
_
_0_._
_
°._
0
_
_
.e"l ._ _._
r_,__
m_
cQ
o
.,.-i C;
0 o
r_ >
.r-I t...t
u
-,"_
_o
0 4_
0
_
r/?
u'l 0
-_'-I
"0
0
_
_ u o
• rt _/)
<_o
4_
o r_.,_ _,
>
,o,..1
_o
Ot_._ U?
5.5-47
5. 5. 4. 5
Coast
Mode
The
three-axis
cold
gas
attitude
control
system
is designed
to main-
tain an optical or inertial reference during the nonthrusting portions of the Surveyor flight. The spacecraft pitch and yaw optical references are provided by a narrow field of view sun sensor; the roll optical reference is provided by a Canopus sensor whose field of view is 5 degrees latitude and 8 degrees longitude. three body-fixed rate The spacecraft inertial integrating gyros. references are provided by
The actuators used in the coast mode are the cold gas jets. The on-off operation of these jets, plus the deadbands built into the system at the gas jet amplifiers, cause the spacecraft to function in a three-axis limit cycle. In the steady state, line within the three-dimensional cepting a bounding plane, a within the deadband along a analogous to the motions of dimensional, planar-sided, determines that the velocity amounts (as caused by a jet one plane. The motions are velocity conditions. the Surveyor attitude coasts deadband of 8, _, %b space. along Upon a straight inter-
gas jet pulse is emitted_ driving the system back new straight line. These motions are entirely a ball bouncing internally within a closed threesix-sided polygon, wherein the law of reflection components of the ball change by discrete pulse), these amounts being constant for any aperiodic and are a strong function of initial
The non-g sensitive drift rates of the integrating rate gyros were measured during Mission B by slaving the spacecraft to the drifting inertial references and observing the drift rates by means of the telemetered optical references. The were I) Z) 3) 4) principal items as follows: Limit Sun cycle and star gas of analysis for the flight control system coast
phase
frequency tracking used
and errors
amplitude and tracking noise
Nitrogen Results
of gyro
drift measurement the coast reference for phase, together with their correspondin Table 5. 5-1Z. Table 5. 5-13 is a phase.
The major events ing times, are presented summary of the analysis
for for
results
the coast
5. 5-48
TABLE 5. 5-12.
Event Sun lockon
MAJOR EVENTS AND TIMES
GMT, day; hr: rain: sec 263:12:48:13 Command Automatic sequence
Star
map
(begin) acquire (manual 1 1 maneuvers lockon)
263:18:37:34 263:19:11:57 263:19:26:24 263:21:35:22 264:04:44:00
0714 0716 0700 0704 0714
Canopus Begin End Begin
gyro gyro
drift check drift check
premidcourse
Conclusions
and
Recommendations
i)
Limit cycle behavior was as predicted except that the roll limit cycle occasionally exhibited the tendency to double-pulse at each side of the deadspace (ideally, it should always single-pulse). About Z0 percent of the fuel consumed during the sampled limit cycle period was a result of double pulsing. However, since the fuel penalty was very low (about 0. 00Z pound) and the limit cycle amplitude was unchanged, this additional pulsing was readily tolerated. Sensor noise performance. did not affect the coast mode control system
Z)
3)
Extrapolated fuel consumption was about what (after allowing for the additional double-pulsing Details
was predicted of the jets).
Analysis
Limit Cycle Frequency. The three-axis by a crosscoupling of the torques resulting from coupling is shown by the following:
limit cycle is characterized a gas jet pulse. This
1)
A
pulse
from about
the the
No.
1 gas
jet pair yaw axes.
causes
a change
in rotational
velocity
roll and
z)
A pulse from the No. 2 or 3 gas jet pair causes rotational velocity about the pitch and yaw axes.
a change
in
5. 5-49
TABLE
5. 5-13.
SUMMARY
OF
RESULTS
Limit Mission Inertial: Optical: 61 64 B sec/pulse sec/pulse
Cycle
Frequency Predicted 117 80 sec/i)uls, sec/pulse _ (Reference (For noise: Figure 1_ 7, Canopus Reference 1). page 3) sensor 7,
Limit Mission Inertial, Roll Pitch Yaw Optical, Roll Pitch Yaw degrees: 0. 441 0. 45 0. 37 Sun Mission Roll Pitch Yaw null: null: null: from to B -0. -0. -0. 08 07 10 DSIF and degrees: 0. 47
Cycle B
Amplitude
(Single
Axi
s) Predicted
0. 44 From DSIF drift data, test 0. 44 0.44 gyro
0. 42 0. 43
0. 44 From DSIF 264:00:03:01 264:01:00:00 data, to 0.44 0. 44 Star Tracking Errors, degrees Specification ±0. +0. +0. i0 i0 10
Nulls obtained data: 264:00:03:01 264:01:00:00
Sun The sun no effect
and
Star
Tracking levels
Noise were low enough to have
and star error signal noise upon limit cycle performance. Fuel Consumption
Mission 0. Predicted 0.
B 020 pound (For 16 hours of limit cycle operation}
018
pound
(nominal)--See Budget'S-limit cycle Gyro Drift measured Measurements in flight* Pitch 78 263:21:35:25. 0.24
Reference corrected for period
a
8, "Fuel 16-hour
Drift Check 1 *Taken from
rates
(deg/hr) Roll -0.
are
as
follows: W aw i. 09
263:19:26:00
to
5.5-50
Consequently,
limit
cycle
frequency
determination
is simply
a matter
of simultaneously examining the pitch, yaw, and roll error signals and counting slope changes, making sure that a pulse is not counted twice because of the system crosscoupling. A 148-minute sample of the optical limit cycle (Figure 5. 5-12) had a mean time between gas jet pulses of 64 seconds. The data were taken from DSIF tapes of 264:00:30:00 to 264:02:58:00. A 77-minute sample of the inertial limit cycle (Figure 5. 5-13) had a mean time between gas jet pulses of 61 seconds. The data were taken from DSIF tapes of 263:19:Z6:00 to 263:20:43:00. Limit Cycle Amplitude. The roll optical and inertial deadspaces determined from the roll error sensors during limit cycle operation. the roll optical and inertial deadspaces were consistent throughout the period.
were Both
sampling
In pitch and yaw, there is an additional measurement consideration. A No. 2 or 3 gas jet will fire whenever the sum or difference of the pitch and yaw error signals exceeds either's single-axis deadspace voltage. Hence, a pure pitch or yaw deadspace measurement can only be made when one or the other is at null. This point will result in the maximum possible swing of the error signal which Table 5. 5-13 were the maximum indicated sample period. is met at null. total deadspaces The values observed recorded during in the
Tracking Noise. Because the single-axis deadspaces are approximately equal for both inertial and optical modes and because the mean time between gas jet pulses was about the same for both the optical sample and the inertial sample, it is certain that optical sensor noise had no harmful effect upon limit cycle operation. errors, recorded in Table limit cycle operation. 5. 5-13,
were
Tracking taken to be
Errors. The tracking the optical nulls during
Fuel Consumption. Both samples had double-pulsing at a deadband boundary which accounts for about 2Z percent of the gas jet pulses. (Doublepulsing is detected from the telemetry signals by noting the magnitude of the error slope change at a boundary. ) The predicted nominal limit cycle fuel consumption (corrected for a 16-hour coast mode) is (0. 075 pound) (16/66) -- 0. 018 pound (see Reference 8, '_Fuel Budget'S). The overage is thus (0. 020 - 0. 018 pound) = 0. 00Z pound. This ii percent overage is accounted for by the double-pulsing noted above.
5. 5-51
Gyro
Drift drift check was The drift rate (FC-5 and FC-6) made values and from 19:26:24 obtained from Canopus sensor until 21:35:22 on plots of the primary error signal were as
Z0
A gyro September. signals s:
sun
error follow
Roll Pitch Yaw
gyro gyro gyro
(S/N (S/N (S/N
72) 70) 51)
= = =
-0. 78 0. Z4
deg/hr deg/hr
I. 09 deg/hr is shown in
Table
The non-g sensitive 5. 5-14. The data do
drift history of the three gyros not appear to predict a trend.
The in-flight, i. e. , zero-g, Mission B gyro drift rate values cover range of values that compares very favorably with the limits of ±l deg/hr placed on earth-based measurements of non-g sensitive drift rate. Since in-flight conditions are zero-g along all axes, as compared to earth-based conditions of zero-g along only two of the three axes, there is no valid method of directly comparing in-flight zero-g and ground-based non-g sensitive drift rates. Based gyro responses that the gyros Gas upon these in-flight drift rate values obtained observed during all portions of Mission B, were operating in a normal manner. Thrust the Level "Surveyor Functional Requirements Specification, " along with the it is concluded
a
Jet
Although
224510E (Reference 5) does not directly dictate the minimum allowable thrust level, it does infer these levels by specifying the minimum allowable gas jet torque values, as presented in Table 5. 5-15. It is apparent pound. The pound. from these data that the minimum allowable gas jet system was designed for a nominal gas jet thrust thrust value
is 0. 05Z of 0. 057
In References 9 and i0 a method is proposed whereby the gas je[ thrust level can be determined from the time response of the gyro error signal received during a roll maneuver. This method also mentions that a weighting factor may be required for the basic equation derived in the references. The equation is
T-
z Rt P
c
5. 5-52
'_
]
I:
]i ::
i; 111: i':: I I_
t:?t=t +-+-_-)--_I ' i "
H
::
*+
+
I--_+
4_--
I:::
;-;.4......h _-
t'
I :: + ._+_ .-t F---_,Z:
i ¸,
......... '..... 7 ::"
!
I!
:il :.:!i
: I I'
......
p.:
_
,
........ i::i:: 11:,/:_
'+"_ ""U: .....
IH --ttt:' ..... ?
: "
.:
;..[ _ I j:
: .... :
:
.. .: ;
':1;1
.:ski[ Jl; li::k
:::] :::i '-:LL
ii. ::: --:_
!1
...+ .....
..........
:::i
If
I
.
,+
+ ........
........... b-b..... it
:....
LK -! ......... 1 :id!i,: :. -:1i-- : ..... ....
!
":!':';
t:'"
:,
"
'
TIM[
IMIN}
Figure
5. 5-1Z.
Optical
Limit
Cycle
I
5. 5- 53 FOLDO,U]_
)
P_ECEDING
PAGE
BLANK
NOT
,,_.,_ ......
32.
36.
qO.
_lq.
qS.
5C-2 TRRN$1T ORTR, MIN
[I.T.:263/1
Figure
5. 5-13.
Inertial
Limit
Cycle
5. 5-55 ¥OLDOg_
Fl_¥_
/
P?,ECEDING
PAGE
_,:__";'_,,. NOT _
FILI',,L.,-..
TABLE
5. 5-14. SUMMARY DRIFT RATE
OF GYRO MEASUREMENTS
NON-G
SENSITIVE
Pitch On-Time, hours I00 115 139 235 260 296 309
Gyro
(S/N
70)
Roll On-Time, hours 78 118
Gyro
(S/N
72)
Yaw On-Time, hour 66 If6
Gyro
(S/N
51)
Drift Rate, deg/hr
Drift Rate, deg/hr
s
Drift Rate, deg/hr -0.02 - 0. 487 -0. 13 -0. 73 -0. 30 -1.24 E-W -0. 57 N-S -0. 58
-0. 09 -0. 19 0 0. 38 -0. O65
128 152 248 322 334
0 -0. 33 -0. 95 -0. 055 -0. 055
126 149 162 207 211
321 344 357 402 406 432 474
-0. 065 -0. 05 0.17 0.43 0. O95
357 370 415 487
-0.2 -0. 48 -0. 19 -0. 762
237 279
-0. 38 -0. 571
0.19
0
SC-2 flight Note: 0.24
SC-2 flight group, S/N 9. P/N -0.78 23500-9, S/N
SC-2 flight i; inertial I. 09 reference
Flight control sensor unit, P/N 235100-1,
5. 5-57
TABLE 5. 5-15.
MINIMUM ALLOWABLE TORQUE VALUES
GAS JET
Spacecraft Axis Roll Pitch Yaw
Minimum Torque Requirement, in-lb
4. O0
Moment Arm, 77 45 68 inch
Number of Gas Jets
Minimum Value,
Thrust pounds
0. 052 0. 047 0. 052
4.25 7.00
From an analog computer simulation the average weighting value was determined various parameters in the previous equation I
Z
program of the gas jet system, to be 0. 85. The values for the are as follows:
= =
189
slug- ft 2
R
6. 4 feet
_c t P
: :
0.500 5.5
deg/sec seconds
tp represents the time from maneuver gyro output reached its first maximum. shown in Figure 5. 5-14. Multiplying of 0. 85, the corrected time is 4. 7Z for roll was T = 0. 056 pound. 5. 5. 4.6 Premidcourse Maneuvers
command This
initiation to the period time was approximated,
when as factor level
the value of tp with the weighting seconds. Thus, the gas jet thrust
In order
to accomplish
the required roll maneuver An attempt
velocity
correction,
it was
necessary to perform a positive yaw maneuver of II0. 5 degrees.
of 75. 3 degrees and a positive was made to reconstruct the
total premidcourse maneuver phase from the beginning of the first roll maneuver to vernier ignition and to compute the roll axis pointing error exclusive of any tracking data. Several variables affect the accuracy of an angular maneuver, includ-
ing precession rate accuracy, precession comn_and time, gyro drift, and initial attitude errors due to biases and limit cycle. When several maneuvers are performed with large time intervals between them, attitude errors due to gyro drift must be included. A list of all parameters affecting midcourse maneuver accuracy is presented in Table 5. 5-16 along with the allowable 3u values and actual performance values whenever possible.
5. 5-58
Figure
5. 5-14.
Roll
Gyro
Response
Time
During
Premidcourse
Roll
5.5-59
TABLE
5. 5-16.
PREMIDCOURSE
ATTITUDE
ERROR
SUMMARY
3_r Parameter Requirement References
Measured Value Pitch Yaw = 40.087 degree = -0.04 degree Comments
Primary sun sensor null with respect to FCSG roll axis
0. Z degree
5 , paragraph 4.3.1. 1
Canopus with FCSG plane Pitch/yaw cycle
sensor
null
0. Z degree
respect to roll/pitch
5, paragraph 4.3.1.2
+0.055
degree
limit
0.3
degree
5, paragraph 4.3.1.1
+0.05 +0.066
degree/ degree
Based error start of
on
sun yaw
sensor at
signals
Roll
limit
cycle
0.3
degree
5, paragraph 4.3.1.2
+0.073
degree
Based error start
on Canopus signal at of roll
Gyro scale
torquer factor
0.
15percent
11, paragraph 3. Z. 5.1.3 Spacecraft sion rate O. Z percent from star O. 4498 ± deg/sec precesdetermined map O. 0008 was
Precession current accuracy Precession current drift source source
0.
13percent
0. i percent
Timing accuracy
source
0. Z second 0. 0Zpercent
±
Roll Yaw
- 40.028 degree ....0.01 degree
Based on timing errors determined in subsection 5.5.4.6
Gyro alignment to FCSG roll axis
0.14
degree
11, paragrapn 3.2.5.1.4
Pitch Yaw
= +0.037 degree = *0. I00 degree
FCSG/spacecraft roll alignment
0. i degree axis
5, paragraph 4.1.3.7.1
+0. 023 YaW itch =
degree +81 | ]
seconds = - 3 seconds! = -0. degree = +0.2 degree =+0.05 degree 2
Gyro drift
non-g
sensitive
1.0
deg/hr
5 , paragraph 4.3.1.5
Roll Yaw Pitch
Based
on
measured in roll and
-0. 78 deg/hr for 16 minutes 1 second;
+l. 0 deg/hr in yaw for l l minutes and 56 seconds; deg/hr in pitc}
+0.25 Total error ignition attitude prior to +0. along 39 degree neBative with
yaw axis O. Z degree uncertainty
5.5-60
Determination
of
Precession
Times
With onds
The register was loaded with 377 bits for roll and 553 bits for yaw. a clock rate of Z. 5 cps, the respective times are 150. 8 and 221. Z secwith a maximum error of 0.20 second. used to determine the the maneuvers was The results are as
The gyro error signal telemetry data were actual precession time. The sampling rate during 20 times/sec, giving a resolution of 0. 05 second. follows (Figures 5. 5-15 and 5. 5-16): T T = 150. = 221. 744 182 seconds seconds (roll) (yaw) Error
Attitude
Maneuver
Reference IZ develops two orthogonal equations that specify spacecraft thrust axis pointing error during midcourse thrusting. equations were derived for a roll-pitch rotation sequence. Rewriting neglecting error in the following these sources equations: equations that are for a present roll-yaw only rotation sequence after engine ignition
the The
and results
Error
along
pitch
axis
= sin %b (_SAE
+4#RE
) + 8A E
cos
_
cos
_5
- %bAE
cosd2
sin
O 0
Error
along
yaw
axis
= - _bRE
- cos_
_A E
- sinqb
@AE
where
(qb, O, are
_)AE rotation
are
spacecraft errors.
inertial
reference
alignment
errors
and
(d_, qb )R E
Use of _ = 75. 3 degrees, _ = ll0. 5 degrees, and the errors listed in the summary chart results in an attitude error of 0. 39 degree along the negative yaw axis and an error of 0. 04 degree along the negative pitch axis. The resultant pointing error has a 99 percent circular probable uncertainty of 0. Z0 degree,
5. 5-61
°_.-g
ol.-I
(D _>
O
0
,_..4
_4
I
Lt_
_Z
5. 5-6g
68189-2--269
or-4
E
°r-_
_D
>_
0
-,-I
_D
I
_4
©
o_
°,-_
5.5-63
Precession
Rates
Accuracy of the precession rates imposed by the "Surveyor System Functional Requirements Specification' (Reference 5) is 0. 5000 + 0. 0011 deg sec. The precession rate obtained during the star mapping phase indicates that the positive precession rate is 0. 4998 deg/sec with a data granularity of 0. 0008 deg/sec. 5. 5. 4. 7 Nitrogen Gas Consumed in commutator during the
mode
Since the nitrogen tank temperature is not available I, an accurate estimate of the nitrogen gas consumed
premidcourse attitude maneuvers could not be made. If it is assumed that the nitrogen tank temperature did not change appreciably during the maneuvers, the estimated gas usage was 0. 06 pound. This compares favorably with an expected value of 0. 055 pound (Reference 4).
5. 5-64
5. 5. 4.8
Midcourse The desired
Velocity midcourse
Correction burn duration of 9. 850 seconds was entered
into the spacecraft magnitude register 5 minutes before the planned ignition time of 05:00:00. Bulk printer data indicated ignition (magnitude register started to count down) at 05:00:02. 5. Within a few seconds after ignition, flight control telemetry signals indicated hard-over pitch, yaw, and roll gyro errors, roll actuator position, and acceleration error. Vernier engine strain gage telemetry signals indicated thrust on legs 1 and 2, but zero on leg 3. The leg 3 throttling signal telemetry went hard-over to the maximum thrust command position and remained there throughout the burn. Loss of thrust a lateral axis. on leg Effects 3 caused the spacecraft due to the roll actuator to spin-up initially and nonsymmetrical a rate
about
inertia properties of the spacecraft caused the ensuing motion to become tumbling about all axes. During the 9. 85-second burn time, the tumble built up to 1. 25 rps, as indicated by fluctuation of the receiver automatic gain control or secondary sun sensor telemetry signals. Following vernier engine cutoff, the tumble 32 percent by action of the gas jet attitude control were inhibited by ground command 14 minutes and after 50 percent of the premidcourse fuel load had tumble rate by then had been reduced to 0. 85 rps. Flight control system performance during
rate was decreased system. The gas jets Z0 seconds after cutoff been expended. The
the midcourse
burn
appeared to be normal under the prevailing circumstances. The flight control system outputs (vernier engine throttling signals and roll actuator position) behaved in a predictable manner during the vernier ignition transient and throughout the burn period. Engines 1 and 2 had an ignition delay time of less than 120 milliseconds (see subsection 5. 5. 4. 8) and responded well to their respective throttling signals. The calculation of vernier dispersions, as done in the case data, was not possible for SC-2. masked by the engine 3 failure. of Flight Control System Behavior engine startup and shutdown impulse of SC-I from pitch and yaw gyro telemetry The relatively small effects produced were
Description
period Figure errors,
Behavior of the flight control system during the 9. 85-second burn is depictedby the real-time SPAC brush recordings shown in 5. 5-17. The inertial sensor signals--pitch, yaw, and roll gyro and acceleration error (the processed accelerometer signal)-are and FC-15, throttle-valve analog channels
given by telemetry analog channels FC-16, FC-17, FC-49, respectively. The flight control system outputs - vernier signals, and roll actuator positionare given by telemetry
5.5-65
Ill,lllillltll
.............. llHLP,441B4_fH I'i'_+,i I + IIi
I
IIIll
'
,,,,,,,,,,,+ ............ ............. ............
i i III] [ l],ill ] I II I I I I E IJ_-J I i [ ] ]
II
I i I i r i
+4........
+"1."1 IIIII
II i i i
[ I I1 i:' IILIII
i 1 i
_lllt
I I i i
!: I
1
,
l I I I I I I
IfHl,,,,,,,,, ............. ,,+.HIHH+ 4-: t t H }4 {Hq44-H.i .... i, t t td f H4+H444H-H HH-h'H H4+_
_m_:;+,
Figure 5. 5-17. Midcourse 5.5-66 Thrust
H!
FC-Z5, FC-Z6, FC-Z7, and FC-43, Figure 5. 5-17 are the three vernier from the four secondary sun sensor from 30 seconds before ignition until
respectively. Also shown in strain gage signals and the indications cells. The time period shown extends 1 minute and 30 seconds past cutoff.
pitch
Pitch/Yaw Behavior. and yaw motion diverged
With the loss of engine 3 thrust, beyond the gyro telemetry range
spacecraft (±7. 5degrees)
within Z seconds after ignition. Although no direct indication of angular rate is available, the rate buildup during the burn period is apparent from the increasing fluctuation rate of the secondary sun sensor signals. Final rotation rate at the end of burn was 448 deg/sec (l 24 rps) as indicated by receiver automatic gain control fluctuations. Following engine cutoff, pitch and yaw gyro errors moved from stop to stop, indicating polarity changes in pitch and yaw body rates due to the tumbling motion. This general behavior exhibited in flight has been duplicated by analog computer simulation (see subsection 5. 5. 4. I0). In the transient motion which yaw motions were obtained. 1 and Z relative to pitch and occurred at ignition, negative pitch and This follows from the locations of yaw body axes as shown in Figure 5. 5-18. 3,
positive engines
Also, yaw divergence to the fact that engine which is evident from
was slower than pitch divergence. This is attributed 2 was commanded to a lower thrust level than engine a comparison of FC-Z5 and FC-26 in Figure 5. 5-17.
The engine geometry also contributes to this, since equal thrust on engines l and Z would cause pitch motion to diverge somewhat faster (17 percent) than in yaw. This initial motion has been duplicated in a mixed flight control electronics analog computer simulation (see subsection 5. 5. 4. 10), where it was found that the small residual attitude errors (less than ±0. 4 deg/axis) due to gas jet deadband and electronic nulls have a strong influence on the initial thrust transients. The tumbling motion that continued after cutoff is difficult to describe since it involved fluctuating angular rates about all three body axes. It is further complicated by the fact that the spacecraft is nearly an inertial sphere, i. e. , the principal moments of inertia are nearly equal, so that an analytical description is extremely complex. The motion has been closely duplicated by analog computer simulation (subsection 5. 5.4. i0), however, so only a few remarks on the general nature of the motion will be made here. From the pitch, yaw, Figure 5. 5-17, it is apparent polarity periodically (period and roll gyro error traces shown in that pitch and yaw angular rates reversed = 13 seconds) and that roll rate was of constant
polarity. This is consistent with the general form of the theoretical solution for the unforced tumbling motion (Reference 13). In the general case (for zero cross-products of inertia), the angular rates about the axes of least inertia will fluctuate periodically about zero, whereas the rate about the axis bias of maximum level. inertia will fluctuate at twice the frequency about some
5. 5-67
\
\
/
\
/
J
Figure
5. 5-18.
Vernier
Thrust
Chamber
1,ocations
5.5-68
The Pitch Yaw Roll Cross
SC-Z
inertia
properties I xx I YY I zz = = = = =
were 208.
as
follows
(Reference
14):
3 (slug-ft 2)
204. 7 215. 4 -7. 3 7.8
products
I xy I xz I yz
=
-1.8
Thus, inertia, reversed
since
the
roll
axis
was did
aligned not change
roughly sign,
with
the
axis pitch
of
maximum and yaw rates
roll angular rate sign periodically. Vernier Behavior Throttling of the
whereas
Signals vernier engine throttling (Figure generated control signals (FC-25, FC-26, is and a
FC-27) linear attitude following
is shown in the brush recordings combination of command signals control loops and the acceleration relationships (Reference 11): A AT 1 A &T 2 A AT 3 0. 171 0. 041
5. 5-17). Each signal by the pitch and yaw loop according to the
A
-0. 147
_o.o lix /4
0. 222 0. 333 Y 0. 333/ I_ATz A commands, and AT z is the net A AT 3 (FC-27) became saturated ignition, since all three control in
A where thrust
A are pitch command. and yaw moment
L x and Ly increment
the
Thus, as seen positive direction
in Figure immediately
5. 5-17, after
c_annels v)_ere commanding more thrust from engine 3. On the other hand, AT 1 and AT 2 (FC-25 and FC-26) were each commanded up by the acceleration loop and down by the attitude loops, and approached compromise levels inside the saturation limits. Assuming that all control channels are saturated in the polarities throttle-valve signals can the predicted values agree indicated by the SG-Z telemetry, the resulting be calculated as shown in Table 5.5-17 where well with those observed in flight.
5.5-69
TABLE
5.
5-17.
COMPARISON OF THROTTLE-VALVE
CALCULATED SIGNALS
AND
OBSERVED
Throttle
Valve n_illianlper Leg 150 36 -192 Z
Signals, e s Leg 150 148 127 3
Command
Source
Leg 150 -175 64
1
Acceleration Pitch Yaw loop loop
loop
Net,
calculated
39
-6
>8 0 ":_
Net,
observed
4O
-5
80 '_'_ ':"
':"Maximum *-'::Telemetry
capability saturation
of vernier level.
valve
amplifier
s.
As noted from Figure 5.5-17, the acceleration became saturated shortly after ignition and remained This follows from the fact that the commanded midcourse level (0. 1 sion system, engines 1 maximum even with g), requiring was not achieved. and 2 were capable total of Z14 pounds. both engines operating earth 220
error signal (FC-15) saturated thereafter. acceleration
pounds of thrust from the vernier propulA review of vernier engine data shows that of 106 and 108 pounds, respectively, or a Thus, the 0. 1 ge level was unattainable at their maximum levels.
5. 5-70
m
Roll The
Actuator
Response responded normally, under the circumstances,
roll actuator
throughout the midcourse burn period. At ignition, a positive roll error signal caused the actuator to deflect a maximum of 2.7 degrees in the negative direction. Then, due to coupling of the uncontrolled tumbling motion into roll, the roll gyro error reversed, causing the actuator to move to its positive 6-degree travel limit where it remained thereafter. Calculations that follow show that the actual actuator response agrees well with predictions based on the roll control system transfer function and the observed roll gyro error. Initial Transient. During the failure analysis, a question concerning an apparent inconsistency in the initial actuator response was raised at Hughes and JPL. The following discussion and analysis are included here to cover this point. A review obs e rvations : of telemetry data at midcourse ignition yields the following
i)
Roll gyro error {FC-49, generally referred to as roll precession command) shows no apparent change over a ?50-millisecond interval beginning at ignition. Roll actuator ignition. deflection (FC-43) shows an immediate response at
z)
Since the roll actuator is driven by a signal derived from the roll gyro error signal, the above observations would lead one to believe that the actuator was responding improperly to an error signal null.
The following analysis shows that the roll actuator response was in fact normal, and that this apparent inconsistency was caused by a combination of the following factors: l) 2) Granularity Filter the of roll gyro telemetry signal telemetry (0. 033 deg/BCD) (0.28 second) shown to
time-constant raw telemetry The plot extends signal behavior.
of roll gyro data for Also
circuit
in Figure indicate
A plot of 5.5-19. general
for 6 (FC-43) and Oz (FC-49) is several seconds beyond ignition shown is the result of a calculation
to check actuator response channel transfer function. etry value of 6 is -1.88 ond difference in FC-43 prediction beyond 6 = -1. that the
relative to that predicted from For the calculation checkpoint degrees. To approximately account and FC-49 telemetry circuit time O z occurring 5. 5-19, the value. It
the nominal roll chosen, the telemfor the 0. 25 secconstants, the 0. 25 prediction is thus second yields concluded
of 6 is based on the values of Oz and the checkpoint. As indicated in Figure 79 degrees, nearly equal to the observed observed actuator response was normal.
5. 5-71
Figure 5. 5-20 shows a plot of 6 and @zobtained from a closed-loop analog computer simulation (including actual flight control electronics hardware) of the midcourse startup. The scales in Figure 5. 5-20 are identical to Figure 5. 5-19. The @ztrace (which includes the effect of a 0. Z8-second telemetry filter time-constant) is observed to change very little over the initial Z50-millisecond interval. In fact, approximately 200 milliseconds are required to develop an amplitude equivalent to 1.0 BCD and, with the spacecraft telemetry system, no change would have been registered over this interval. Thus, the analog computer result, which yielded a 1.7-degree peak of 0z, indicates no 0z BCD change over a 200-millisecond interval. Since the peak 0z in the SC-Z case was I. 35 degrees, no BCD change would be expected for an interval of 1.7 x 200 milliseconds 1.35 -- Z50 milliseconds
which agrees with the observed result in SC-2.
Conclusions. l) Roll The following response conclusions during have been phase reached: appeared normal.
actuator
midcourse
z)
Zero BCD change in roll gyro error over interval is consistent with analog computer Jet System Operation
the first Z50-millisecond study results.
Gas
and
29
The cold gas attitude control system remained active until 14 minutes seconds after midcourse ignition, when the gas jet amplifiers were
inhibited by ground command. As a result, 2. 19 pounds of nitrogen(50.5 percent of premidcourse load) was expended in reducing the spacecraft tumble rate from 448 deg/sec to 306 deg/sec (31. 7 percent reduction). During this time, two gas jets were on continuously, and one jet was on about 60 percent of the time. The net specific impulse of the system was about 64 seconds over this time period, yielding a rate reduction efficiency of approximately 1.0 deg/sec per Ib-sec of total jet impulse. Gas was Jet Duty operating Cycle. with The following calculations show an average of 2. 6 jets thrusting. that the gas jet Reference 15
system
shows that a large pure pitch or pure roll angular cable gas jets on continuously. Because of phase a large yaw error signal would hold the No. Z and of the time.
rate would hold the applidetector voltage saturation, 3 gas jets on for 80 percent
5. 5-72
13 _5
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5. 5-74
cycle ship
In Section 6.4.2 and steady-state for current system
of Reference 16, a relationship between gas jet amplifier input is derived. The parameter values can be expressed as
gas same
jet duty relation-
ton Duty cycle = ratio to total 3. 2h (9 -v) 25.6h _ (v-l) 8.0 of on-time period
(9-v)(v-i)
where input
V "_
voltage voltage
deadspace
Thus, v = 7.4 for an 80 percent duty cycle and, using the measured SC-2 yaw deadspace of 0.22 degree, the yaw phase detector output voltage would be equivalent to (7.4)(0. ZZ) = i. 63-degree error. Assuming that all phase detectors saturate at the same voltage level, roll and pitch levels of saturation are 2. 7 and 2.4 degrees, respectively. The observed thrusting were 1) 2) 3) Saturated three pairs Roll: Pitch: Yaw: signals of gas polarities as follows: of the saturated gyro signals following mid-
course
negative negative positive from jet these amplifiers gyros would have the (after going through be commanded jet would be following effect the summing signal upon the matrix): equivalent
1) z)
No. 1 CCW jet to 2.7 degrees.
would This
on
on by a roll continuously. on
No. 3 CCW jet would be commanded signals and off by the roll signal. would be equivalent to 3.5 degrees. c ontinuou sly.
Net
by the pitch command to This jet would
and this also
yaw amplifier be on
5. 5-75
31
No. and
2 CCW jet would be commanded commanded off by yaw signal.
on Net
by pitch command
and roll signals to amplifier times nominal
would be equivalent to 1. 29 degrees, which is 5. 84 deadspace. Duty cycle for this jet would then be
ton T or, this jet would the gas jet system be on about 60 was operating
15. 84-i) 8.0
0.60
percent of the time. Thus, with 2.6 jets thrusting.
on
the average,
used
Fuel Consumption. in the fuel consumption Nitrogen Nitrogen Nitrogen Isp 869sec_ weight weight used
Spacecraft consulting analysis calculations as follows: before after in rate midcourse midcourse reduction 17 notes ignition. z 4. 34 pounds _ Z. 15 pounds = Z. 19 pounds that the gas
team
data
were
for
Determination. _
Reference midcourse
jets were
enabled
Isp is calculated
from
the following
equation-
(time)(total Isp weight (869
thrust) used pound
of nitrogen seconds)
(2. 6 x 0. 0622)
2. 19 pounds -- 64. 3 seconds where the thrust per jet is taken to be the thrust value under ditions as recorded in the SC-2 Surveyor Flight Control Data summary The fuel budget data. value for Isp is between the calculations and 73 seconds, lower bound as measured (60 seconds) in Reference used for 18. An full flow Package con-
Isp less than 73 seconds was expected because Isp decreases with decreasing temperature, and it is thought that the gas jets were materially cooled by convection as the nitrogen flowed for this long thrusting period. Rate Reduction Efficiency. mined from SPAC automatic gain The initial tumbling rate was 448 been reduced reflects how impulse A plot of spacecraft tumble rate (detercontrol data) is shown in Figure 5. 5-21. deg/sec and, after 14. 5 minutes, it has
to 306 deg/sec. In the following, I is an efficiency figure that many deg/sec of body rate is eliminated for each lb-sec of from the cold gas jet control system.
expended
5. 5-76
460
440
420
o z o w a400
(1) .i w n, 380 o O_
G
z "7 360
340
320
300 0 200 TIME FROM END 400 OF MIOCOURSE 600 BURN, SECONDS 800 IO00
Figure
5. 5-21.
Spacecraft
Tumble
Rate
After
Midcourse
5. 5-77
aw-(X)(a )or Z
A¢o=
sp
1 _ x (aw)(I
142 deg/sec seconds
sp
I = 64.3 sp &W
= 2. 19 pounds
x
(14Z) deg/sec (2. 19) (64.3) = i. 01 Ib-sec
This rate in previous analog efficiency figures in different ways
reduction efficiency figure is consistent with values obtained computer studies. Reference 1 records the following (Table 5. 5-18) for a 5 deg/sec initial body rate distributed among the pitch, yaw, and roll axes.
TABLE
5. 5-18.
RATE
REDUCTION
EFFICIENCY
Initial Rate Magnitude, deg/sec Yaw Pitch Roll 5 0 0 3. 535 0 3. 535 2. 882 0 5 0 3. 535 3. 535 0 2. 88Z 0 0 5 0 3. 535 3. 535 2. 882
Impulse Expended, Ib-sec
Efficiency Figure, deg/sec ib-sec 1.33 1.49 0.96 1.40 1.35 0.91 I. 07
3.75 3.35 5.20 3.58 3.70 5.50 4.68
5. 5-78
5.
5. 4.9
Postmidcourse
Vernier
Engine
Firings
In order to bring about leg 3 vernier engine ignition, 39 additional vernier engine firings were programmed and executed between the time of the midcourse firing command (0721 at 264:05:00:02) and the retro firing command (0730 at 265:09:34:17) (see Section 4. i). Thirty of these firings were for commanded durations of_0. Z second, seven for _Z. 0 seconds, one for _2.5 seconds and one for _21.5 seconds. (The midcourse commanded duration was 9.85 seconds. ) 0 seconds) in terms of error derived from The strain gage and in Figures 5. 5-23
thrust SPAC thrust through
A summary of the longer firings (burns) ( ->2. commands, strain gage response, and acceleration brush recordings is presented in Figure 5. 5-22. command data are plotted with greater resolution 5. 5-32,
preference rate and subsequent
Postmidcourse burn analysis concentrated on the longer to the 0. Z-second burns due to the disparity between vernier on time. * Analysis of the latter burns is to be to submission of this report. The general aspect of the 2. 0-second firings was as
burns in sample performed
follows:
1)
At the fire command, legs 1 and 2 thrust commands immediately dropped from midthrust to minimum thrust, and the leg 3 thrust command immediately increased from midthrust to maximum thrust. The strain gage readings for legs 1 and Z increased from zero to the commanded thrust (approximately, making allowance for transducer drift due to spacecraft perature variations). The leg 3 strain according to the thrust command, but acceleration and temperature. accelerations and temgage did not respond did show the effects of
2)
Acceleration each firing Roll Gas actuator jets were
error period. (FC-43) inhibited.
(FC-15)
signal
was
saturated
(+)
during
3) 4)
remained
hard
over
at
+6
degrees.
5)
Pitch (F-16), yaw remained saturated Telemetry condition:
(FC-17), (minus, Mode
and plus, 1,
roll and 1100
(FC-49) minus,
error signals respectively).
6)
bits/sec
".:One sample 35 through correspond
every 39, for to mode
1. 6 seconds for the which 5 samples/sec 1, 137 bits/sec and
strain were mode
gages read. 1,
on 1100
all burns except (These sample rates bits/sec. )
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Engine of Thrust Burn
I Command 4{) and Strain Gaga Data,
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2 of Thrust 40 Command and Strain Gage
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For
the
2. 5-second
and
31. 5-second
firings,
the vernier
engines
were ignited at about 90-pound thrust command level (retro mode). Legs 1 and 2 immediately throttled down an increment close to that experienced for the 2. 0-second firings, after which throttling control was exercised. The leg 3 command stepped to maximum at ignition and remained there until engine cutoff. Engine throttling during the first 2. 5 seconds of the 21. 5-second firing was quite similar to that during the 2. 5-second firing. The remainder of the flight control signals were the same as for the 2. 0-second Burn An firings. Duration accurate indication of commanded burn duration was obtained from
the magnitude register (FC-18) (Table 5.5-19) The register countdown was assumed to be linear with time, allowing extension of the straight line established from two or more data points during the countdown, and thus the determination of the times of zero and full count. The accuracy of burn duration measurements made in this manner is limited by the on-board clock and data processing technique. Of course, the absolute time is in question by ± 50 milliseconds (i.e. , ± one word time) due to the granularity of the telemetry system. Isnition Time
The data needed to bracket vernier engine ignition times is listed in Table 5. 5-20, based on the apparent telemetry response of EP-4 (22-volt bus current), FC-25, FC-26, FC-27 (thrust commands: legs l, 2, and 3), FC-18 (magnitude register), and P-18, P-19 (strain gages: legs l and 2). The interval between "before" and "after" times is due not only to engine ignition delay uncertainty, but also telemetry granularity. The size of the interval depends on telemetry mode, bit rate, and location of the ignition event relative to the data words in the frame (Table 5. 5-19). The signals listed above were used to determine the last known time prior to ignition. The ignition response was determined from strain gages alone. Figure 5. 5-33 graphically shows the method for burn 27. Observations I) relative to the ignition times are as follows: 1 and 2 TCAs
Flight acceptance were 0. 088 and Ignition times
test ignition 0. 081 second, in Table
times for the leg respectively. 5. 5-19 are maximums.
2)
listed
3)
Effect of data mode and bit rate be seen from the time increment in Table 5. 5-19.
on measured ignition between strain gage
times can samples
5.5-I15
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a)
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1
b)
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preceding
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a) b)
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command change
is within 3 milliseconds '_ FC-25 reading. 2 was probably normal
after
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(i.e.,
<0.
100-
2)
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a
normal
ignition. 2, 3, 15, and ignition times. 21 probably
3) 4)
Ignition times for burns 1 (midcourse), did not experience exceptionally long Based on data analyzed, it cannot anomalous vernier engine ignition occur on any of the vernier engine Nutation
be stated with certainty that performance did or did not burns.
Spacecraft
Apart from other detectable characteristics (e. g., thermal drift), the strain gage data (Figures 5. 5-23 through 5. 5-32) show the effect of cyclic acceleration following each vernier engine burn. Amplitude and period of oscillation can be used as an indicator of the consistency of the applied torque are presented Examination damped nature of the impulse in Table of for the 2-second 5. 5-19. data just prior to burns. Amplitude and period data
and
after
each
burn
shows
the
oscillation. Firing to all data Anomaly (Burn 27)
Postmidcourse It vernier on burn is possible engine during 27 from the
establish vernier discussed
the anomalous performance of the leg 3 engine burns and the leg 1 vernier engine above. The leg 3 anomaly is considered
5.5-119
a continuation of that occurring at the midcourse correction and causing loss of spacecraft control. However, the anomalous performance of leg l during burn 27 does not appear to be (directly) connected with the midcourse failure.
The burn shutoff after the from an attitude loss of spacecraft follows : 27 extra-performance, commanded firing control standpoint could result. consisting of vernier termination, is of particular during the terminal descent Burn 27 anomaly supporting engine interest phase, since data is as
1)
Strain
gage
data
show
(Figures i thrust
5.
5-22
through termination
5. 5-32): of commanded
a) b) cl
Continuation interval. Leg 2 cutoff
of leg
after
per
command
(strain
gage
data). after termination throttling command ) nutation compared to the of
Leg 1 response to throttling commands commanded interval. (Also, absence response for leg 2 in the same interval. A definite change by in character of the
of
d)
postburn 27,
experienced other 2-second
the spacecraft burns (Table show the following: (i.e., that
for burn 5. 5-19).
2)
Thrust
commands Programmed Command response. register level
a) b)
3) 4)
duration changes
-_2
seconds). to strain gage
correspond
Magnitude
(FC-18) data
shows
proper
commanded
duration.
Combined telemetry with strain gage-indicated duration measured. proper time. )
show engine cutoff on leg 2 compatible time of ignition and commanded (Thus, cutoff command was received at
the
5.5-120
J
5. 5.4.10
Retro
Firing were ended transmitted on 22 the SC-2 mission: GMT, power on September as part of
the
The following commands retro firing sequence which Command Thrust Enable Manual Reset Load with Mode Reset Retro phase gas
hr:min: 09:19:06 09:30:33 09:30:53
sec
jets
lockon nominal magnitude 8. 0-second 2 Group sequence AMR IV mode mark on thrust register delay bias
09:31:12 09:31:51
09:32:19 09:32:55 09:33:14 09:34:17
Emergency
The emergency AMR mark command was apparently received at 09:34:19. 178 as indicated by start of the magnitude register countdown (Figure 5. 5-34a). The clock counted down smoothly for the desired 8 seconds, at which time ignition of vernier engines 1 and 2 occurred as indicated by the telemetered strain gage signals (Figure 5. 5-34b and c). The earliest indication of retro ignition by means of the retro ignition latch going high, was at 09:34:28. 578. At about the time of retro ignition, vernier engines 1 and 2 were shut off, as indicated by the strain gages, and remained off until all data were lost at 09:35:00. The vernier ignition latch (FC-28), vernier engine command signals (FC-25, FC-26, and FC-27)(Figures 5.5-34e through 5. 5-34g), and the magnitude register signals (FC-18) remained normal during this time. At increased from approximately retro ignition, acceleration 7. 1 to I0.3 g and remained time it gradually increased along the Z axis at this level for to ii. 5 g when
around 18 seconds, after which data were lost (Figure 5.5-34h). Estimated based on nitrogen a pressure gas of
pounds
remaining 1340 psi
at
when 43°F.
data
were
lost
was
1.62
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5.5-129
5. 5.4. ii
Postmission
Tests
and
Analyses Flight Control and other perti37 comprise
Several special tests were performed in the Hughes Laboratory to assist in postmission analyses. These tests nent analyses are discussed below. (References 20 through control IDCs documenting the postmission activity. ) SC-2 Gyro Error and Thrust Command Telen_etry
Accuracy
It was determined that the gyro error telen_etry signals were not appreciably affected by a degradation in gyro transfer function and that their accuracy is better than that associated with the thrust command telemetry signals. A comparison of tolerance allotments for the gyro error and thrust commands telemetry circuit components is shown in Table 5. 5-21. Although Reference 12 discusses these tolerances in detail, the following clarification of the values listed should be noted:
l) z)
Specification allowance cent, -20 percent.
on
coil resistance
is 400
ohms,
+33
per-
The 20 percent value listed was taken as a convenient "symmetrical" number approximating the maxi1_ur_7. It is a useful crude limit if baseline coil resistance and ten_perature are not available. The 2 percent value listed is probably the best attainable if a reliable baseline resistance measurenzent is available and corrections are made for coil temperature. Acceptance test data (taken at Reaction Motors Division of Thiokol prior to delivery) on resistance of the solenoid valve coils are presented in Table 5. 5-22. All values were within 3 percent of nominal. IXote that coil temperature changes following application power caused an appreciable resistance change 4. 5 minutes prior to the 21. 5-second burn. A the telemetered _null" output was observed. of thrust-phase on leg 1 in the 9 percent drop in
3)
4)
Data presented in Table 5.5-21 do not include tolerance on performance itself (accuracy of thrust developed at a given of coil differential current); this is not insignificant.
engine level
Since a variation in gyro transfer function (no_1_inally 44 mv/deg) directly affects the gyro error telemetry signal scale factor, the transfer function histories of the three gyros were investigated for their possible effect on postmission analysis. These data are presented as Figures 5. 5-35, 5. 5-36 and 5. 5-37. The results are summarized as follows: Yaw Gyro: with some tests. On mv/deg for at midcourse, A gradual increase of transfer function with time is noted evidence of leveling off in the last i00 hours of inertial lab this basis, it is believed that the best estimate of 43. 2 yaw gyro is the transfer function, as same as that measured operating in the at AFETR. spacecraft
5. 5-130
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M
5. 5-131
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5. 5-132
5.5-133
TABLE COMMAND
5.
5-21. AND
RELATIVE GYRO ERROR
ACCURACY TELEMETRY,
OF
THRUST PERCENT
Telemetry Circuit Error Pitch Roll Thrust and gyro command yaw gyros <2 1 2
Errors Electronic Components in Serie 3 and 3 and 3, 5, and
of s s 5 5 5 2O 2 Valve Coil Error Total Error, RSS 6.2 6.1 22. 8.2
TABLE
5. 5-22. OF
TORQUE THROTTLE
MOTOR COIL VALVES
RESISTANCE
Leg Valve serial number ohms to pin B* to pin D 396. 398. 5356
1
Leg 5376
2
Leg 5380
3
Resistance, Pin Pin A C
3 0
395. 406.
5 l
High
current
in
this
coil
closes
valve.
5.5-134
Pitch Gyro: A slowly midcourse is estimated AFF, TR measurement.
decreasing at 42.?
trend mv/deg,
is apparent; the about 1 percent
value at below the
Roll Gyro: A moderately decreasing course is estimated at 39. 2 mv/deg, measurement. Inasmuch data, and normally midcourse telemetry as the telemetry calibrations
trend about
is apparent; value at mid2 percent below the AFETR
were
based
on
the
AFETR
no data exist to indicate that the gyros were operating other than during the flight, it is concluded that the transfer functions at were not significantly calibrations. different from those that established the
While they are a very insensitive indicator, the dynamics of the premidcourse yaw maneuver were examined for any evidence of low controlloop gain that might have been attributable to low gyro transfer function. No such evidence was apparent. Gyro Telemetry Saturation. Limiting characteristics of a demodulator prior to the telemetry pickoff point, and the telemetry output limits of 0 and 5 volts, result in a gyro telemetry saturation characteristic typified by the solid line in Figure 5. 5-38. AFETR data for the pitch and yaw gyros, overplotted on this figure, fit the nominal reference very well, with just a hint of the typical break in slope at 6 degrees. To obtain a better idea of the saturating behavior Control Laboratory of actual hardware, data were obtained in the on a prototype inertial reference unit (Figure Flight 5. 5-39).
The yaw gyro in that unit exhibited characteristics those of the SC-2 pitch and yaw gyros and validated curve angles for use in correction above 6 degrees. 2?.-Volt Thrust Phase SC-2 gyro data telemetry
closely comparable to the nominal saturation indications at
Bus
Current
During
Midcourse that existed 2Z-volt
during
An attempt was made to duplicate the midcourse velocity correction.
the current waveforms Measurement of the
thrust phase bus the SC-I "ZAP" 13) on the FCSG
power and power control waveforms was performed using flight control electronics unit (FCEU) (P/N 273100-6, S/IN flight acceptance test console with its associated roll
actuator simulator (T284828). Three vernier engine prop valve solenoids (S/N 230, 247, and 236) were obtained for this test. The voltage and current waveforms that appear on the Z2-volt power and the FCEU power control circuitry under various operating conditions of the roll actuator are shown in Figures 5. 5-40, 5. 5-41, and 5. 5-42. Figure 5. 5-43 depicts the applicable control circuitry in the FCEU. In addition, an attempt was made to obtain the current waveforms by simulating the transient roll conditions that existed during midcourse. The spacecraft roll rates and roll acceleration were simulated with the inertial reference unit mounted on a Genisco rate table. Figure 5. 5-44 shows the 22-volt thrust phase current waveforms
5. 5-135
68189-2-316
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5.5-136
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5. 5-137
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Figure
5. 5-40.
Current
Waveforms
5.5-138
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A
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0
Figure
5. 5-40
(continued).
Current
Waveforms
5.5-139
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g
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Figure
5. 5-41.
Voltage
and
Current
Waveforms
5.5-140
,,D i !
Ii
cO / COA,.r7 ,_HA3_ ON O_ j y Figure 5. 5-42. Voltage
,3
2 / r_
.0]4-
"t;
Waveforms
--V
czo
and
VRI6
5.5-141
68189-2-322
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0
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5.5-142
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AvI.5_"C
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A-I 3
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Figure
5. 5-44.
22 volt
Thrust
Phase
Current
Waveforms
5.5-143
that were obtained by monitoring on an oscilloscope the voltage across a l-ohm resistor inserted in the ZZ-volt thrust phase return line at IA7-H.
Figure 5. 5-45 etry (roll gyro nals. Numbers is the brush recording of the roll precession error) and the roll actuator position feedback for the roll rates and roll acceleration were command telemetry reduced telemsigfrom
the brush recording. These data were used to isolate the unregulated 2Z-volt bus current due to the roll actuator from that of the vernier engine propellant valves, gas jets, and gyro heaters. Dynamic Telemetry Versus Static Signal Calibration of Canopus Sensor Mapping
sensor
In order to determine star mapping channel
more effectively the calibration of the Canopus in space, the sensor mapping signal was mea-
sured for both static and dynamic conditions. Comparisons of the mapping circuit telemetry output under both conditions for 0.67X Canopus (lower lockon threshold), i. 0X Canopus, and 1.5X Canopus (upper lockon threshold) are shown in Figure 5.5-46. The static calibration corresponds to what is observed in a mission during the gyro drift check while the dynamic calibration, which was done at an equivalent spacecraft roll rate of 0. 5 deg/sec, corresponds to what is observed during the normal star mapping phase of the mission. These data were used in conjunction with SC-I and SC-Z star mapping data to more precisely establish the calibration of the sensor in space, as discussed in subsection 5. 5.4.4. Computer Simulations
Analog and digital computer programs have been used to simulate the midcourse firing._:= Solne of the simulation was done with the SC-I ZAP electronics FCEU as part of the closed loop. The spacecraft electronics contain many large signal nonlinear effects that become important for operation when the gyros and accelerometer are hard over. The best match with SC-2 obtained using a mixed simulation puter. The best data match over all-analog computer simulation. telemetry data over the first Z seconds was of SC-I ZAP electronics and analog coma 25-second period was obtained with the The better long-term data match was for the differences in the test assumed small engine startup no thrust from engine 3.
obtained using all analog by compensating electronics and that of SC-2. These data delays and initial gyro angles, as well as
The first simulation attempts were closed loop (analog simulation of electronics and equations of motion). When the result showed discrepancies with the telemetry data, representation of the nonlinearities in the electronics was suspect. While these were being measured, the simulation continued open loop with programmed thrusts acting into the equations of motion.
_",-'See Reference
19.
5.5-144
)
Figure
5. 5-45.
Roll
Actuator
Test
Results
5. 5-145_
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5.5_149
Figures 5.5- 47 and 5.5-48 show the details of this, and Figure 5.5-49 is representative of the results. Spacecraft motion is duplicated well, yielding a nutation period of 14.4 seconds and a total angular rate of 417 deg/sec, which is very close to the observed spacecraft data of 13. 0 seconds and 448 deg/sec. However, the initial pitch and yaw gyro transients produced were faster than observed in flight. The roll gyro trace provides a good match, peaking at I. 5 degrees and crossing over at 3.5 seconds. This simulation did not incorporate any special large-signal electronics features. The gyro electronics amplifier and demodulation have saturations and gain changes well below the level of the gyro stops which attenuated actual telemetry signals. Mixed Simulation. Tests to determine the characteristics of the
electronics showed complicated saturation nonlinearities and transient characteristics which would require a vast amount of equipment to duplicate on an all-analog simulation. A mixed simulation incorporating all suspect parts of the electronics, as shown in Figure 5. 5-50, was set up. The exact duplication of SC-Z electronics was not possible since dynamic ranges are required only to be greater than some minimum level, and thus are not controlled. The use of actual hardware command switching was also made possible with this setup. By assuming small engine delays and gyro initial conditions, Figure 5. 5-51 was obtained. It is the best combination of engine delays and initial gyro angles within known tolerances and knowledge of SC-Z electronics. Figure 5. 5-52 also is derived from the _nixed simulation, but has slightly different engine delay and somewhat different initial conditions. The best match was obtained with initial gyro angles in a direction to reduce startup thrust, with engine 2 ignition lagging engine I. The thrust traces show the same form as that of the SC-2 data, but engine Z is higher than indicated by the telemetry. The gyro traces match well with flight data. At 6 seconds, the yaw acceleration changes sign, and there is a thrust dip due to the acceleration loop. This did not occur in the flight data, and is the result in the thrust profile differences at the beginning. However, the period of final oscillation and total angular rate is the same as SC-Z, indicating the integral of thrust is correct. The roll gyro and roll actuator angles match well with the flight data. Analog Only. With the knowledge gained by using the mixed simulation, a second all-analog computer simulation was attempted. However, the spacecraft gyro amplifier and demodulator were used in the recording of the gyro traces, as shown in Figure 5. 5-48. A passive network was used for the attitude loop shaping. The best long-term results were obtained from this setup. The period of nutation was 13 seconds, and the angular rate was 432 deg/sec. The roll gyro and actuator are very close to an exact match. Figure 5. 5-53 term results. motion. shows Figure the initial transient, and 5. 5-55 includes vehicle Figure angular 5. 5-54 shows the longrates to detail the
dynamics assumption
Thus, the provides that
present a close engine
computer match
with no
model of observed thrust.
spacecraft SC-2
and behavior
flight control under the
3 produced
5.5-150
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5. 5-49.
Open
Loop
Simulation
Results
5. 5-153
68189-g-331
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Figure
5. 5-53.
Final
All-Analog
Simulation
Results:
Initial
Transients
5.5-159
Figure
5. 5-54.
Final
All-Analog
Simulation:
! <,_g-Term
Results
5. 5-160
Figure 5. 5-55.
Final All-Analog
Simulation Results:
Vehicle Angular Rates
5. 5-161
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+ +-.-+
if{ i l! 777
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if. 7. II. l. 1_
Figure
5. 5-56.
Digital
Simulation:
Thrust
Command
T1
"+4q :_'"" t ..... p
',::_.
-::r :i!iH],!- ii:
]_H*,It;r
"+!iiiiiH :::".....
.... iii4tti!_ :-:
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...........
:m ++_,+ ......... i{_{ ii!il!_::
+.+++_ ........ ,<,< v+.,_ ....... _F!i4 ! -+-, +_-+4 .........
.....
Figure
5. 5-57.
Digital
Simulation:
Thrust
Command
T2
5. 5-162
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; .
.
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L
I', _',,_,I._' _ ........ .... I,,' i l ....
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Figure
5. 5-58.
Digital
Simulation:
0
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Figure
5. 5-59.
Digital Simulation:
0y
5.5-163
TC
: I ¸
i
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Figure
5. 5-60.
Digital
Simulation:
-0
z
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: i ! I LII
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:_:
5. 5-61. Digital Simulation: -Delta
5. 5-164
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\
5
REFERENCES
"SPAC Report, SC-Z II October 1966. W. S. Hicks,
Flight
Control
Performance,
" IDC
2255.
1/1803,
"Preliminary
Analysis Report No.
of AC-7 GDC-BNZ
Flight
Data, 28
" Genera[ September
Dynamics/Astronautics 1966.
,
66-053,
"SC-Z
Telemetry
Calibration
Handbook, 1966.
" Hughes
Aircraft
Company,
Specification
.
291032,
iZ October Gas 19 Jet
R.H.
Bernard, IDC
"Revised 2223/843,
Consumption 1965.
for
66
Hour
Surveyor
Flight,"
.
February
"Surveyor System Functional Specification 224510 Revision L.S. Crowell, 23 November "Analysis 1966.
Requirements, E, 7 January Star
'_ Hughes 1966.
Aircraft
Company,
6_
of SC-2
Acquisition,"
IDC
2223/2570,
.
R. H. Bernard, "Limit Cycle Behavior Control System with Simulated Canopus Analog Mechanization," IDC 2223/772, O.N. Hertzmann, IDC 2223/2341, "SC-I June Mission 1966.
of the Surveyor Gas Jet Attitude Sensor Noise Included in the 15 February 1965. -- Flight Control Report, "
.
Performance
23
.
K. Kobayashi, "A Mission Analysis,"
Method of Determining Gas Jet Thrust IDC 2253. 4/25, i March 1966. Mission Analysis HandbookSSD 68117R, 5 May 1966.
Level--
Post
10.
"Surveyor Spacecraft Post Hughes Aircraft Company, "Spacecraft Flight Detail Specification Control, 234600
Flight
Control,
"
Model A-Zl," Hughes Aircraft Revision E, I0 June 1965.
Company,
IZ
E.I. Axelband, "Analysis Midcourse Thrust Vector," W.T. York, L.M. 1966. S. Thomson, 1961. Bronstein,
of Inertial Pointing Accuracy of Surveyor IDC 2242/2206, 17 June 1963. to Space Dynan_ics, John Wiley, INew
13_
Introduction
14.
"Principal
Axes
of SC-2,"
IDC
2292/114,
8 November
15.
Kubo,
"FCSG
Z35000-9
S/IN
I Retest
Data,"
IDC
ZZZ3/2298,
14 June
1966
5. 5-167
16. 17. 18. 19. 20. 21.
22.
E.I. Axelband, "Surveyor 2242/Z780, i November E.T. 2255. H.T. P/N Pfund, 1/1786, "SPAC Zl SC-Z
Coast 1963. Quick 1966.
Phase
Attitude
Control
System,
_' IDC
Quick-Look
Report
No.
2,"
IDC
September
Lew, "Flow 235700," IDC
Versus Altitude Thrust, Surveyor 324Z/Z423, iZ July 1963.
Gas
Jet
Valves
--
R. I-I. Bernard, Analog Computer R.O. Engine Crook, Logic
"Restoration and Mechanization, "Effect of Absence
Updating of Surveyor Coast Phase _'IDC 2223/77, 29 July 1964. of Pull-Down Resistor in 1966. Valve Vernier
Circuit,
_'IDC
2223/2415,
8 Septe_l_ber
B. IN. Smith, "SC-2 Modulation Test," R.O. Ii,"
Failure Review IDC 2223/2476,
Board 14 and
Flight Control/Prop 15 October, 1966. Board Action Item
Crook, _'Reply IDC 2323/Z477,
to SC-2 Failure Review 27 October 1966. SC-2 27 Failure October Re\-iew 1966. Review 1966. Review 1966. SC-3
10-19-
23.
R.O. Crook, '_Reply to 10-19-9," IDC 2223/2478,
Board
Action
Item
24.
R.O. Crook, "Repty to SC-2 Failure I0-19-8," IDC 2223/2540, 27 October R.O. Crook, "Reply to SC-2 Failure 10-Z7-Z, " IDC 2223/2542, 31 October S. Kubo, "Special Tests for Surveyor IDC 2223/2544, i November 1966. B.N. Supply 1966. Smith and R on O Croo
Board
Action
Item
25.
Board
Action
Item
26.
Post
Mission
Analysis,"
27.
Voltage
Solenoid
k, "SC-Z FRB Action Valve Operation,"IDC
Item II-3-X 2223/Z552,
Effect of I0 l'4ovember
28.
R.O. Crook, '_Reply to SC-2 Failure Review Board 10-27-4," IDC 2223/2553, I0 November 1966. S. Kubo, "2ZV Current I 1 November 1966. W. IN. Turner, IDC 2223/2555, Due to Roll Actuator,_r IDC
Action
Item
29.
2223/2554,
30.
"SC-2 Gyro 14 I'_ovember
and
Thrust 1966.
Command
Telemetry
Accuracies,
31.
S. Kubo, "Special Test for Surveyor IDC 2233/2558, 14 November 1966. S. Kubo, "Operation IDC 222.3/2447, 16 of the iNoxember Roll
SC-2
Post
Mission
Analysis
If, :'
32.
Actuator 1966.
at
Reduced
Bus
Voltage,
''
5.5-168
\
33.
H.D. Marbach, Startup," IDC L. S Crowell, 23 November
"Rol[ Gyro and Actuator 2223/2534, Zl November "Analysis 1966. of S/C 2 Star
Response 1966. Acquisition,"
at SC-Z
Midcourse
34.
IDC
2223/2570,
35.
B.N. Smith and M. R. Buehner, The Two Second Burns, and The 28 November 1966.
"SC-Z TM 20 Second
Strain Gage Burn)," IDC
Data (Midcourse, 2223/2533,
36.
P.L. Welton, "Results of Mixed, FCE/Analog of SC-Z Midcourse Thrusting," IDC 2223/2562, B.N. Smith (Midcourse, 2223/2563, and M.R. The Two 2 December Buehner, "SC-Z Vernier Second Burns and The 20 1966.
Computer Simulation 30 November 1966. Engine Second Ignition Burn)," Times IDC
37.
5. 5. 6
ACKNOWLEDGEMENTS
This responsible
section for the
was coordinated contents are: J.R. L.R. H.D. M.R. P.L. R.H.
by
B. N.
Smith.
Those
directly
Angerman Stumpf Marbach Buehner We[ton Bernard
5.5-169
5.6
VERNIER
PROPULSION
SUBSYSTEM
5. 6. I
INTRODUCTION System Description
5. 6. 1. 1
variable percent composed 5. 6-1). assemblies a variable
The Surveyor vernier propulsion system (VPS) is a bipropellant, thrust, liquid rocket system utilizing an oxidizer composed of 90 nitrogen tetroxide and 10 percent nitric oxide (Mon 10) and a fuel of 72 percent monomethyl hydrazine and 28 percent water (Figure The VPS consists of three regeneratively cooled thrust chamber (TCAs) with radiation cooled expansion cones. Each TCA has range of 30 to 104 pounds vacuum thrust. Propellant is supplied to the TCAs from six tanks employing bladders. One fuel tank and one oxidizer tank supply each located adjacent to the TCA near each of the three spacecraft positive TCA landing
expulsion and are legs.
Propellant expulsion is accomplished by pressurizing the propellant tanks on the gas side of the bladders withhelium gas. The helium is stored under high pressure in a spherical pressure vessel. The helium tank, together with the pressure regulator, dual check and relief valves, and servicing connections, is mounted outboard of the spaceframe between landing legs 2 and 3. Thermal control of the VPS is both active and passive. Electric heaters are installed on two oxidizer tanks, one fuel tank, and on all propellant feedlines to the TCAs. Passive thermal control consists of the application of black and white paint and vapor-deposited aluminum to selected portions of the VPS, together with super insulation applied to the propellant tanks. The feedlines are wrapped with aluminum foil to deter heat loss. 5. 6. 1. 2 System VPS Purpose has three main functions correction during retro during and phase a Surveyor attitude control mission:
The 1)
Midcourse
velocity control
2.) Attitude
5.6-1
0" O0 HELIUM TANK TEMPERATURE O0 ",.0 I ! HELIUM TANK PRESSURE ,_ TEMPERATL.ItE
PRESSURE
CURRENT [_ DUMP VALVE
L c HARG I NO VALVE
--'-_
RELEASE VALVE
O'sc_U'_ECT _ R_E_E
OO-4
/
/
MEL,uM
OXIDIZER _ TANK
OXIDIZER I
I
FUEL
_
,_"
2N
'"/L'-5 _ I
I
rT_
b"l
TEMPERATURE I L_,_,
T_NK ;
L
TAiNK ;
_'1'
-_
t_2m_ ;
L
'_2_'
7
CoTh_ROA1NTLEM'_)LVEToR_I_V_
THROTTLE \ ALVE
_
I
2,__.._
_.._.J
3_
O
IZER
1
.
TCA TEMPERATURE I
2
3_
\
Figure 5. 6-1. Vernier Propulsion System Schematic 5.6-2
w \
3)
Attitude descent
control maneuver
and
velocity
correction
during
the final
A midcourse velocity correction may be required to correct initial launching and injection errors. The Surveyor VPS has the capability of providing velocity corrections up to 50 m/sec with sufficient propellant remaining to successfully land the spacecraft on the moon. The required correction is transmitted to the spacecraft in the form of a desired burn time at constant acceleration of 0. 1 g, which results in a thrust level of approximately 70 pounds for each of the three VPS TCAs. In addition to providing the required velocity change, the VPS also provides spacecraft attitude control during the mane uve r.
Attitude control during firing of the spacecraft retro motor is provided by the VPS. The VPS is ignited approximately I. l seconds prior to retro ignition. Attitude control by the VPS is biased around a total vernier thrust level of either 150 or 195 pounds, depending attitude and velocity at retro burnout. The transmitted maneuver on predictions desired vernier of spacecraft thrust level is
to the spacecraft several minutes prior to initiation of the retro sequence. After retro burnout, the vernier thrust level is increased slow the spacecraft to allow the ejected
to 267 pounds total thrust to further retro motor case to fall clear. Following retro pounds total thrust motor under
ll0
ejection, the radar control.
VPS is throttled to approximately -When the spacecraft intersects
the first '_descent segment, " the VPS, operating in the closed-loop mode with the radar system, "acquires" the predetermined altitude-velocity profile and keeps the spacecraft on the profile. Each succeeding segment of the profile is acquired in a similar manner. At an altitude of 13 feet, the VPS is shut down and the spacecraft free falls to the lunar surface. 5. 6. i. 3 General Prelaunch Performance Summary
Final l September
propulsion 1966 when
preparations for propellant loading
the SC-2 launch were begun on of the vernier subsystem was
initiated. A total of 182. 4 pounds was loaded, of which 72. 2 pounds of fuel and 108. 1 pounds of oxidizer were usable (Reference i). Preloading calculations of the SC-2 propellant capacity (see subsection 5. 6. 4. l) indicated a total load of 182. 50 pounds, of which I08. 2 pounds of oxidizer and 72. l pounds of fuel were usable. The slight differences noted are well within the specified loading tolerance of Reference I. The helium psia at 68°R. tank was Telemetry charged on readings II September 1966 to a pressure of of the tank temperature and pressure check and prelaunch (see subsection 5. 6. 4. 2).
5160
were taken on 16 September. telemetry data, an "on pad"
Based on this telemetry leak rate was calculated
5.6-3
The calculated leak rate was negative, indicating that any leakage during this period was less than the telemetry sensing accuracy. During the joint flight acceptance composite test, high pressure helim leakage was measured at 0. 7 psi/day, which is in agreement with leakage below the telemetry sensing capability. at 75°F. to 85°F. Thermal conditioning of the spacecraft prior to launch was maintained Two hours prior to launch, the shroud temperature was increased
Table 5. 6-I compares the predicted propulsion temperatures with the actual stabilized values just prior to increasing the shroud temperature to 85°F. All temperatures were within the shroud temperature tolerance, and all propulsion parameters appeared normal at liftoff. TABLE 5.6-I. ACTUAL VERSUS PREDICTED TEMPERATURES Prelaunch Temperature Sensor P-4 leg 2 line P-5 leg 2 fuel tank P-6 leg 3 oxidizer tank P-7 leg l TCA P-8 leg 1 line P-9 leg 3 line P-10 leg 2 TCA P-f1 leg 3 TCA P-13 leg l fuel tank P-14 leg 3 fuel tank P-15 leg 1 oxidizer tank P-16 leg 2 oxidizer P-17 helium tank tank Actual,
degrees 71.3 70.1 70.2 71.2 71.5 71.0 70.2 70.4 70.7 70.3 70.8 71.0 71.2 Predicted, degrees 75 75 75 75 75 75 75 75 75 75 75 75 75 Premidcour Actual, degrees 15 se
Predicted, degrees 20-27 42 49 65 19-29 20-23 8O 7O 57
46
43 53 20 20 88 64
56
53 48 34 71
56
49 37 75
5.6-4
Coast
Phase
I (L
+ 30M
to L
+ 15H
45M)
The initial postinjection spacecraft interrogation indicated that all propulsion parameters were normal. Indication of heater operation on the leg Z and 3 feedline heaters was noted at 13:42 and 14:26 GMT, respectively, The temperature drop rate on the leg i line was considerably slower, and the heater did not start cycling until 21:40 GMT. Helium pressure increased from 5168 psia at 71.2°F at L - 2. 5H to
5174 psia at 73°F at L + 15H45M (see Figure 5. 6-2). Leakage calculations (see subsection 5. 6.4.3) indicate a leakage rate of 776 standard cc/hr. The short interval (18.4 hours) used in this computation, coupled with the telemetry sensitivity, place a low confidence in this value. In future reports, similar leakage calculations will not be made over intervals of less than 80 hours The oxidizer system pressure, as indicated by the leg 3 oxidizer transducer, dropped from 215 psia at Z - Z. 5H to 203 psia at L + 15. 5H, just prior to premidcourse maneuvers (see Figure 5. 6-3). Concurrent with the 12-psi pressure drop, the average oxidizer tank temperature dropped from 70 to 45 °F, causing both a decrease in tank ullage temperature and an increase in tank ullage volume resulting from propellant density increase. The pressure profile is similar to that of SC-I (Reference 2). Deviations from the nominal sun during gyro drift measurements, changes of the leg 2 TCA and line. patterns on the TCA and line, causing spacecraft attitude, with respect to the resulted in temporary temperature The attitude deviations altered the shadow the temperature changes.
rate,
At L + 3.5H the heater on
(16:00 the leg
GMT) after cycling 2 line remained on
at a progressively slower while the line continued to cool. drift check and then continbits/sec data, the line tem-
The line temperature briefly rose during a gyro ued cooling. Just prior to the initiation of 4400 perature was 15°F (see subsection 5. 6. 2. 2).
At 17:00 GMT, the leg 2 oxidizer tank was decreasing in temperature slightly faster than had been predicted. At that time, the leg 2 oxidizer tank was indicating 47°F as compared with the 54°F predicted. The actual indication was well within the predictability range, and the only possible effect of the increased temperature drop would be the possibility of enabling the propellant tank heater earlier than scheduled. The most probable cause of the increased temperature drop rate is that the insulation on the tank was more tightly wrapped than on SC-I.
5.6-5
I-
HELIUt4 PRESSURE
"-
"
...........
r-
"
+
"
[ ........
T.....
-t....... L i......._ ........................... .......
] 1
4----i-
...... i................ ,i-i" ,i,,z.,4, ....
r>...... I
d.
..
_ ,
+
1i -4.
i
-0.
I
,t
,t
llz.
,'
_.
_.
Figure
5. 6-2.
lViission
l_lot of Helium
Pressure
\¥itl_out Voltage
Correction
I
_i!i!
Figure
5. 6-3. Without
Mission Reference
Plot
of Manifold
Pressure
Voltage
Correction
5.6-6
Midcourse
Operations
Propulsion system condition just prior to midcourse was normal, and all parameters were within their allowable range. The leg 2 oxidizer tank indicated 34°F, and the leg 2 line indicated 15°F. The helium release squib was actuated at L + 16HZIM39S, and the propellant tank pressure increased from 198 to 769 psia immediately and locked up at 777 psia prior to the midcourse correction. Corrections to this figure indicate a lockup with pressure the 765 of 772 to 775 psia psia (see subsection during 5. 6. 4. 5). regulator This flight compares acceptance favorably test. recorded
At helium release squib actuation, the helium tank pressure 739 psi from 5126 to 4387. The predicted drop was 206 psi. This was caused by ahelium transducer zero shift experienced at squib (see subsection 5.6. 2. I). At 265:05:00:02 GMT, vernier ignition was commanded
dropped difference actuation
on for
a
planned 9. 81-second firing. The leg 3 TCA appeared not to ignite, and the resulting unbalanced moment from the other two TCAs caused the spacecraft to tumble. At the end of the firing, the spacecraft was tumbling at approximately one revolution per second. Since the tumbling rate exceeded the cold gas system correction capability, the gas jets were turned off shortly after firing was terminated. The standard mission ended at this point. 5. 6. I. 4 Major Vernier System Events of the major events concernTable 5. 6-3 summarizes
Table 5. 6-2 lists the time of occurrence ing or influencing the vernier propulsion system. all anomalies affecting the propulsion subsystem.
5. 6. 2
ANOMALY Pressure
DESCRIPTION Transducer Zero Shift
5. 6. 2. i
At helium release squib actuation, an abnormally large pressure droop was noted on the helium tank pressure transducer. Based on computed ullage volumes (see subsection 5. 6. 4. 4), the predicted pressure drop was calculated at 206 psi; the measured drop was 739 psi. A frame-by-frame examination of the data showed a 533-psi drop in helium tank pressure between two consecutive samplings of the helium tank pressure, indicating a flow rate far in excess of system ability. The helium tank pressure decay and the propellant tank pressure rise agree well with experience and with SC-I behavior. Therefore, the instantaneous drop exhibited by these two consecutive telemetry readings indicates a zero shift in the transducer. The helium tank pressure decay and propellant tank pressure rise transients for both SC-I and SC-2 are plotted in Figure 5. 6-4. The corrected pressure
5.6-7
TABLE
5.6-2.
MAJOR
VERNIER
SYSTEM
EVENT
COMMAND
TIMES
GMT, Event Launch Pressure Midcour VPS se ignition day:hr:min: 263:12:32:00 264:04:53:38 264:05:00:02 sec
Mission Time, hr:min:sec 00:00:00 16:21:38 16:28:02
End Dump
of
standard helium
mission 265:09: 265:09:34:17 13:I6 44:41:16 45:02:17
Emergency AMR command Vernier (FC-28 ignition telemetry)
265:09:34:27.
2
45:02:27.
2
Note:
A complete course is
listing given in
of all system
vernier engine subsection 4.
firings 1.
after
mid-
TABLE
5. 6-3.
ANOMALY
SUMMARY
TABLE
Anomaly
Number Helium tank
Anomaly pressure a 533-psi squib cooling heater transducer zero shift actuation. prior to operating. at
experienced helium release
Leg 2 line was course with the
mid-
Leg 3 TCA appeared not to ignite at midcourse, causing the spacecraft to tumble.
5.6-8
Z
Figure
5.6-4.
Helium
Squib
Release
5.6-9
drop taken from this figure is 211 psi, which compares favorably with the computed value of 206 psi. A comparison of pressure decay and rise rates between SC-I and SC-2 shows that flow through the regulator was about equal for both spacecraft. This agreement is evidenced by the helium tank pressure decay curve slopes which are nearly the same; the SC-I value is -150 psi/sec, and the SC-2 value is -141 psi/sec. A comparison of the two propellant tank pressure rise curves indicates that the propellant tank ullages also were nearly the same; the SC-I rise rate was 354 psi/sec, and the SC-2 rise rate was 344 psi/sec. From this, it is concluded that a zero shift did take place and the VPS pressurization sequence was normal. A zero shift of this type was noted in two cases during the vernier system development program (Reference 3). Both shifts were less than 200 psi, and a note inserted in the spacecraft signature list indicated a shift of up to 4-150 psi could be expected (Reference 2). The zero shift is caused by shock loading the transducer during squib actuation, and is a somewhat random function. For future spacecraft, the helium tank pressure will be displayed on an analog recorder so that any zero shift will be readily discernable. 5. 6. 2. Z Leg 2 Line Heater Cycling Termination
The leg 2 feed line assembly indicated 85°F at launch. The line temperature dropped to 20°F at L + iH and began to cycle between 20 and 25°F as the heater thermostat began to operate. The thermostat cycled four times between L + 1 and L + 3H. Each "power on" cycle was longer than the last. At L + 3H34M, the line temperature appeared to stabilize at 24°F, which was below the thermostat opening temperature of 25°F. The line then began to cool; just before the premidcourse maneuvers, it had reached a temperature of 15°F. A gyro drift check from L + 7 to L + 9H caused the line temperature to rise slightly, but cooling resumed at the termination of gyro drift check.
Thermal analysis concluded (see Section 5. i) that the line heater was on during the cooling period and that the heat input from the heater was less than the heat loss from the line to space. To prohibit recurrence of this problem on future spacecraft, minimum duty cycle criteria are being established for heater operation during STV testing to prevent a line with marginal thermal characteristics from being accepted. 5. 6. 2. 3 At maneuver that the firing, control and the tumble Midcourse L + was Anomaly (264:05:00:02 Detailed GMT) a 9. 85-second data midcourse review indicated
16H28M02S initiated.
system-by-system
VPS leg 3 engine failed to ignite, while at least during the midcourse leg 1 and leg 2 engines behaved properly (Reference 5). The flight system immediately throttled leg i and 2 engines to minimum thrust leg 3 engine to maximum thrust; however, the vehicle began to and, at the end of midcourse, was tumbling at approximately 1 cps.
5.6-10
Preliminary SPAC/SCAT analysis of oxidizer line, engine temperature, strain gage, and helium and leg 3 oxidizer temperature data (during telemetry mode 2 32 minutes after midcourse) indicated normal behavior of legs 1 and 2, but no ignition on leg 3 TCA, probably due to failure of fuel to flow to the leg 3 TCA. Following midcourse, a series of 0.2-second and 2. 0-second pulse firings were performed, with a final 21-second firing (see Table 5.6-3). While detailed analysis has not resulted in a conclusive diagnosis of the failure to ignite, some conclusions are pertinent regarding the leg 3 vernier engine:
i)
There during of some
was all
evidence after flow
of
oxidizer at midcourse, midcourse.
flow
at and
less less
than
commanded indication
rate
firings oxidizer
conclusive
z)
Through direct evidence or by demonstrated for all firings. There firings thermal was no ignition showed which indicated
inference,
fuel flow
can
be
3) Subsequent quantitative
on
any
firing anomalies at length
attempt. as detected Section 5. by 1.
possibly analysis,
minor random is discussed
in
5. 6. 3
SUMMARY Summary summary analysis, Conclusions
AND
CONCLUSIONS Effort parameters, along with as determined from predicted values.
5. 6. 3. I A postflight 5. 6. 3. 2
of Analysis
of the VPS performance is given in Table 5. 6-4
the
The i)
following Excessive actuation
conclusions helium was due
are
given: drop noted at release zero shift. squib
tank pressure to a transducer
2)
Cooling of the the line having Propulsion periods. Positive valuable
leg 2 line with the heater operating marginal thermal characteristics. availability was insufficient during
resulted
from
3)
data
thrusting
4)
indication during
of fuel
subsystem
pressure
would
have
been
investigation
of the SC-2
failure. of less than 80 hours in future reports.
5)
Leakage calculation over time intervals are not valid and should not be repeated
5.6-11
TABLE Item Fuel loading loading
5. 6-4.
ANALYSIS Predicted
SUMMARY Actual 73.06 109.34 211 pounds pounds
73. 00 pounds 109. 50 pounds 206 psi
Oxidizer
Helium consumption at squib release Regulator lockup
psi': _
765
to 775
psia
772
psia
;:_Corrected
for zero
shift.
5. 6. 3. 3
Recommendations The i) following Line and recommendations TCA temperatures are made: be available during subject lines. be added propulsion thrusting. to more
should
2) 3)
Line heater performance during STV should be stringent acceptance criteria to detect marginal A fuel subsystem vernier system for both transit pressure telemetry and lunar measuren_ent to provide operations. should additional
to the data
5. 6. 4
SUBSYSTEM Predicted
PERFORMANCE SC-2 System system Propellant
ANALYSIS Loads
5. 6. 4. 1
Oxidizer SC-2
oxidizer
total volume 6 and and 7)
Vto t = 2228. Unusable volume
7 in 3 (References trapped 3 in lines
TCAs
Vtr Unusable V
e
= 12. 6 in volume
(Reference due
8) bladder expulsion inefficiency
to 0. 5 percent 8)
= I i. 1 in 3 (Reference
5.6-12
Loading tolerance = 0. 75 pound
V For calculated zero ullage tolerance at usable = Vto t - Vtr - Ve - loading tolerance oxidizer Loading A -30is based loading
worst-case 0°F,
the
conditions, minimum maximum
the expected
weight of unusable temperature. temperature.
is
on
at 105°F, the is also included.
expected
Wox
usable
= Vtot = (2228. -0.75 = 108.21
( Pox 7)
105°F) (0. 04947}
- Vtr -(12.6)
(Pox (0.
0°F) 05437)
-
Ve
(Pox 1)
0°F)(0.
0. 75 05437)
- (11.
pounds
Fuel SC-2
System fuel system 0 total volume 6 and and 8) bladder expulsion inefficiency 7)
Vto t = 2229. Unusable Vtr Unusable V
e
in 3 (References trapped in lines
volume = 12. volume = 11. 0 in 3
TCAs
9 in 3 due
(Reference to
0. 5 percent 8) pound
(Reference = 0. 75
Loading
tolerance
VtotNET
= V usable
+ Vtr
+ V e
Fuel of I. 5.
loading
is
based
on
a nominal
oxidizer-to-fuel
mixture
ratio
W W f The total net fuel usable -ox
usable _ 1.5 108.21 1.5 72.14 pounds
load
is
WfNET
= 72.
14
+ (12.
9)
(0.
03586)
+ (11.
0)
(0.
03586)
= 73.
00
pounds
5.6-13
For a tabulation of predicted SC-2 loads at [05°IV, see Table 5. 6-5.
To determine the amount of propellant to be offloaded to compensate for the lower than maximum loading temperature, the total loaded propellant must be determined at the loading temperature of 70°F.
Oxidizer W loaded System = Vto t (Pox 70°F)'Pox = oxidizer density
= (2228. W offload--W70oiv-
7) (0. 05109) W105o F
= 113. 86 = i13. 86-
pounds if0. 25 = 3.61 pounds
Fuel W
System -- Vto t (pf 70 °F),pf = fuel density
loaded
= (2229.
0) (0. 03450)
= 76. 90
pounds
W
offload
= W70OF
W105
°Iv + W
offload
105 °F
--76. 90
- 75. 14
+ I. 66 For a comparison 5. 6- 5.
-- 3. 15 pounds of predicted versus actual SC-2 loading, see
Table
5.6-14
TABLE 5. 6-5. ACTUAL VERSUS PREDICTED SC-2 PROPELLANT LOADING SC-2 Predicted at 105°F
Oxidizer Fuel 75.41
SC-2 Predicted at 70oF
Oxidizer I13. 86 Fuel 76. 9O
SC-2 Actual at 70°F
Oxidizer 116.42 Fuel 76. 15
Total loaded gross, pounds 3_ loading tolerance, pounds Offload, pounds Total loaded net, pounds at i. 5 mixture ratio Unusable at 0°F, pounds Total usable, pounds at I. 5 mixture ratio
5.6.4.2 PV where helium helium helium helium helium helium
110.25
0.75
0.75
0. 75
0.75
0.75
0.75
0 I09. 50
1.66 73.00
3.61 109. 50
3.15 73.00
6.33 109.34
Z. 34 73. O6
1.29
0.86
I. 29
O. 86
1.27
0.85
108.21
72.14
108.
21
72. 14
I08.07
72. 21
Prelaunch = WZRT
Helium
Leakase
p
__
tank tank tank
pressure, volume, temperature, in
psia 3
V T Z R W
= = = = =
°R factor
compressibility gas tank constant gas weight,
pounds
5.6-15
B_
Ln(P) Differentiating dW W Dividing by dW dt From Mode
+ Ln(V) and dP P dt, time
= Ln(W) dV,
+ Ln(Z) dZ, and
+ Ln(R) dR = 0
+ Ln(T)
using dT T
W dP P dt 2 telemetry 259:18:22
W dT T dt
GMT
PI
=
5168
psia
(4 Clays before
launch)
T I 263:09:51 GMT P2 T 2
=
531. 7 ° R psia (3 hours before launch)
-- 5168 =
531. 2°R
PAV
=
5168
psia
TAV
=
531. 5°R
Z VHB
-=
1.
17
(Reference in 3 based on
10) expansion data of burst tanks
1300
WAV dP dt
-
PV ZRT
-
(1.
(5168)(1300) 17)(386)(12)(531)
= 2. 335
pounds
=
5168-5168 84. 3
= 0 psi/hr
dT dt
531.
2 - 531. 84. 3
7
-- -0. 00593°R/hr
dW dt
:
2. 335 0 - -- 531
(-0. 00593)
=
+0.00002608
ib/hr
Any
leakage
is
below
the
telemetry
sensing
capability.
5.6-16
5. 6. 4. 3
Coast
I Helium Mode
Leakage
From
2 telemetry GMT PI = 5168 psia (3 hours before launch)
263:09:51
T
I
=
531. 2 ° R
264:04:15
GMT
P2 T 2
= = TAV
5174
psia
(16 hours
after
launch)
533. 0 ° R = 532. l°R
PAV Z
= =
5171 i. 154
psia (Reference
I0)
VHB
=
1300 data
in 3 based on expansion of burst tanks
WAV dP dt dT d-'}-" dW dt
-
PV ZRT
-
(5171)(1300} (i. 154)(386. 2)(532.
I)(12)
=
2. 362
pounds
5174-5168 18.40 533. = 0-531. 18. 40 = 0. 3268 psi/hr
2 = 0. 0978°R/hr
=
2. 362 5171"
(0. 3268)
2. 362 (0.0978) - 532. i
= 0. 0001492
- 0. 0004341
=
0. 0002849 0. 0002849 0.01054
Ib/hr std hr
leakage std = 0.02705 hr cc = 766 std --_ cc ft 3
ft 3
= 5. 6. 4. 4 Helium Following
(0.
02705)(1728)(16. -
4) Squib
std _
Consumption the method
Release in Reference 11, initial gas weight is
outlined (5126)(1300}
WHT
1
=
(1.
16)(386)(12)(532.
+ 72. i = 532.
1)
= 2. 331
pounds
(at 460
I °R)
5.6-17
The
gas
volume
in the
propellant
tanks
(downstream
from 73. 06 0. 03494 3
the
squib)
is
VpT
= 2228.
7 + 2229.
109. 34 0 - 0.05244
= 4457. The initial propellant tank
7 - 2085. gas
0 - 2091.0 is then
= 281. 7 in
weight
Tox
T -
+ _f
o 42.8 + 51.6 2 = 47. 2 * 460 = 507. 2°R 2
g
+ WpT 1 = WOT 1 WFT
_ 1
after
(198)(281. (1)(386)(12)(507.
helium
7)
: 0. 0237 pound
2)
is
The
final propellant
tank
gas
weight
release
WpT2
The amount of helium W The corresponding =
= (1. 028)(386)(12)(532. (771)(281. 7)
transferred - 0.0237 is then = 0.0620 tank
l)
= 0. 0857 pound
0.0857
pound stabilized pressure at 72. I°F is
postrelease
helium
The
prestabilization
pressure
is found
from
PI I where n is the polytropic
= P2 2 exponent from Reference ii.
Since
V I : V 2
n
1.65
5126
= 4920
(0. 975)
psia
1.65
= 5126
4387
(0. 959)
psia recorded)
(versus
5.6-18
The discrepancy of -533 psia between predicted and recorded helium tank pressure must be due mainly to a zero shift in the pressure transducer (see subsection 5. 6. 2. I).
5. 6. 4. 5
Regulator
Lockup GMT
Determination
264:10:30
P2
= 776. 9 psia telemetry)
= 775
BCD
(from
mode
2
Reference Reference Unbalance The equation for
voltage, return, current, correcting
SI $2
= 998
BCD
= 0 BCD 131 BCD signals is
$5--
telemetry
TMcoRR where
= TMIN
D +
6LD
+
6A/D
+
61 +
6E
TMIN
D
=
actual
telemetry
reading
6LD
-- line drop
correction
6A/D
-- analog-to-digital
converter
correction
61
=
unbalance
current
correction
6E
=
reference
voltage
correction
5.6-19
Iunbalanc e (R 1 + R 2) (TMRE 61 (TMREF)
F - TMIN 2
D) (TMIN
D)
lunbalanc e = 131 BCD
: - 2. IZ_A
(-2.62 61
x 10-6)(2
x 103)(998-775)775
(998)(998)
(-5. 24 x 10-3)(223)(775) (998)(998) MV 88 _-_)
_ -0. 9092
x 103 volts
--
(-0.
9092
MV)/(4.
= -0.
186
BCD
TMIN 6E TMRE _
D F -(993-TMRF-F)
775 - 998 (993-998)
775 998 (-5) ---3. 89 BCD
6A/D
TMcoRR
= + 0. 5 BCD,
=
6LD
= -i. 33 BCD
(Reference
12)
775 - 0. 19 - 3. 89 + 0. 5 - I. 33 = 770 BCD
P2COR FAT
R
=
771. 9 psia lockup at 4950 at 4000 psig inlet -- 775 psia psig inlet = 765 psia
data indicates
5. 6-20
5. 6.4.
6
Midcourse Using the
Helium methods
Consumption outlined
Calculations 11
in Reference nz P op VHB V p
Ap
:
whe
re Ap
n z
pressure
drop,
Pinitial
- Pfinal' factor pressure, cubic inches
psi
polytropic exponent helium compre s sability op P propellant propellant ; helium tank volume operating
Vp
psia inches
expended, cubic
VHB
bottle
volume,
W Vp whe re V
OX
OK
Wf +-7 : x + Pf/ b
:
V ox
+ Vf
-
Pox
: :
oxidizer fuel oxidizer fuel time oxidizer fuel burn density, time, weight rate volume
volume
expended, Cubic
cubic inches pounds pounds
inches
Vf W
OX
expended, weight expended,
= -:
Wf
expended, of change of lb/in 3 lb/in seconds
weight 3
quantity,
lb/sec
Pox Pf z_t b
= = =
density,
]_rox
:
]Arox I + ]hroxZ + ]Srox 3
Numerical
subscripts
refer
to TCAs
i, Z, and
3.
5.6-Zl
nz Ap -
P op VHB
At
nz
Pop VHB
oxl
+ Woxz Pox
+ Wo×3
+
*fl
+ wfz
Pf
+ Ycf_
7
As suming
Poxl
Pfl
=
:
PoxZ
Pfg
=
=
Pox3
Pf3
_r oxl whe re F Isp is is TCA TCA thrust, specific
+
" Wfl
=
" W1TOT
-
F1 ISPl
pounds impulse, seconds
MR
-
ox i
1
whe re
wf 1
MR 1
=
TCA
propellant
mixture
ratio
F1 Wox 1 -isPl
Wox
1 I+MR1/Isp 1
MR 1
Similarly
Wfl
:
+MR
ISPl
5.6-ZZ
Substituting
into the
original
equation
Ap
nz _%[t_, _o_ _ [
#_ _-_7_ +k_÷_] _
Pf
(__
Oox
+
MR3 I_M--R;]._ F3 IsP3
+_ _-q
For
normal
operation
nz Ap _
P op V
At HB
MR =
n
Fn
+MR
Poxn
_-_pn +
m=l
+ R
m
Pfm
>1
F m
For
no flow
from
I TCA3:
Ap
v_
Po_ntb _
_ v..v._I _-_n+ =
Poxn
I
Fm
P fro
From
SC-Z
flight data
and
SC-Z
TCA
log books
F I commanded F Z commanded F 3 commanded IsPl ISpZ IsP3
= = = = = =
73. 6 pounds 52. 6 pounds 104 265 Z66 Z75 pounds seconds seconds seconds
5.6-23
MRI MR2 MR3 P op n
Z
= =
1.49 I. 54
-- 1.55 -- 754 -- I. 59 = 1.17 9. 8 seconds psia
At b
VHB Toxl = 56 °F Tfl
=
1300
cubic
inches
-- 57 °F
Po×|
-- 0. 05205
ib/in 3
Pfl
=
0. 03485
ib/in 3
Tox 2
=
38 °F
Tf2
=
46 °F
Pox2
=
0. 05288
ib/in 3
Pf2
=
0. 03506
Ib/in 3
Tox For
3
=
46 °F
Tf3
=
54 °F flow on
Pox3
=
0. 05250
ib/in 3
%3
=
0. 03491
Ib/in 3
normal
commanded
all three
TCAs:
_P
= (I. 59) (i. 130017) (754) (9, 8) [3. 191 + Z. 266 + 4. 379 %- 3. 202 + Z. 223 + 4. 3481
io 7119 2o6psi o81
For assumption of no leg -_P For assumption of no leg _P For assumption of no leg Ap The results of these 3 oxidizer = I0. 57 flow (15. 13) -- 160 psi 3 fuel flow = 10. 57 (15.26) or = 161 psi
3 oxidizer = i0.57
fuel flow -- i15 been psi summarized in
(i0.88) have
calculations
Table 5. 6-6. From the measured pressure drop of 168 psi, it can be concluded that the propellant flow on the leg 3 TCA was nearly equivalent to normal oxidizer or fuel flow above, but not both.
5.6-24
TABLE 5. 6-6.
CALCULATED HELIUM CONSUMPTION FOR VARIOUS MIDCOURSE FLOW ASSUMPTIONS
TCA Propellant Flow As sumptions Leg l Normal Normal Normal Normal Leg Z
Normal Normal Normal Normal
Calculate
d
Measured Pressure Drop, psi (for comparison)
Leg 3 Normal Fuel only Oxidizer only None
Pressure Drop, Ap, psi
2O6 160 161 i15
168 168 168 168
5.6-25
5
REFERENCES Report - SC-2, " Rev. C, Hughes Aircraft
"Mass Properties Monitoring Company, SSD 64226R.
.
"Surveyor Aircraft
I Flight Company,
Performance SSD 68189R.
Final
Report,
" Volume
III, Hughes
°
E. Goller to Distribution, HT8VA TAT, '_ IDC 2227. T.B. Shoebotham to E. T.
"Failure Analysis I/I173, 3 December Pfund, "Zero
Transducer 1965.
S/N
7-5705
,
Shift of UPS
Helium
Pressure
Transducer
,
(PI), '_IDC
2227.
1/1330, "Results Thrusting,
17 February of Mixed " IDC
1966. FCE/Analog 2223/2562, Computer
P.L. Welton to S. C. Shallon, Simulation of SC-2 Midcourse 30 November 1966. R.A. SC-1, R.A. System R.A. System Laird and to Distribution, SC-2," IDC
°
2227.
"VPS 1/1543,
Propellant 13 April
Tank 1966.
Capacities
of
SC-6,
.
Laird to Propellant Laird to Propellant
Distribution, Inventory, Distribution, Inventory,"
"Unmanifolded " IDC 2227.
Surveyor Vernier 1/761, 2 June 1965.
Propulsion
°
"A21 and A21A/l14 IDC 2227. 1/1110,
Vernier Propulsion 29 September 1965. - QualificaNo. Division,
.
"Technical Manual, Operating and Maintenance Instructions tion Thrust Chamber Assembly, Model TD-339, " Publication 8984-H2A, Thiokol Chemical Corporation, Reaction Motors Denville, New Jersey, 15 December 1965. "Hydrogen Handbook, '_AFFTC, TR-60-19. "Surveyor the S-6 Vernier Program," System IDC
G.F. Pasley to Distribution, Utilization as determined 6 October 1965. 12.
by
TAT
Helium 2227.
1/1125,
"TM Correction Factors - Special Company Specification 227152.
Test
No.
62B,"
Hughes
Aircraft
5.6.6
ACKNOWLEDGMENTS
The
following J. J. T. Amelsberg W. B. Putt
people
contributed
to
the
propulsion
subsystem
section:
Shoebotham
5.6-Z6
5. 7
MECHANISMS
SUBSYSTEM
5. 7. l
INTRODUCTION section of the report is concerned with the mechanical spacecraft landing legs, omnidirectional antennas, and positioner (A/SPP). For purposes of this report, these are collectively defined as the mechanisms subsystem. constituting Landing Omnidirectional A/SPP Mechanisms operations of The automatic gear the main headings for this analysis effort performantenna/
This ance of the solar panel mechanisms Items 1) 2) 3)
include:
deployment antenna solar subsystem deployment panel deployment during nonstandard flight
4)
performance
was
Performance satisfactory. panel signals to from is
the above equipment and landing gear, omnidirectional were landing gear
functions
during antennas, the allotted antenna
the and
mission auto-
matic solar Telemetry
deployments for the
completed within and omnidirectional
time span. mecha-
nisms continued which resulted of the A/SPPwas performance
indicate normal conditions throughout the abnormal midcourse maneuver. not normal while the spacecraft was covered in the section on nonstandard the major mission events subsystem performance. from Reference 1. and
spacecraft spinning However, stepping spinning. This flight operations. times pertinent to All Centaur command the
analysis and event
Table 5.7-1 lists of the mechanisms data were taken
5. 7. 2
ANOMALY
DESCRIPTION
There DegradedA/SPP operations, and is not
were
no anomalies performance, attributable anomalous.
in the discussed to
mechanisms in the loading
section
subsystem performance. on nonstandard from spacecraft spinning
is fully considered
abnormal
5.7-I
TABLE 5. 7-1.
MISSION MAJOR EVENTS AND TIMES Mission Time, min:sec 0. 00 ii:50. 22 to 51. 20 11:51. 26 to 53. 96 12:00. 19 to 01. 17 GMT, hr:min:sec 12:31:59. 824 12:43:50. 044 to 51. 024 12:43:51. 386 to 53. 784 ]2:44:00. 014 to 00. 994
Event Launch Centaur extend landing gear command Legs extended (V-l, V-Z, and V-3 on) Centaur extend omnidirectional antenna command Omnidirectional extended (M-l, antennas M-Z on)
12:00. 56 to 02. 96 12:31. 76 to 34. 16 18:3 I. 89 to 34. 29 22:43. 88 to 46. 28
12:44:00. 386 to 02. 786 12:44:31. 585 to 33. 985 12:50:31. 717 to 34. ll7 ]2:54:43. 708 to 46. 108
A/SPP solar panel unlocked ( M-14 on) A/SPP solar panel relocked (M- 11 on) A/SPP roll axis relocked (M- 13 on)
5. 7. 3 SUMMARY AND CONCLUSIONS 5. 7. 3. 1
subsystem Performance Parameters and actual values for the mechanisms
Table 5. 7-2 compares expected performance parameters. Conclusions
5. 7. 3. 2
Mechanisms subsystem performance omnidirectional antenna deployment, and was excellent in all respects. No problem Landing leg deployment approval test deployment time respectively).
during landing gear deployment, automatic solar panel deployment was indicated. favorably with the type 2. 31 and 2. 34 seconds,
time compares (two deployments:
5.7-2
TABLE 5. 7-2.
PERFORMANCE
PARAMETERS
Parameter Time from Centaur extend landing gear command to legs extended indications (V1, V2, and V3 on) Time from Centaur extend directional antenna command omnidi r ectional antenna s extended (M-1 and M-2 on) time auto omnito
Expected Value, Nominal Z. 3 seconds
Measured
Value
I. 34 to 3. 74 s ec onds
Z. 4 seconds
0.37 to seconds
2.77
Solar axis deployment (A/SPP solar panel deployment) Roll axis deployment (A/SPP solar panel deployment) Total A/SPP solar time launch launch axis launch transit transit
365
seconds
360
seconds
time auto
255
seconds
252
seconds
panel
auto
6Z0
seconds
61Z
seconds
deployment Solar Polar Elevation Roll Solar Roll axis axis axis axis axis
position position launch position position position position
355
degrees
355.0
degrees
0 degree 0 degree -59.9 270 degrees
-I. 1 degrees 0. 0 degree -59.8 271.4 -0.4 degrees degrees degree
degrees
0 degree
The omnidirectional acceptance test data at -20°F; and omnidirectional antenna
antenna deployment omnidirectional B was l. 9 seconds.
time antenna
agrees Awas
with 2.4
the flight seconds,
times Table telemetry matic
Automatic solar panel deployment time corresponds recorded during SC-Z solar thermal vacuum retest, 5. 7-Z. Positions of the various A/SPP axes, using values, were near the expected values at launch solar panel deployment, again as shown in Table 5.
closely as shown corrected and after 7-2.
to the in auto-
5.7-3
J
5. 7. 4
DETAILED Landing
ANALYSIS Gear Deployment
5. 7. 4. I
Table 5. 7-3 shows the expected and actual times for the Centaur programmer extend landing gear command and indicates deployment completion. The uncertainty in actual times is due to the telemetry data sampling rates. The expected times are based on Centaur actual times and nominal landing gear type approval test deployment times. No anomalies of any type were noted concerning landing gear deployment. 5. 7. 4. Z Omnidirectional 5. 7-4 gives Antenna the Deployment and actual times for the Centaur pro-
Table
expected
grammer extend omnidirectional antennas command and time of deployment completion. The uncertainty in actual times is due to the telemetry data sampling rates. The expected times are based on Centaur actual times and nominal SC-? omnidirectional antenna flight acceptance test deployment times. No anomalies occurred in connection with omnidirectional antenna deployment. 5. 7.4. 3 A/SPP Automatic solar Solar panel Panel Deployment of the M-9, panel A/SPP was completed in the
Automatic
deployment signal solar
prescribed manner. at ?.63:12:44:25 - 27
Telemetry GMT. M-4,
vehicle unlock,
separation, occurred followed at 1Z:44:32-34.
Solar axis stepping commenced immediately and continued until solar panel relock, which initiated roll axis stepping. The solar panel relocked at 12:50:32 - 34, and the roll axis relocked at 12:54:44 - 46. It is not possible to determine precisely the response of the solar and roll axes motors to the applied stepping pulses since there is no means of counting the number of pulses applied to the stepping motors during automatic deployment. However, several indicators provide substantial evidence that the response of each axis was essentially 100 percent. Figure 5. 7-1 is a plot of the roll and solar angles versus time during
automatic solar panel deployment. Assuming the multivibrator pulse rate to be essentially constant, a significant number of n_issed steps would be indicated by a nonlinearity in the plot. A study of the plots shows no such nonlinearities. A comparison of the SC-2 automatic solar panel deployment data the corresponding data from SC-2 STV 2IB and STV retest (Table 5. 7-5) shows close correlation in deployment times. with
During SC-2 STV 2B, the automatic deployment was completed in I0 minutes and 24 seconds. The number of stepping pulses required was recorded (oscillograph of I_P-17), and responses of the solar and roll axes were calculated to be 97. 8 and 99. 3 percent, respectively. Panel deployment took 10 minutes and 20 seconds during SC-Z STV retest, but no response calculations were available since I_P-17 was not recorded. From the above comparisons, it can be assumed that the solar and roll axes responses were essentially i00 percent 5.7-4 during automatic deployment.
_D
O
O
O
M
_D
5.7-5
T
TABLE
5.7-3.
LANDING
GEAR
DEPLOYMENT
TIME
Event Centaur landing Legs (VI, programmer gear command extend
Expected, hr :min: s ec 12:43:51. 274
Actual, hr:min:sec 12:43:50. 044 to 51. 024
extended V2, and
indications V3 on)
12:43:53.
574
12:43:51.
386
to 53. 784
TABLE
5.
7-4.
OMNIDIRECTIONAL
ANTENNAS
DEPLOYMENT
TIME
Event Centaur programmer extend omnidirectional antennas command Omnidirectional extended (M-1 antennas and M-2 on)
Expected, hr:min:sec 12:44:01. 774
hr
Actual, :min: 014
s ec to 00. 994
12:44:00.
12:44:04.
174
12:44:00.
386
to 02. 786
Table andpostautomattc transit based SC-2
5. 7-6
shows deployment.
A/SPP
position Included
and also
related are the
data known
for prelaunch Iaunch and positions from the
locked axes positions on corrected telemetry Spacecraft Telemetry The maximum This difference result
and the corresponding data and calibration Handbook. between is reasonabIe predictions
calculated coefficients
and
measured the following
values known
was 1.4 degrees. SC -2 uncertainties: Potentiometer
considering
Calibration Solar Polar Elevation Roll axis axis axis
Curve
Errors,
rms
Degrees 0. 0. 94 32
axis
0. 45 0. 16 98
Other
signal
processing
errors
(all
axes),
rms
0.
5.7-6
TABLE
5.7-
5.
SC-2 MISSION, STV 2B, SWITCH CLOSURE TIMES
AND
STV
RETEST
Item MM-II M-13 14 on on on (solar (solar panel panel unlock) relock)
SC-2 Mission 12:44:33. 985
SC -2 STV 2B 21:14:01 21:20:05 21:24:25 364 seconds
STV
Retest
17:44:13 17:50:18 17:54:32 365 seconds
12:50. 34. 117 12:54:46. 360 108
(roll axis
relock)
Solar axis (M- II on Roll axis (M- 13 on
stepping time M- 14 on) stepping time M- l 1 on) time on)
seconds
252
seconds
260
seconds
255
seconds
Total deployment (M- 13 on M-14
612
seconds
624
seconds
620
seconds
5.7.
4.4
Mechanisms Operations
Subsystem
Performance
During
Nonstandard
Flight
ance
This subsection after midcourse(whenthe of this the retro period.
is
concerned with spacecraft Mechanism
mechanisms was spinning) telemetry
subsystem performuntil contact was signals were normal
lost
after firing throughout
rocket.
At 265:02:44:58, the polar axis was commanded in a positive direction for 240 steps. Since this axis is not pinned, there should have been a response, but telemetry indicates that no motion occurred. It is concluded that the stepping motor was unable to overcome the forces induced by spacecraft spinning. At 265:06:35. 19, command 0631, unlock This command was an error, since these start of automatic solar panel deployment. solar panel (transit), was pin pullers had been fired The command should have four stepping since the solar
sent. at the
been 0635, unlock solar panel (lunar). commands were sent to the solar axis axis was still locked in transit. At immediate
Two hundred and with no response,
265:06:54:33, the correct command, effect of this command was a jump
0635, in axis (see had
was transmitted. The position from 271. 4 to Figure aligned 5. 7-2). It is itself normal
249. 8 degrees, corrected for reference voltage likely that in the latter position, the solar panel to the axis of spacecraft a series of 87 negative
rotation at that time. Thirty solar axis stepping commands
seconds after unlock, was given. At eight
5.7-7
,r-4
D_ D_
O
o,-_
O
_D rj
_v
o
i
D-
_4
°r-4
5.7-8
TABLE
5. 7-6.
A/SPP DATA
PRELAUNCH/POSTDEPLOY (TELEMETRY MODE 4)
POSITION
Raw
Data Post
Corrected
Data-':: Post
Signal
Prelaunch, bcd
Deploy bcd
.........
Prelaunch, bcd
Deploy, bcd
M-3 M-4 M-6 axis M-7 S-I
solar polar
axis axis
903 378 52O
668 379 521
897 375 516
663 376 517
elevation
roll axis reference
336 999
502 i001
334
498
voltage S-2 reference return S-5 commutator unbalance current 120 119
Position
Indications
Based Corrected
on Calibration BCD Data
Coefficients
and
Prelaunch Predicted Angle, Signal degrees Indicated Angle, degrees Predicted Angle, degrees
Postdeploy Indicated A ng 1e, degrees
M-3 M-4 M-6 M-7
solar polar
axis axis axis
355 0t 0 -59. 9
355. 0 -1. i 0.0 -59. 8
270 Ot 0 0
271 -1.1 0.0 -0.4
4
elevation roll axis
'::Corrected per Test Requirement MS Specification 3023926 A. (Corrections conversion were not applied as these in the calibration coefficients. ) *':-'From ".:*'::From t Polar prelaunch countdown Z63:13:3Z:51 (telemetry axis not pinned; based data,
112
through for line corrections
MS drop are 1966,
117 in System Test and analog-to-digital already included 10:28:57.492 data. GMT.
20 September mode 4). on 8 September
1966
alignment
5.7-9
TAB
LE
5. 7-7.
A/SPP
STEPPING
COMMAND
LOG
Start Time, day:hr:min:sec 263:13:20:16 13:20:21 13:21:44 13:21:49
Stop Time, day:hr:min:sec 263:13:20:21 13:20:23 13:21:49 13:21:51
C ommand 0402 0401 0405 O4O6
Quantity 10 5 10 5
265:02:44:58 06:35:19 06:41:39 06:42:51 06:45:41 06:46:07 06:46:32 06:46:55 06:47:17 06:47:43 06:54:33
265:02:52:12 06:40:27 06:41:44 06:43:47 06:45:46 06:46: ll
0403 0631 0401 0402 0401 0402 0401 0402 0401 0402 0635
240 5 I0 II0 I0 I0 i0 I0 I0 34 1
06:46:37 06:47:00 06:47:23 06:49:35 Unlock (lunar) solar panel
06:55:06
06:55:49
0402
87
5.7-10
steps 11
per degrees.
degree,
the
axis the
should result
have was
rotated only
upwards a 3-degree
approximately maximum motion, while stepping polar axis, the solar Table 5.7-7 sumSC-2 mission.
However,
which then settled back to a net commands were still being sent. motor was not able to counteract marizes all stepping commands
change of 2. 2 degrees As in the case of the the effect of spinning. transmitted during the
5.7.5
REFERENCE
Report
"AC-7 Preliminary GD/C-BNZ66-053,
Test Results, 28 September
" General 1966.
Dynamics/Convair
5. 7.
6
ACKNOWLEDGEMENT R. J. Hausauer coordinated the mechanisms subsystem section.
5.7-11