1!1 ;:;:;::::}:;::::::: · illustrations fi qure 1 - descent burn throtti ing profile 2 -...
73
.< . ' ' ;::::·: ·: · · :::::::::;:: M S C- 0 4112 ;:;:;::::}:;:::::::: . SUPPLEMENT 5 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION I i ri f i i f :I::� APOLLO 14 MISSION REPORT .·.: . ·:·:·:·:·:· -. . .. . .·.·.·.· .· SPLEMFNT 5 DESCENT PROPULSION SYSTEM FiNAL FLIGHT EVALUATION �:}�{:�:�{-"(rUSA-Tfl-1-69491) APOLLO 14 IUSSION ·:·:·:·: · :·:·:·:· B!POBT. SUPPLEMENT 5: DESCENT :::::::: : : : : : : : : : BOPULSION SYSTEſt FINAL FLIGHT EVALUATION :::::::::::::::;: ( NASA) 61 p HC $5.25 CSCL 21& 873-29813 Unclas G3/28 13464 MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER 1972 •. J ,, � ·:- i, ' f � · ,·
Transcript of 1!1 ;:;:;::::}:;::::::: · illustrations fi qure 1 - descent burn throtti ing profile 2 -...
.< . ' '
;::::·:·:··:::::::::;:: M S C- 0 4112 ;:;:;::::}:;:::::::: . SUPPLEMENT 5 1!1 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
Iirifiif:I::� APOLLO 14 MISSION REPORT .·.: .·:·:·:·:·:· -. . .... ·.·.·.·.· SUPPLEMFNT 5
DESCENT PROPULSION SYSTEM FiNAL FLIGHT EVALUATION
�:}�{:�:�{-"(rUSA-Tfl-1-69491) APOLLO 14 IUSSION ·:·:·:·:·:·:·:·:· B!POBT. SUPPLEMENT 5: DESCENT ::::::::::::::::: i'BOPULSION SYSTEft FINAL FLIGHT EVALUATION :::::::::::::::;: (NASA) 61 p HC $5.25 CSCL 21&
873-29813
Unclas G3/28 13464
MANNED SPACECRAFT CENTER HOUSTON.TEXAS SEPTEMBER 1972
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APOLLO 14 !f.I�3SIOH REPO'R�
SUPPLE!v!Ei1T 5
DESCENT PROPULSION SYSTEH
FINAL FLIGHT EVALUATION
PREPARED BY
TRW Systems
APPROVW BY
Owen G. Morris
Manager, Apollo Spacecraft Program
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l�C-04112 Supplement 5
N/.'!'IONAL AERONAUTICS Arm SPACE ADMINISTRATIO!l
MANNED SPACECRAFT CENTER
HOUSTON, TEXAS September 1972
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P ROJECT T ECHNICAL REPORT
APOLLO 1 4
LM-8 DESCENT PROPULS I ON SYS T EM
F I NAL F L I GHT E VALUATION
Follow-On Contract to NAS 9-8166
Prepare d for
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17618-H21 9-RO-OO 1t
SEPTEMBER 1971
NATI ONAL AERONAUT I CS AND SPA CE ADM I NISTRAT I ON MANNED S PACECRAFT CENTER
HOUSTON, TEXAS
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I PROJECT TECHNICAL REPORT
APOLLO 14
LM-8
DESCENT P ROPULS I ON SYSTEM F I NAL F L I GHT E VALUAT I O N
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17618-H 219-RO-OO
. Follow-On Contract to NAS 9-8166 SEPTEMBEP 1971
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P ,-epa red for NAT I O Nt�L t\ER ONAUTICS AND SPACE ADMiri!STRAl ION
MAtlNED SPACECRAFT CE NTER H OUSTON, TEXAS
Prepared by A. T. Avveni re
S. C. Wood Propulsion Systems Section
Chemical and Mechanical Systems Department
NASA MSC
C onc urre d by: �(.:tft- ·-- ·�,.A' w. R. Hammock, Manager Descent Propulsion Subsy$tem
Conc u rre d by: C .'J!-4 '"' -t .. ;.,� . w. Yodz1s, L iet/
Primary Propu ion Branch
TRW I"JPI•I ONtlll
I
TRW SYSTEM5
Approved by: �� 5.,.,./1/.._ R. J. Smith, Manager Task E-99/705-3
Approved by:�� �;non, Manager Chemical and Mechanical
Systems Department
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CONTENT S
1. PURPOS E AND S CO P E
2. SUMW\RY .
3. I NT ROOUCTI ON
4. FTP S TEADY-STATE PERF!,Pt1Ai�CE ANALYSIS
5.
-. ' .
8.
'l.
Analysis Technique
Analysis Result�
Critique of Analysis Rr�ults
Comparison with PrefliJht Performance Predicti�n
Engine Performance a� ·Le:ndar,_. Interface Conditi0ns
SIMULATION OF THROTTLED P ERFO RtiAN CE RESULTS
OVERALL PERFORMANCE
POGS EVALUATION AND PRGrELLANT LOADING
Propellant Quantity Gaainq System
Pronellant Loading
PRESSURIZATION SYSTEM EVALUATION
ENGINE TRANSIENT ANALYSIS
Start and Shutdown Tran;i�nts Throttle Response
1�. REFERENC ES.
1. LM-6 DPS ENGIN E AN D FEED SYS TEM PHYSICAL C�APACTERISTICS .
FLIGHT DAiA USED W FTP :,7(,;(JY- :T ATE ANALYSiS . . .
3. DESCEN T PROP ULSION SYSTEM STE ADY-ST ATE FTP PERFORMANCE
4. DESCENT PROPULS ION SYSTEM THROTTLE D P ERFORMANCE .
5. OPS PROPELLANT QUANTITY GAGING SYSTEM PE RFORM ANCE
6. OPS START AN D SHUTDOWN IMPULSE SU,.,..ARY
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10 12 14 15 16 17 18 18 19 20
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ILLUSTRATIONS
Fi qure
1 - DESCENT BURN THROTTI ING PROFILE
2 - COMPARISON OF PREFLIGHT PREDICTED AND INFLIGHT THROAT EROSION • •
) - ACCELERATION MATCH
.1 - OXIDIZER INTERFACE PRE .liFE MATCH
S - FUEL INTERFACE PRESSLIPF. MAiCH .
5- PROPELLANT QUANTITY G,.rJNC. SYSTEM MATCH, OXIDIZC� TANK NO. 1 • • •
7- PROPELLANT QUMTITY (.�fliNG SYSIEM MATCH, OXIDILL,' �ANK NO. 2 . • •
R- PROPELLANT QUANTITY l,�t.ING SYSTEM MATCH, FUEL TAil< NO.
9 - PROPELLANT QUANTITY GA!lltlG SYSTEM MATCH, FUEL TA:W NO. 2
lC - CHAMBER PRESSURE MATCH • • •
1 1 - COMPARISON OF PREFLIGHT PREDT C.TION AND IN FLIGHT PERFORMANCE
12 - COMrARISON OF PRE FL IGHT PREOIC�ED A�D ANALYSIS PROGRAM SIMULATED THROTTLE COMMAND VOLTAGE (THROTTLING REGION).
1 ". - COMPARISON OF PREFLir.�·T PREDICID AND ANALYSIS PROGRAM SIMULATED MIXTURE RATIO (THROTTL!NG REGION) . •
1 � - COMPARISON OF PREFLIGHT PREDICTED AND AN�LY5IS PROGRAM S IMIJLATEO ENGINE SPECIFIC IMP I ILSE (THi\Oi�LING REGION)
15 - MEASURED RtGULAIOR Oli7L ET P�!:SSURE (JQ30Lr;. 16 - MEASURED REGULATOR OUTLET PRESSIJPE (GQ3025P).
17 - MEASURED AUTOMATIC C OM�'AND VOL i AGE
18 - MEASURED OXIOI�ER INTERFACE PRESSURE
19 - MEASURED FUEL INTERFACE PRESSURE
20 - MEASURED CH�MBER PRESSURE • •
21 - MEASURED PROPELLANT Q UANTITY, OXIDIZER TANK NO.
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ILLU STRATIONS (Continued)
Figure
22- MEASURED PROPELLANT QUANTITY, OXIDIZER TANK NO.
23 - MEASURED PROPELLANT QUAtlTITY, FUEL TANK NO.
24 - MEASURED PROPELLANT QUA NTITY, FUEL TANK NO. 2
25 - MEASURED INJECTOR ACTUATOR POSITION . . . .
26 - MEASURED SUPERCRITICAL HELIUM SUPPLY PRESSURE
27 - SHe TANK PRESSURE (POWERED DESCENT)
28 - DESCENT s··�.GE S't STEM SCHEMATIC • .
Pag()
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1 . �liRPOSE NJD SCOPE
The purpose of this report is to present the results of the postflight
analysis of the Descent Propulsion System ( DPS ) performance during the
Apollo 14 Mis sion. The primary objective oft�_ analysis was to det�nnine
the steady-sta �e perfonnance of the �;.;) during the descent phase c.Jf �he
m an ned lun ar landing.
This report is a supplement to the Apollo 14 Mission Report. In ad-
dition to further analysis of the DPS, this report bri:1g:; toyether infonna-
tion from other reports and memorandums an alyzing specific an omalies an d
performance in order t o present a c omprehensive description of the DPS
operation during the Apollo 14 Mission.
The following items are the major additions a�d changes to the results
as reported in Reference 1.
(1) The performance values for the DPS burn are presented.
(2) The analysis techniques, prob lems and assumptions are discussed.
(3) The analysis results are compared to the preflight perfonnan ce
prediction.
(4) The Propellant Quantity Gagin g System is discussed in greater
detail.
(5) Engine transient perfonnance an d throttle respnnse arc: discussed.
(6) Estimated propellant con sumption and residuals are revised.
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2. SUMMARY
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The pe�forn�nce of the LM-8 Descent Propulsion System during the
Apollo 14 Mission was evaluated and found to be satisfactory . The aver-
age engine effective specific impulse was 0.1 second higher than pre
dicted, but well within the predicted lu uncertainty. The enqine per
for�nce corrected to standard inlet conditions for the FTP portior of
the burn at 43 seconds after ignition was as follows: thrust, 9802 lbf;
5pecific impulse, 30 4.1 sec; and propellant mixture ratio, 1.603. These
values are +0.8, -0.06, and +0.3 percent different, respectively, from
the values reported from engine acceptance tests and were within speci-
fication limits .
Several flight measurement discrepancies existed during the flight .
1) The chamber pressure transducer had a large drift, exhibiting a maxi
mum error of about 5 psi at appro xi mate ly 150 sec after engine ignition . a This problem has occurred during other flights but never to t.his magni
tude. (Previous errors were less than 1 psi.) 2) The fuel and oxidizer
interface pressure measurements appeared to be low during the entir'� flight.
The discrepancy, which does not inc lude the regulator outlet measurement
bias, is assumed to be a measurement bias (-.92 and -2.44 ps i for oxidizer
and fuel, respec tively) . 3) The propell ant quantity gaging system per
formance was not within s pecification. Testing at the White Sands facil
ity prior to the flight showed that the des ign could not meet the s pec
ification performance; consequently , this out-of-specification perform-
ance was expected.
The l ow l evel sensor actuation time was as predicted, indicating
that the new anti-slos h baffles installed for this flight performed well.
2
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3. INTRODUCTION
The Apollo 14 Mi ssi on was th e seventh flight and si xth manned flight
of the Lunar Module (LM). The mi ssi on was the thi rd successful lunar
landing.
The ;,pace vehi cle was launched from Kennedy Space Center (KSC) at
4:03:02 p. m. (EST} on January 31, 1971. The la�nch was delaye d about
40 mi nutes due to unsati sfactory weather condi ti ons. At 108:02:27 hours,
the Descent Burn (POI) was i ni ti ated and lasted about 765 sec . The burn
was started at the mi ni mum throttle setti ng and after approximately 27 sec,
the thrust was manually i ncreased to the fixed throttle posi tion (FTP).
The targeted time for throttle up was 26 sec . However, due to a faulty
abort signal recorded prior to the POI burn, a lll(l.nua 1 rather than an
automatic thi'Ottle up was requi red. An automati c descent was maintained
to approximately 643 seconds after throttle recovery, at whi ch ti me the
astronauts ass umed s emi- manual control of the final landing phase. The
engine was commanded through a substantial number of throttle changes by
the LM commander during this final landing phase. Lunar landi ng occurred
at 1 0d:l5:09 AET ending the DPS mi ssion duty cycle. Successful vent i ng
of the descent propellant tanks occurred about 1 min. after touchdown at
about 108 hr. 16 min. AET . After a lunar stay of approxi �tely 33-3/4
hours, the APS was ignited and the ascent stage of the LM li fted off and
was inserted into lunar orbit about 1 2 -l/2 minut�s later. Descent data
te, �inated at ascent liftoff.
The actual ignition a nd shutdown times for the DPS firing are
1 08:02:26 hours and 108:15:10.8 hours , res pectively. The throttli ng
profile for the DPS burn is shown in Figure 1 .
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The Apol lo 14 Mission uti lized LM-8 which was equipped with DPS
eng�ne S/N 1032. The engine and feed system characteristics a��
sented 1n Tab le 1 .
The DPS burn was preceded by a two jet +X Reaction Contro l System
(RCS) ullage maneuver of 7 seconds to sett le prope llants.
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4. rTP STEADY-STATE PERFORMANCE
�- Analysis Technique
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The mdjor a n a l ys is effurt for this report was concentrated on deten�in
�ng the f l ight stea dy-stat e performance of the DPS dur i ng the fixed throttlr
position \FTP) pr..rtion of the i)escent Bur·n. A recr1,struction of the
throttl ed portion of t::r- Descent Bur n was attempted, however, due to thL
rJpid changes in the eng1ne tr,rust ,- �ten ex;H�ri ence . r1 ... ring thi, [10rt,C"Jn
of the burn, a detai led anfllysis was not possi b l e . Th� performanct analysL
nf the FTP region was uccomp1ished by use of the Arol lo Propuls1on Pnalys1s
Program whi ch utilizes a m1ni mum vari ance techni q u e tu "best" correlate
the available flight data. The program embodi es error moaels for the
v ar i ous f l i ght data that are used as inputs, and by 1 t erati ve �thods,
arrives at est i mates of the system performance n istory and prope.llant
weights which "best" (m i nim um - v ariance s ense) recon c i le the data .
The reconstruction of the throttled portion was made us ir.g a
s im u lat i on tech n iq ue a nd h an d a dj ust i ng var i ous initial parameters to
ach ieve a reas onable fit to the data.
Analv !l is Res u l ts
The engine performance diJring the FTP portion of the Descent Burn
was s atisfactory.
The Apollo P rop u ls i on A na l ys is Pr ogram (PAP) r es u lts presented in
this report are based on r ecvnstr uttl ons us ing data from the fli ght
measurements listed in Ta bl e 3.
The propellant d�nsities were calculated from samp l e s peci fic gravity
data from �SC, assumed interface tem peratures based o n the f i i ght bu l �
propellant temperatures, and the flight interface press ures.
5
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The initial vehicle weight was obtainP.d from Reference 2. Th�
initial estimates of the propellant onboard at the be4inning of the
analyzed time segment were caltulated from the loa1ed propellant weights.
rke damp weight was also adjusted for t0nsumables such as RCS propel lant,
water, etc., used betw�en �qnition and the start of the analyzed time
segment. During the Descert Burn approximately 120 lbm of consumables
other than the DPS propell�nt were used. Of the amount, 90. 1 l bm was
RCS propellant. Since there was little RCS act�vity during the �nalyzed
portion of tne burn, it .,...,s assumed that the non-DPS consumed weight was
used at a rate of .08 lbm/sec.
The DPS steady-state FTP performance was determined from the analysis
of a 340 second segment of lhi hurn. The segn�nt of the burn analyzed
commenced approx1ma tely 33 seconds after DPS ignition (FS-1) and include1
the flight time between 108:02:59 hours and 108:08:3� hours gro�nd elapsed
time. Engi ne throttledown to 60 perc· nt occ urred 8 seconds after the end
point of the analyzed segment.
The results of the Propulsion Analysis Program reconstruct;on of the
FTP portion of the Descent 8�Jrn are �resented in Table 3 along !�ith the
preflight values. The va�ues ore�enteo are end point conditions of the
segment dnalyzed and are considered representat;ve of the actual flight
v alues throughout the segment. In general, the actual values are within
1.0 percent of the predicted values. A portion of the d;fference can be
attribute� to the differerse bPtw!�n predicted �nd actual regulator outlet
pressure ( l psi a at the e·.� of t<-- FT0 burn ) . The infl ight throat ero�if',, agreed wel: witn predicted values. At tlie
t>nd of the HP portion of the burn, the inflight tiwoat erosion was within
1'l of the pred;cted value of 9.07' • .
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Critique of Anal ysis Resul ts
Figures 3 through 10 s how the anal ysis program output p l ots which
pres ent t�e fil tered fl ight data and the accuracy with which the data was
matched by the PAP program. The accuracy is represented by the residual ,
which is defined as the difference between the fil ter�d data and the
program cal cul ated val ue. The figures pre3ent�d are thrust accel tration,
oxidizer interface pressure , fuel interface pressure, quantity gauinq
system for oxidizer tank 1 and 2, quantity gaging system for fuel tank
and 2, and chamber pressure. The chamber pressure p l ot indicated how ba dl y
the chamber pressure measurement behaved during the burn. Because of
this , chamber press�re was not used in the PAP program as a mea�urement .
A st rong ind�cation of the val idity of the anal ysis program simu1ation
can be obtained by comparing tne thrust accel eration history as determined
from the LM Guidan ce Computer (LGC) �V cata to that computed in the simul a
tion. Figure 3 shows the thrust accel eration derived from the 6V data
and the res idual between t�e meas ured and the computed val ues. The
time history of the residual has an essential ly zero mean and a smal l
neqative sl ope.
Several probl ems were encountered with fl ight data whil e anal yzing
the s teady-s tate performance at FTP. Several assumptions were necessary
in order to obtain an acceptabl e match to the fl ight data. These probl ems
are discussed below.
The regulator outlet p ress ure is redundantl y samp l ed by measurements
GQ 3018P and GQ 3025P. The pressure indicated by GQ 3025P was about 2 psi
lower than that from GQ 3018P. Bas ed on earlier anal yses and prefl ight tests ,
t he dat a from GQ 3018P was used for the analys is . It s hould be noted that
tests made at KSC s everal weeks prior to launch on the hel ··•m regulator
7
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; ndi cated about a 1 psi diffe renee between the two measurements. Figure 1 1
shows the helium regulator pressure used to drive the program. Its shape
was dictated by the interface pressure data which, although they were
biased , agreed well with each other as to shape. Analysis indicated that
the helium regulator measurement GQ 301 8P was biased by about 1 psi.
The inflight chamber pressure (Figure 10) could not be used due to a
large drift as previously stated.
The filtered value of the fuel interface pressure (GQ 411 1P) was
severely biased by -2. 1 psi, although this is within the instrument accuracy.
The oxidizer interface pressure was also biased by - .9 psi.
During the throttled portion of the bum , intermittent data dropout
occurred from about 530 seconds to 580 seconds after ignition. As previously
stated , the throttle region is driven by the actual commanded thrust (GH1331 V) .
Since no attempt was made to smooth the data during the throttled portion
)
of the burn, other than to eliminate " wild" points, a poor acceleration 0
1 ··:
match during this time was obtained , with the calculated values higher than
the measured values.
Comparison with Preflight Performance Predictions
Prior to the Apol lo 14 Mission the expected inflight perfcnmance of
the DPS was presented in Reference 3 . The preflight performance rep ort was
intended to bring together all the information relating t o the entire
Descent Propulsion System and to present the results of the simulation of
its operation in the space environment.
Tne predicted steady-state and related three sigma dispersions for
the specific impulse , mixture ratio and thrust during the FTP portion of
the Desce nt Burn are presented in F igure 1 1 .
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Engine Per formance at Standard Inlet Conditions
The flight perfonnance prediction of the DPS engine was based on the
data obtained irom the engine acceptance tests. In order to provide a
common basis for comparing engine performance, the acceptance test and flight
performance is adjusted to standard inlet conditions. This allows actual
engine performance variations to be separated from pressurization system
and propellant temperature induced variations. The standard inlet condi
tions performance values were calculates for the following conditio11s:
Standard Inlet Conditions
Oxidizer interfar.e pressure, psia Fuel interface pressure, psia Oxidizer interface temperature, °F Fuel interface temperature, °F Thrust acceleration, lbf/lbm Throat area, in2
2 22.0 2 22.0
70.0 70 . 0
1 .0 54.4
The following table presents ground test data and flight test data
adjusted to standard inlet conditions. Comparing the corrected engine flight
performance at FTP durinq the Descent Burn to t' corrected ground test
� I
Ground Test FliQht Engine Prediction Analysis Characterization Results e
Thrust, lbf 9732 9802 I Specific Impulse, sec 304.3 304.1 I Mixture Ratio 1 .598 1 .603
data shows the flight data to be 0.8% more, 0.06% less, and 0.3% more for
th.-ust, specifi•: impulse and mixture ratio, respectively. These differences
are within the engine repeatability uncertainties and within the performa"ce
specification ranges.
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5. Slftt.lLATIO N OF THI()lTLED PERFORMANCE RESULTS
The DPS throttling perfonnance was simulated by utilizing the prediction
mode of the Apollo Propulsion Analysis Program. By this method , the measured
value of the regulator outlet pressure (GR 301 8P) drives the program and the
measure d value of throttle command voltage (GH1 331 V) determines the engine
throttle setting. The program then calculates values of the remaining
flig ht measurements and engine performance. In this mode the program does
not compare calculated measurements with flight measurements and a minimum
variance match is not performed.
Based on the FTP analysis , it was deterreined that a -1 psia correction
should be made to the regulator outlet pressure (GQ301 8P). For the simulation ,
the initial values of throat erosion LM vehicle and propellant weights
were obtained from the end point conditions of the FTP analysis. The damp
weight was adjusted for non-DPS consumables, as in the FTP analysis , at a
rate of 0. 22 lbm/sec to account for the remainder of that weight lost
during the bum.
The DPS throttling performance simulation was conducted starting at the
end of the FTP analysis (FS-1 + 374 seconds) and continued for 388 seconds.
This includes all of the powered descent burn after throttle down and
includes the flight time between 1 08:08:48 hours to 108:1 5:09 hours.
Typical values of the simulation results are presented in Table 5.
Since throttle recovery occurred 1 0 seconds earlier than predicted due to a
high thrust at FTP, a cof11)arison between the original pref1ight and flight
data is not realistic. This is due to trajectory differences which would
force the guidance sys tern to conwnan d difference thrust levels than predicted.
1 0
I I I
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.,., ......
•
0
•
. .. ._,, .
� - . •
�-� \-. � '
" .. , ' t(.".
; �
However, another "preflight" was made assuming the high thrust level at
FTP. This al lows for s�mil ar commanded thrusts durin g the throttl e region, •
and so �hese preflight val ues appear in Table 5 for comparison with the
fl ight data.
Figures 12 through 14 present pl ots comparing the modified prefl ight
predicted and the analysis program simulated values of throttl e co11111and
percent , mixture ratio , and specific impulse. Also shown is the
original preflight. T he difference in t he profil es between the original
preflight and the analysis simul ation is due primarily to the trajectory
differences caused by the high thrust at FTP.
Figures 15 through 26 present the inflight val ues of the measured
propulsion parameters. The major portion of the FTP data has been del eted
to obtain better resolution. In general, the FTP data shown is representa-
tive of the deleted segment .
11
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... - :·
.� . .
, . .
• J
"• li
'. l• ·'
6. OVERALL PERFORMANCE
•
oil>.
When the results of the FTP analysis and the simulation of throttled
operation are combined, the overall performance during the Descent Burn and
the total propellant consumption for the mission can be evaluated. The
following table presents a comparision of the propellant consumption, average
mixture ratio (MR) and overall effective specific impulse (ISP}. The vehicle
effective specific in1pulse was computed based on spacecraft weight reduction
due to both DPS propellant consumption and non-DPS consumables (approximately
0. 08 lbm/sec during FTP and 0.22 lbm/sec during throttled operation}. The
engine effective specific impulse was calculated considering only weight
reduction due to DPS propellant usage. Contributions from RCS activity are
not included.
Preflight Prediction
Analysis Program
Gaging System*
Propellant -�-<msJJ�ti '!n_U�� Oxidizer Fuel
10758 6729
10809 6794
10742 6731
*Based on gaging system and AV measurements
Average MR
(0/F)
1.601
1. 591
1. 596
Vehicle Effective 1
Engine 1 Effective1
Jsp(sec) I sp(sec) 300.4 302.5
299.8 302.6
300.4 I 303 .2
The mea�ured propellant quantities consumed are based on final gage
read i ngs and measured initial loads. Due to l oad i ng and gaging system in
accuracies, the uncertainties in the consumed propellants are :±.85 lbm and
153 lbm (3o} for oxidizer and fuel, respectively. The uncertainties in
mixtur� ratio and effective specific impulse resulting fran these uncer
tainties are �0.016 and ±1.59 respectively. Both the predicted and analysis
program results are within these uncertainties.
1 Calculated from FS-1 plus 33 seconds.
1 2
)
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rt· ��· lf4.:./. , �:f.t�� .
.. ' .Y.·
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The values of effective specific impulse presented in the tabl e are
dependent on both the vehicl e weight change and the thrust vel ocity fJ<�'�"�
The analysis indicated a thrust velocity gain of 7017 ft/sec. The
vel ocity gain used in computing the val ues of Isp usinq the g aging c,yr.ter,,
readings were taken directl y from the acceleration data (GGOOOl Y) wr11Lh
measures the g ain in velocity for each two-second segment of the burn.
The total me asured thrust •Jel ocity gain , 6997 ft/sec, includes the con-
tribution of both the DPS engine and RCS activity . T he simuiation was in
error by 20 ft/sec due to the previously mentioned data dropout in the
commanded thrust data (G41331V). The propellant consumed as cal cul ated in the
simul ation, was corrected to account for the error induced by the data dropout.
The uncertainty in effective specific impulse due to measured propel lant
usag e and velocity gain uncertainties is tl.2 seconds. The engine effective
specific impulse for both the prediction and analysis are within this un
certainty. Due to the rather large uncertainties rel ated to the gag ing
system, it is felt that the best estimate of DPS performance is given by the
dnalysis results, rather than from gaging system data above.
Both the anal ysis results and measured results are within the predic
t ion 3· uncertainties of ±5.91 sec and ±0.0225 for effective specific
impulse and mixture ratio, respectively. The difference bet ween the pre
dicted vehicle effective specific impulse and that calcul ated from the
PQGS measurements and the analysis program was due to more RCS usag e
(non-DPS consumabl e) than predicted .
13
·-
L .
--
1 . . ..... ....
·' ·�
•
, .
7. PQGS EVAULATION AND PROPELLANT LOADWG
At ignition cf the descent engine, all propellant gages should be
reading off scale. However, the fuel tank 2 (FU2) gage never did read off
scale, while the FUl gage intermittently droppad slightly belm·J the ma>'.imurn
reading of 95%. The oxidizer gages initially performed as expected.
All gages were indicating propellant consumption by about 50 seconds
after ignition. At that time the gages were reading 93.5%, 95.0%, 91.5%,
and 90.5� for OXl, OX2, FUl and FU2, respectively. The reconstruction
analysis indicates that the fuel gages were out of specification limits
during the first 140 seconds of the ana)ysis but were within the expected
range as d etermined by the WSTF tests. Therefore, thry were not included
in the analysis until 140 seconds when they were respectively 0.6% and -O.BY
for FUl and FU2 between the calculated and measured values. The oxidizer
difference wa� +0.4� and +0.4% for OXl �nd OX2, respect�vely. Throughout
the burn, a difference of from 1 to 3� exists in the l�Quid level measure-
ment bctiiC'cn the two oxidizer tanks. The difference in propellant levels
was due to propellant t,·ansfer bet\leen the oxidizer tad;s caused by s;mbl1l in�
of the engine com111anding the thrust vector to pass thr:n .. .:.:h the vehicle center
of gravity. At the end of the FTP portion, a predict.d difference of 4.0�;
versus a 111easured diff erenr.e of 3. 4�; \-Ja� obse1·ved. At touchdO\·m, agreement
between the average reading of both oxidizer and both fuel tanks were within
0.1�: for oxidizer and 0.3' for the fuel. The oxidizer t;:.ages indicated
a diffet�:ltll' of i.7: beti-IL'Cn the 'l\':o tanks a� compared to a ptedictE:d value
of 37�. At the end of the burn the propE:ll<:mt gages wert:� reading approxintate1y
6.1�:, 3.3�, 5. 1', und 3.4: fo1· OXl �X2, FUl and FU2, rc!pective'ly.
The expected accul'acics for the gaging system, basc.d on tests conducted
at WSTf ( Reference 4) are presented in the fo1lm-Jing taLlc:
- ·-- - · · · · - -· -· "· · - ·�
14
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)
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,., -
a p¢ -
-�""'·;; ... �.. £ ... . <
.... ·t� • .. -'.·'
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.. _, � · .. �. '
,, • ;
i
0
•
EXPECTED PROPELLANT GAGING SYSTEM ACCURACY
Quantity Remaining in Tank
100-50%
50-25%
25-8 %
8-0 %
* Percent of Full Tank
Accuracy For Each Oxi dizer
Gage*
2.7%
1.0%
0.5%
1.0%
Quanti ty Remai ni ng i n Tank
1 00-60�G
60-20%
20-0 %
. ,.��. '
Accuracy Far Each Fue 1
Gage*
3.5%
2.0%
l.m:,
The specification limit of the PQGS i s ±U of full tank capaci t y for
quantiti es above tJ� load and below 8% load. When the PQGS i s integrated
i nto the vehi cle and telemetry effects are consi dered , the 1% value i s
increased to 1.3%. In the 8% to 25% range, the speci ficat ion requi rement
i s ±0.5% of full tank capaci ty. However, the WSTF tests i ndicate that
these specificati ons cannot be met.
In the analysis of FTP, propellant was transferred from OX2 to O Xl
at a rate o f 0.67 lbm/sec with an initial unbalance of 11.6 lbm. No fuel
transfer was simulated. Table 5 presents a comparison of the measured
data and the best estimate of the actual values at various time points
during the descent burn. Whi le the di fference be tween the measured and
computed values were frequently outside the specification limits, they were
generally within the expected accurac y of the gaging probe based on WSTF
results. At engine shutdown, the quantities of propellants remaining in the
tanks were computed to be 5.6l, 2.6%, 3.4%, and 3.4% for OXl, OX2, F Ul and
FU2, respectiv ely. This is equivalent to remaining tanked quantities of
477 lbm of oxidizer and 244 lbm of fuel. Of these quantities, 374 lbm of
oxidizer and 221 l bm of f uel are usa ble to depletion (including burning usable
15
- ----.......... --·----- "·--·- - · - --- ·..._. - "" - ;�. "!' . _,, ..... .. . __ ,_....,., , .. ,.
.. � ..
··- . ... _ ............ -
•
. " . •• .1 • .... ---... ' .....
propella nt s in feedlines ) . Using the co mputed propellant flowrates and
the computed propellant resid uals at engine shtttdnwn,60 second s of hover
time remained to fuel dcpl�tion. The measured propellant quantities re-
maining at engine shutdnl'm , 439 lbm of usable oxirlizer and 294 lbm of usable
fuel, ind icated a hover time of 77 seconds to oxidizer depletion .
The propell ant low l<'vel sens or \·Jas trig<yred at 711 se cond s intu the
burn . Gaging system data in dicates that the sensor was triggered oy the
OX2 probe. At the time of s ign al , the measured re adings were 8.4%, 5.5;�,
7.5%, and 5.6�{ for OXl, OX2, Fl.:l and FU2 gages, respectively. Bdsed on pre-·
dieted times, the low level sen�or v1as triggered on time . This indicated
the new anti-slosh baffles in stal le d in the des cent propellant tanks operated
as expected, and premature triggering of the 10\'J 1evel s ensor was prevented
(was approximdtely 20 seconds prt.'llli:lture on Apollo 12).
\ /
Pri or to propPllar;t lodtling, ck·rasity detC'nnin'-ltions 1·1ere made for eoch 0
propellant to pc:;tatdish th!.! ilrnounl of off-loadin�l of the planned overfill.
An average oxi d izer densi\y of 90.33 lbm/ft3 dnd an average fuel de nsit y
of 56. 52 1 bm/ft3 at a prc � s�a re of 240 psi a and � temperature of 707� 1·1ere detenninPd froil: the Sa!1!plP'>. The t<JnLfd propc:llr.nts \·Jere 7037.7 lbrn of fuel
and 11283.9 ltJ"' of oxidizu. The toiul tanl--ed propel1ant� 1vere 18321.6
±20 lbrn.
16
.... ':::"'�:- . . . "?
. '�
+ -
()
• �l '
•
..
8. PRESSURIZATION SYSTEM EVALUATION
:-� .... . • . , .
• "...; r )ji.._
The DPS Supercr1tical Helium (SHe) system performed essentially as
ant i ci pated duri ng the Apollo 14 Mission . The predi cted and fli ght
measured values for the p ressure ri se rates are shown below for compari son.
Prefli ght Apollo 14 Parameter Predi cti on Mi ssi on
On-Ground Pressure Ri se Rate, psi / hr 8.15 8 . 0
Coast Pressure Ri se Rate, psi /h r 6.65 6.2
Lunar Stay Pressure Ri se Rate, psi /hr 4.9* 6 . 0
*Apollo 12 Mi ssi on value duri ng a 30-hour lunar stay.
Both the lunar ri se rate and ground-to- coast sh i ft i n ri se rate experi enced
duri ng the Apollo 1 4 Mi ssi on were h i gher than p redi cted. The ground- to-coast
shi ft in ri se rate of 1.8 psi/hr . although hi gher than predi cted, i s close
to the 1.5 psi / hr value. The lunar ri se rate di fference can be explai ned
wi th the following i nfonmati or.. Exi sti ng SHe system bottles are ei ther
doubly- evacuated or si ngly-evacuated types. The pressure ri se rates of SHe
bottles in which the annulus has been si ngle-evacuated arP. h i gher than the
doubly-evacu ated annulus bottles and cons i derably hi gher when the bottles
are less than full. Because Apollo 14 was t he fi rst successful fli ght that
contained a single-evacuated type of SHe bott le and the preflight lunar rise
rate shown above was obtained from Apollo 12 flight data for a doubly- evacuated
bottle, the higher value of lunar rise rate for Apol lo 14 appears reasonable.
An Apollo 14 postflight simulation of the SHe system generated with
the latest version of the SHe system computer program is presented i n
Figure 27. The comparison with flight data shows that a close match to
the S He bottle pressure during the DPS engine burn w4s obtai�ed.
,_ .
17
� .
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t t ' ,
at m- z:a . .
•
9. ENGINE TRANSIENT ANALYSIS
The mission duty cycle of the OesCP11t Propulsion System for Apollo 14 included one start at the minimum throttle setting, and one shutdown at
approximately 29% throttle. Much throttling occurred during the Descent
�urn, all of whi c.h was commanded by the LGC.
Start and Shutdown Transients
Reference 5 presents the tect:,,ique used in detennining the time of
engine fire switch signals (F�- 1 and FS-2) for the Descent Burn. This
method was developed from White Sands Test Facility (WSTF) test data and
d�sume� that approximately 0.030 seconds after the engine start command
(FS- 1) an oscillation in the fuel interface pressure occurs, as ob�erved
from tne WSTF tests. Similarly, 0.092 seconds after the engine shutdown
signal (FS-2) another oscilldtion 1n the fuel interface pressure occurs.
Thu�. start and shutdown oscillations of the fue·l interface pressure were
noted arad the appropriate time lead applied.
The ignition delay from FS- 1 to first rise in chamber pressure was
approximate ly 0. 55 seconds. The delay time compared favorat�y with the first
bum delay observed dur·ing Apollo 13. Comparir g the delay time for the POI
burn of Apollo 13 with other POI burn de lay times can not be dor.e since a l l
other POI bums were second burns; that is, they all followed .t 001 burn.
The start transient from FS-1 to 90% of the minimum steadywstate
throttle setting required 2.14 seconds with a start imp� lse of 710 lbf-sec.
The transient time was wel l within the specification limit of 4.0 seconds
for a minim um throttle start. The start transient from 90% to 100% of the
mirlimum throttle setting required 0.1 3 seconds with an impulse of 160 lbf·sec.
18
- . - , . ......,. _..... . .. -- '- --� f - ·- '
.. •--r·
)
0
.,,
. - •• "'J.· '-<�; - .... -
• ,'11:;. ·"'-, _'{(< •
.... -
•
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.... .. ' tl
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-- --- ........ - '
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The shutdown t rans i ent requi red 1 . 07 sec onds from FS- 2 to 10% of the
steady-s tate thrott l e se�ti ng with an i mpul se of 976 l bf-sec . The s pecifica
tion li mi t on trans i e nt shut down t i me i s 0.25 seconds, howe ve r , thi s appl i es
o nl y to shutdowns from FTP . The re i s no speci fi cat� on l imit on impul se.
Throttle Response
Uuri ng the Descent Burn the engi ne was commanded to many di ffe rent
t h rust l e ve ls . Al l t'lrottl e commands we re automatic. The first throttl ing
manett ver , mi nimum (14% of ful l thrust ) to FTP, which was Pxecute d 27 seconds
into the burn , require d approxi ma tely 1 second. The engi1 1e the n re mained
at FTP for 352 seconds . The seco nd comman� . from FTP to 59% , occurred 380
seco nds afte r i gn i t i on and required appro ximatel y 0.5 sec ond. This val ue
r:ompa res favorabl y with simi� _. · ma neuve rs on previous fl ights. Litt l e
throttl ing was pe rformed during the ne xt 1 45 sec onds. The L M Gui dance
Compute r the n conmande d a ramping dec rease in the t; t rottl e setti ng from 60%
to 33% over 118 seconds. At this time the Spacecraft Co�� nder se l ected
guidance p rogr am P-66 wh i ch allowed hi m to �elect the vehi c le rate of
descent wi th the LGC s till controll ing the Descent Engine . During the sub
seque nt 123 seconds o f the burn , the L GC commanded approximately 60 throttle
changes i n the 28% to 33% r ange . The c ommand time from one throttle setting
to the next was gener a l ly less than 0. 30 sec on�s. The requi rement for the
larqe number of th rottle ch ange s w as di rectly at tri buted to the spacecr a ft
attitude . As the as tronaut pi tched or rolled the vehicle , a diffe rent
engi ne th,.ottle setti ng was necess ary to ma i nt a i n the select.,d rate of desce nt. i
Whi le no th rottl � res ponse sp�cifications exi st fo r commands of the type
given duri ng the l atte r po rt i on o f the burn , the res ponse of the DPS engi ne
wJs c ons i de red s at i s facto ry .
19
_. _...,.. ., _ . -. ··-- . . .
--
• i
1.
2 .
3.
4 .
5 .
-'· -, : \'•
._,,. � ...,.\').
·,. ;
- f ,._ -
10. REFERENCES
•
. , ' ..... �
NASA Report MSC-01 855 , "Apo1 1o 14 Mi ssion Report, " May 1971.
SNA- A- D-02 7 (1 1 1 ) . Rev . 3., " CSM/ LM Spacec raft O perational Data Book, " Vo l . I-II , Mass Pro perties, 31 Ap ril 1971.
TRW Rep ort, " Apol lo Mi 5 S i on H 3/ LM- 8/DPS Prefl ight Pe rfonnan ce Report , " P . T. Avve n i re, Novembe r 1970.
GAC LED- 271 -98, " PQGS Accuracy Study , " I. Gl ick and S. Newman, 1 4 May 1969.
MSC �roo randum EP22-41-69, " Transient Anal ysis of Apo l l o 9 LMOE, " from EP2/ Sys terns Anal ysis Section to EP2/Chi ef. Primary Propul sion B ranch, 5 May 1969.
20
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t tiL
L . . .
0
· . . ..
• ' , , .
.. _,... . . - - .
•
TABLE 1
LM- 8 DESCENT P ROPU L S I ON ENG I NE AND
F E ED S Y ST EM PHY S I CAL C HARACTE K I ST I C�
ENG I NE
E ng i ne Number
C h amber T hroat Area , i n2
Nozz l e E x i t A rea , i n2
Nozz l e Expans i on Rati o
FEED SYSTEM
Ox i d i zer P rope l l an t T a n k s , To t a l
Amb i ent4 Vol ume , F t 3
F u e l Prope l l an t Tanks , Tot a l
Ambi ent Vol ume , F t 3
Ox i d i zer Tank to I n te rface l bf - sec2 Res i s tance , l bm- ft5
Fue l Tank to I nterf ace 1 bf-sec2 Res i s t ance , _.;..;__,;;..;;..::---1 bm- f t5
1 032
54 . 1 97 1
2569 . 7 3
47 . 4 3
1 26 . 03
1 26 . 03
431 . 2 1 5 2
684 . 787 2
1 TRW No . 0 1 82 7 - 6 260- R-00 , TRW L EM Des ce n t Engi n e S e r i a l No . 1 03 2 Acceptance Tes t Pe rformance Report P a rag raph 6 . 1 0 , dated 7 October 1 969 .
2GAEC Co l d F l ow Test5 .
3Approx i mate Val ues .
2 1
t I
... .
.. ..
' ' '
' � .
.. !,
� ' ' .
,• '�-
.... •,\; J " .. •.J•
.. . .. � .... , ' ' ... ! ./
Mea s u remen t NumbP r
GQ301 8P
GQ361 1 P
GQ41 1 1 P
GQ3603Q
GQ3604Q
GQ41 03Q
GQ41 04Q
GQ37 1 8T
GQ37 1 9T
GQ421 8T
GQ42 1 9T
GQ4001 X
.. . , ,_ , . , .... , · -
.. .
.. ... �� . ..... .t ";
. . . • \ ' I
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TABLE 2
F L I GHT DATA USED I N FTP STEADY- STATE ANALY S I S
Samp l e Rate Des cri Et i on Range S amEl eLSec
Pres s u re , Hel i um Reg . Ou t . Man i fol d 0- 300 ps i a 1
Pre s s u re , Eng i ne Fuel I n terf a ce 0- 300 ps i a 200
Pr·e s s u re , E n g i ne Oxi d i zer I nterface 0-300 p s i a 200
Qua n t i ty , Fuel T d n k No . 1 0-95 pe rcen t 1 00
Qu an t i ty . Fuel Tank No . 2 0-95 percent 1 00
Qu ant i ty , Ox i d i zer Ta n k No . 1 J-95 pe rcent 1 00
Quan t i ty , Ox i d i zer Tank No . 2 0-95 percent 1 00
Temp e ratu re , Fue l Bul k T a n k No . 20- 1 20° F
Tempe ratu re , Fuel Bu l k Tank No . 2 20- 1 20° F
Temperature , Ox i d i zer Bu l k Tank No . 20- 1 20° F
Temperatu re , Ox i d i zer Bu l k Tank No . 2 20- 1 20° F
PGNS Down l i n k Data 40 B i ts l / 2
22
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• .. !:,.' ' ���i" - "" � .� ,' � .t,'a .. �; J � :'- - j .
; � ,t· ,. � ilf,. ... .., �:ir1 · .. :f' I· �� . -t" ,1 .,.¥(. . •. , • .,� :.. i: �
-"' •, 'J' -,.!t.� �·�� . �--- - ��h�i.fj'>-h·' )�-�.).,:-·.;':·ilc .:� ... J 0 ' •
-�--;9- ' , } -��.1 �� ._ �;;� � l_"� � -������:it��-:_ : - l- � t " �· .. t, '
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, 0 0
TABL E 3
D ESCEflT PROPULS I ON SYSTEM S TEADY- STAT E FTP P E R � O RMANCE
PARAMETER
I NSTRUMENTED
Reg u l ator Out l et Pre s s u re , ps i a
Ox i d i ze r I nterface P ressure , ps i a
Fuel I n terface Pre s s u re , ps i a
Engi ne Chambe r Pre s s u re , ps i a
Ox i d i zer Bu l k Ten pe rature , Tank No . 1 , ° F Ox i d i zer Bu l k Temperatu re , Tank No . 2 , ° F
Fuel Bu l k Temperature , Tank No . 1 , °F
Fuel Bu l k Temperature , Tank No . 2, or
DERI VED
Ox i di zer F l owrate , l bm/sec
Fue l F l owrate , l bm/sec
P rope l l ant Mi x tu re �a t i o
Vacuum Spec i f i c l�pu 1 se , sec
Vacuum Thru s t , l bf
Throat E ros i on , %
--
FS- 1 + 43 S ECONDS FS - 1 + 3 7 3 SECONDS
PRE D I CTED MEASURED CALCULATED PRED ICTED MEASURED CALCULATED
244 . 8 245 . 7 244 . 5 244 . 8 246 . 9 245.7
223 . 9 223 . 1 223 . 9 222 . 8 223 . 1 223 . 9
223 . 9 221 . 6 224 . 0 223 . 0 221 . 8 224. 1
1 04 . 5 1 05 . 8 1 04 . 9 99 1 00 97. 3
68 6C - - - 6ti 68 - - -
68 68 - - - 68 68 - - -
68 68 -- - 68 68 - --j
68 1 68 - - - 68 68 - - -
1 9 . 81 - - - 1 9 . 9 20 . 2 - - - 20 . 4
1 2 . 4 - - - 1 2 . 5 12 . 7 - - - 1 2 . 8
1 . 597 - - - 1 . 603 1 . 596 - - - 1 . 60
304 . 4 - - - 304 . 2 301 . 1 - - - 301 . 2
9 7 92 - - - 986 1 9901 - - - 9990 - . 6 1 - - - - . 6 3 9 . 07 - - - 8 . 23
•»t J.At t_ t _, Jt. � � �. • "" •• ...... ........ � ........ ..... --...... - - . ..- - ._ ... . � . .. � ., ... ,...,_� - . -;.1.:;. """ " " �-- · - ·
-� .. �t:'-. . �
;.,�<:;' -'': . ..
?,' � ':..� \ • ;!'"\
I I
..
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� . .... . .... �-, . · ·� � � ��?· .. · . . . . ' , t / - '
� .f!. , \
.,
N .,.
F l? ! j .S TJJ'# ·. HJ. ¥1 . -·;· �-: : ·· ... - ·,,, . . . v 7>•' • • ·.r ....... ,,._. _ _.., __ _;__, __ , _ ___,.,�,...
PARAMETER IN�TRI.Jot[NTED
Regu l ator Outl et Press ure , psi a
Ox i d i ze r I n terface Pres s u re , ps i a
Fuel i n terface P re s s ure . ps i a
E ng i ne Chamber Press ure , psi a
Oxi di zer Bul k Temperature , Tank No . 1 , ° F
Oxi di zer Bul k Tempera ture , Tank No . 2 , °F
Fuel Bul k Temperature , Tank No . 1 , °F
Fuel Bul k Temperature , Tank No. 2 , 0f
Throttle Command Vol tage , VDC
DERI VED
Oxi d i zer Flowra te , l bm/sec
Fuel F l owrate , l bm/sec
Propel l ant Mi x ture Rati o , 0/�
Vacuum Spec i fi c Impu l s e , sec
Vacuum Thrus t , l bm
Throat E ros i on , � - - -·-···---- - -- --- -- -
'- �
TABLE 4
DESCENT PROPuLS I ON SYSTEM THROTTLED PERFORMANCE
FS- 1 + 396 Seconds Predi cted Measured S i mu l ati on
245 247 246
236 234 236 236 232 236
6 2 - 61
68 68 68
68 68 68
68 68 68
68 68 68
0 8 . 40 8 . 40
1 2 . 5 - - - - 1 2 . 2
7 . 8 - - - - 7 . 6
1 . 603 - - - - 1 . 605
303 . 4 - - - - 303 . 2
6 1 59 - -- - 6003
9 . 25 - - - - 8 . 53 -�-- -
0
----
FS- 1 + 606 Seconds Pred1 cted Measured r S 1 mul ati on
246 I 247
240 239
240 239
37 -
68 68
68 68
6 8 68
68 68
5 . 57
7 . 7 - - - -4 . 8 - - - -
1 . 604 - - - -
298 . 1 - - - -
3726 - - - -
1 0 .96 - - - -
-- -
I I 245
239
239
37
68
68
68
68
5 . 5 7
7 . 6 4 . 8
1 . 598
298. 2 3697 1 0 . 46
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TABLE 5
DPS P ROPE LLANT GU.'HH I TY GP G I N G S Y S T E �1 P E RF O RMANC E
P a rame te r
Oxi di ze r Tank No . 1 Me a s u re d Quant i ty , pe rcent Cal cul ate d Q�anti ty . pe rcent
Di ffe rence , pe rcen t
Oxi di ze r Tank No . 2 Meas ure d Quant i ty . pe rcen t Cal cu, a ted Quan t i ty . pe rcent
o; ffe rence . pe rcent
FtAe1 Tank No . 1 r�asu re d Qual'\ t i ty . pe rcent Cal cu l a te d Quan ti ty . pe rcent
Oi ffe rence . perce n t
F ue l T ank No. 2 Measu re d Quan ti ty . pe rcent Ca l c u l a ted Quant i ty , pe rcent
-Di f fe rence . o e rce n t --- - - � - ------
4 3
9 4 . 6 95 . 0 - 0 . 4
-95 . 3
-
92 . 5 94 . 0 - 1 . 5
9 1 . 3 94 . 0 - 2 . 7 --
T i me ( F rom Descent Burn I � n i ti on 1 73 24 3 343 445 541
7 3 . 3 6 1 . 5 4 3 . 3 2 9 . 9 1 1 9 . < 72 . 9 6 1 . 0 4 3 . 8 30. 2 1 9 . 7
0 . 4 0. 5 - 0 . 5 -0 . 3 0 . 2
72 . 3 59 . 4 40 . 7 2 7 . 7 1 6 . 6 7 1 . 9 59 . 2 4 1 . 1 2 7 . 2 1 6 . 7 0 . 4 0 . 2 - 0 . 4 0 . 5 - 0 . 1 7 1 . 9 6 0 . 1 42 . 0 28. 6 1 8 . 2 7 1 . 3 5 9 . 1 41 . 4 2 7 . 7 1 7 . 5
0 . 6 1 . 0 0 . 6 0 . 9 0 . 7
7 0 . 6 58. 3 40. 6 2 8 . 0 1 . . 5 7 1 . 3 59 . 1 4 1 . 4 27 . 7 1 i . 5 - 0 . 7 - 0 . 8 - 0. 8 0 . 3 u. o
� � - -
sec 6 4 5
1 2 . 4 1 2 . 1 0 . 3
9 . 2 9. I 0 . 1
1 1 . 0 9. 8 1 . 2 9 . 7 9 . 8 · 0 . 1 ------
,.�·-.
745
6 . 8 6 . 5 0 . 3
4 . 1 3 . 5 0 . 6 5 . 6 4 . 3 1 . 3
3 . 8 4 . 3
- 0 . 5
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I I I I I
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S'teiay.:State Tllrott l e Pos i t i on . Percent
Tot•1 Vacu� St•rt J.,ulse ( FS- 1 to 9U . s te•dy-state ) . l b f - sec -Start Ti� (FS- 1 t� ')� steady-state ) . sec
Coast T i me fro�� Po 1 o r Burn , Mi nutes
Steady-State Throttle Pos i ti on . Percent
Tot•l Vacuum Shutdown I�Y l se ( FS-2 to 1 0 . Steady-State ) , lbf- �ec
f-· Shutdown T i � (FS-2 t� 1 � Ste•dy-State ) . �ec
Repeatabi l i ty . 1bf- iec
Total Vacu� Shutdown I�lse (FS-2 to Zero Thrust) fro- Veloc i ty Gain Dat• . l bf-sec
1 Reference :: .
2Data Unavai l ab l e .
�
FE 7., "n ' . I 't ) 'II v.Lu •• til I ?f ...... eO'' a l ll!J\!I.I .. M#IJ ....... .,. IM'f' ,....,._;_,��
TABLE 6
OPS START MC' SHUTOOWN I MPL: 5E SUMMARY
Apo l l o 1 4 Apol l o 1 2 Apo l l o 1 0 Apo l l o 9 Apol l o 9 Apo l l o 9 Apo l l o 5 Apol l o 5 SPEC I F I CAT ION LM-8/0PS- 1 LH-6/0PS-2 LH-4/0PS-2 LM- 3 1 DPS - l
STARTS
1 3 . 1 1 6 . 2 1 3 . 1 1 2 0 7
7 1 0 591 728 flO�
, I .: • 1 4 1 . 77 2 . 1 3 2 . :i I
i F rnn- r rom I Laun' c : 72 Launch
SIIJTDOWNS . 27 . 0 23. 4 FTP 40
2 976 • 54(: . 204 1 - - - l
1 . 23 2 . 06 0 . 34 1 . 1 i
s s 2948 1 777 - - - - - -
LM- 3/0PS-2 LM- 3/DPS - 3
1 2 . 7 1 2 . 7 · . . ?'1 �5 v ---- -
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1 7Jj
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LM- l / D P S - 2
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3 - - -
3cnavai l ab l e due to APS " F i re- i n- the-Hol e " maneuver 5Not aopl i cabl e to l unar l andi ng s hu tdown 4Spec i H cat i on v a l ue for shu tdowns rerfor. . . ed f r<r FTP on1.: .
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