NASA Apollo 11 Mission Report PDF
NASA Apollo 11 Mission Report PDF
NASA Apollo 11 Mission Report PDF
MSC-00171
I
\
:Jj
\
AS-102 BP-15 Nominal launch and Sept. 18, 1964 Cape Kennedy,
exit environment Fla.
PA-2 BP-23A Second pad abort June 29, 1965 White Sands
Missile Range,
N. Mex.
PREPARED BY
/
Mis s i on Evaluat ion Te am
APPROVED BY
c� � L.c.r--
George M. Low
Manager , Apollo Spacecraft Program
--.,__
CONTENTS
Section Page
/
4.3 EARTH ORBIT COAST AND TRANSLUNAR INJECTION 4-1
4.4 TRANSPOSITION AND DOCKING 4- 2
4.5 'TRANSLUNAR COAST 4-2
4.6 LUNAR ORBIT INSERTION 4- 3
4.7 LUNAR MODULE CHECKOUT 4-4
4. 8 DESCENT PREPARATION . 4-4
4.9 UNDOCKING AND SEPARATION 4-7
4 . 10 LUNAR MODULE DESCENT 4-7
4 . 11 COMMAND MODULE SOLO ACTIVITIES 4-9
4 . 12 LUNAR SURFACE OPERATIONS 4-10
4 . 13 LAUNCH PREPARATION 4-16
4 . 14 ASCENT 4-17
4.15 RENDEZVOUS 4-17
r .#-
4 . 16 COMMAND MODULE DOCKING 4-18
-
'"• 4 . 17 TRANSEARTH INJECTION 4-19
1
Section Page
Section Page
Section Page
A ampere
ac alternating current
AGS abort guidance system
A-h ampere-hour
ALDS Apollo launch data system
arc sec arc second
-�
hr hour
Hz hertz
I inertia
in-lb inch-pound
kW-h kilowatt-hour
LM lunar module
M mega-
mg milligram
min minute
mm millimeter
msec millisecond
N north
NA not available
PM phase modulation
degrees Centigrade
degrees Fahrenheit
a angle of attack
j.l micro-
1-1
1.0 SUMMARY
The space vehicle was laun ched from Kennedy Space Cente r , Flori da ,
at 8 : 32 : 00 a . m . , e . s . t . , July 16 , 1969 . The act ivities during e arth
orb i t checkout , t ranslunar inj e ct i on , t ranspos ition and docking , space
craft e j e ct i on , and translun ar coast were s imi lar to those of Apollo 10 .
Only one mi dcourse corre ct i on , performed at ab out 2 7 hours elapsed t ime ,
was required during t rans lunar coas t .
The spacecraft was insert e d into lunar orb i t at ab out 76 hours , and
the circulari z at i on maneuver was performed two revoluti ons later . Ini ti al
checkout of lunar module systems was s at i sfactory , and after a planned
res t peri od , the Commander and Lunar Module Pi lot entered the lunar module
to prepare for des cent .
s ei smometer , and a las er retro-refle cto r . The Lun ar Module Pi lot evalu
at ed his ab i lity t o operate and m ove ab out , and was ab le to trans late
rapi dly and with confiden ce . Forty-s even pounds of lunar surface materi al
were collecte d t o b e returned for an alys i s . The surface explorati on was
concluded in the allotte d time of 2-l/2 hours , and the crew reentered the
lunar module at lll-l/2 hours .
The return flight started with a 150 -se c ond firing of the s ervi ce
propuls i on engine during the 31st lun ar revoluti on at 135-l/2 hours . As
in trans lunar flight , only one mi dcours e corre cti on was require d , and
passive thermal control was exerci s ed for most of trans earth coast . In
clement weather nece s s it at e d moving the landing point 215 mi les downrange .
The entry phase was n ormal , and the command module lande d in the P ac i fi c
Ocean at 19 5-l/4 hours . The landing coordinates , as determined from the
onb oard computer, were 13 degrees 19 minutes north lat it ude an d 169 de
grees 09 minutes west longitude .
2.0 INTRODUCTION
A complete analysis of all flight data i s not pos sible within the
time allowed for preparation of this report . Therefore, report s upple
ments will be published for the guidance and control sys tem, propul s i on,
the biomedical evaluation, the lunar surface photography, the lunar s ample
analysis, and the traj ectory analys i s. Other supplements will be publi sh-
ed as need is identifi e d .
I n this report, all actual times are e laps ed time from r ange .zero,
established as the i ntegral s e cond be fore lift-of f . Range zero for thi s
·miss ion was 13:32 : 0 0 G .m . t . , July 16 , 1969. All references to mile age
dis tance are in nauti cal miles.
3-l
The descent orbit insertion maneuver was performed with the descent
propulsion system at 101-1/2 hours. Trajectory parameters following this
maneuver were as planned, and the powered descent initiation was on time
at 102-l/2 hours. The maneuver lasted approximately 12 minutes, with
engine shutdown occurring almost simultaneously with the lunar landing
in the Sea of Tranquillity. The coordinates of the actual landing point
3-2
After an 8-hour rest period, the crew began preparations for ascent.
Lift-off from the lunar surface occurred on time at 124:22:00.8. The
spacecraft was inserted into a 48.0- by 9.4-mile orbit from which a ren
dezvous sequence similar to that for Apollo 10 was. successfully performed.
---�,
r
,...
3-4
Time ,
Event
hr : mi n : s ec
Time,
Event
hr:mi n : s e c
Docking 128:03:00
Ascent stage j ett i s on 130:09:31. 2
Separat ion maneuve r ( from ascent stage ) 130:30:01*
Trans earth inj ect i on maneuver 135:23:42. 3*
Second midcours e correction 150:29:57. 4*
Command module /service module s eparat i on 194:49:12.7
Ent ry interface 195:03:05. 7
Landing 195:18:35
NASA·S-69·3700
l
•0 Night • 11 MSFN
US L ift-off
Terminate battery B charge
16 hours charging time)
l nsertion
CYI
Systems checks
TAN
CRO
T
Eat
12
HSK
l
CYI
Prepare for translunar injection
maneuver
22
TAN
CRO
Eat
1
23
Initiate battery A charge
CSM/S·IllB separation
Docking
MSFN
Evasive maneuver
TV !GDSl
T
Initiate passive thermal control
Eat
T
28 t nitiate cabin purge 36
Eat
1
29
T
Sleep
TV IGDSl
1
/;--
31
Terminate cabin purge
48
Initiate battery 8 charge +Eat
-.l
NASA-S-69-3702
T T
52 MSFN MSFN
Eat
1
Eat
f
56 11
Fuel cell purge
Eat
' f
Lunar revolution count tCSMl Lunar revolution count tCSMI
Ground elapsed time ound elap>ed time
"' Day
1
• Night
76 81
I
f MSFN
77 83
78 84
Eat Waste water dump
MSFN
TV
1
MSFN .
Termtnate battery A charge
79 85
-1
Eat
Sleep
,.. _::I._
9
t
80 94
T
Second lunar orbit insertion
maneuver
..... ·<
Initiate battery A charge
10 Eat
MSFN
1 1
3
1 81
(dl 76 to 95 hours.
95
� �
Lunar revolution count (CSMl Lunar revolution count ICSMl
'
ound elapsed time Day Ground elapsed time liay
T _r-
... 102
_l
Night Night
95
Lunar landing
Postlanding activities
l
Lunar Module Pilot transfer to LM
-�-
LM systems check
1
15
98 105
Eat
1
100
MSFN Undock
16
106
MSFN
ICSMl r· '" "'
Prepare for egress to
101 109
Commander egresse s from LM
--
Descent orbit insertion maneuver Lunar Module Pilot
egresses from LM
., . '"""T
Day Day
• Night
't 122 Night
1
110 MSFN Lunar surface activitie � MSFN MSFN
s
IL� ll ICSMl ILMl
Eat
24
TEat
ILMl
_l_
ICSMl
112
MSFN
iCSMl
Lunar photography
l
Fuel cell purge
19
l
113 124 MSFN
ICSMl
1
25
114 125
20 Equipment jettison from LM
MSFN
l l
ICSMI
T
Sleep
Coelliptic sequence
initiation maneuver
1 121
T
Sleep 126 MSFN
t
26
,, 23
+
Constant differential
__. __,.-
height maneuver
Eat
24 ICSMI
1 122 1 127
Ill 110 to 127 hours.
' .J
Lunar revolution count ICSMl Lunar revolution count ICSMl
Ground elapsed time Day Ground elapsed time Day
...
Night
... 133 Night
I
127
I
26
Terminal phase initiation
1
MSFN MSFN
128 135
Docking
27
-1
137
Initiate passive thermal control
Eat
130 138
LM jettison
131 TEat
148
Initiate battery A charge
1 1
t
Fuel cell purge
�
1
Eat
1
MSFN
29
I 133 1 lgl 127 to 149 hours.
149
fl
NASA-S-69-3707
::;�
Ground elapsed time ed time Day
� 149 MSFN
�:�:
Day
• Night s
Night
J::- l n
�
Terminate passive thermal control
!51
T
!59
Eat
1
."T
!53 160
1
T
S leep
j_
r
!54 171
Initiate battery B charge
Eat
Waste water d u mp
.r�
!55
Terminate battery A charge
!56
l
I hi !49 to 175 hours.
Figure 3-1.- Continued.
3-14
NASA-S-69-3708
T
Eat
T
Eat
176
1 191
1
Ter minate passive thermal control
Terminate passive thermal control
1
TV
1 CM/SM separation
178 195
Entry i nterface
Initiate passive ther mal control
Landing
181 196
TEat
182
Sleep
j_
4 .2 LAUNCH
The entire S-II stage flight was remarkably smooth and quiet and the
.l aunch es cape tower and boost protective cover were j ettisoned normally .
The mixture ratio shift was accompanied by a noticeable acceleration
decrease . The S-II /S-IVB st aging sequence occurred smoothly and approx
imately at the predicted time . The S-IVB ins ert i on traj ectory was com
pleted without incident and the automatic guidance shut down yielded an
insertion-orbit ephemeris , from the command module computer , of 102 . 1 by
103 . 9 miles . Communicat i on between crew members and the Network were
excellent throughout all stages of launch .
The digital autopilot was used for the transposition maneuver sched
uled to begin 20 seconds after spacecraft separation from the S-IVB. The
time delay was to allow the command and service modules to drift about
70 feet prior to thrusting back toward the S-IVB. Separation and the be
ginning of transposition were on time. In order to assure a pitch-up
maneuver for better visibility through the hatch window, pitch axis con
trol was retained in a manual mode until after a pitch-up rate of approx
imately 1 deg/sec was attained. Control was then given to the digital
autopilot to continue the combined pitch/roll maneuver. However, the
autopilot stopped pitching up at this point, and it was necessary to re
establish manual control (see section 8.6 for more discussion of this
subject). This cycle was repeated several times before the autopilot
continued the transposition maneuver. Consequently, additional time and
reaction control fuel (18 pounds above preflight nominal) were required,
and the spacecraft reached a maximum separation distance of at least
100 feet from the S-IVB.
Two peri ods of ci s lunar mi dcours e navigat i on, using the command
module compute r program ( P 23 ) , were planne d and execut e d . The first,
at 6 hours, was primarily to estab li sh the apparent horiz on alt itude for
opt i cal marks in the computer. The first determinat i on was begun at a
di stance of approximately 30 000 miles, whi le the s e cond, at 24 h ours,
was designed to accurately determine the opti c al b i as e rrors. Excess
time and fuel were expended during the first peri od becaus e of di ffi culty
in loc at ing the sub stellar poi nt of each st ar. Ground-supplied gimb al
angle s were us ed rathe r than those from the onb oard computer . This t e ch
ni que was devi s e d because c omputer s ol ut i ons are un constrained ab out the
opti cs shaft axi s ; there fore, the comput e r i s unable to predi ct i f lunar
module struct ure might b lock the line of s i ght to the star . The ground
supplied angles prevente d lunar module struct ure from occult i ng the star,
but were not accurat e in loc ating the pre c i s e s ubstellar point, as evi
denced by the fact that the s extant reti cle pattern was n ot parallel to
the horizon . Additi onal maneuve rs were required to achi eve a parallel
reti cle pattern ne ar the point of hori zon-st ar superp osition.
The digital aut opi lot was used to initiate the pas sive thermal con
trol mode at a pos it i ve roll rat e of 0 . 3 deg /se c, with the posit ive lon
gitudi nal axis of the space craft pointed toward the e clipti c n orth pole
.during trans lunar coast ( the eclipt i c s outh pole was the di rect i on used
during trans earth coast ) . After the roll rat e was estab lished, thrus ter
firing was prevented by turning off all 16 switches for the s ervi ce mod
ule thrusters . In general, this method was high ly successful in that it
maintained a s ati s factory space craft attitude for very long periods of
time and allowed the crew to s leep without fe ar of either entering gimb al
lock or encountering unacceptab le thermal condit i ons . However, a refine
ment to the procedure in the form of a new c ompute r routine is requ i re d
to make it foolproof from an operat or ' s viewpoint. [Editor's not e : A
new routine ( rout ine 6 4 ) i s av ai lab le for Apollo 12 .] On s everal occa
s i ons and for s everal di fferent reasons, an incorrect computer-entry
procedure was us ed, resulting in a s li ght waste of react i on control pro
pellant s . Sat i s fact ory plat form alignment s ( program P 5 2, opti on 3 ) using
the opti cs in the res olve d mode and me dium speed were possible whi le ro
t ating at 0 . 3 deg/sec.
The space craft was ins ert e d into a 169 . 9 - by 60 . 9-mile orbit b as ed
on the onboard computer with a 6-minute s ervi ce propul s i on maneuver.
Procedurally, thi s firing was the s ame as all the other servi ce propuls ion
4-4
maneuvers, except that it was started using the bank-B propellant valves
instead of bank-A. The steering of the docked spacecraft was exception
ally smooth, and the control of applied velocity change was extremely
accurate, as evidenced by the fact that residuals were only 0.1 ft/sec
in all axes.
Two entries were made into the lunar module prior to the final activ
ation on the day of landing. The first entry was made at about 57 hours,
on the day before lunar orbit insertion. Television and still cameras
were used to document the hatch probe and drogue removal and initial entry
into the lunar module. The command module oxygen hoses were used to pro
vide circulation in the lunar module cabin. A leisurely inspection period
confirmed the proper positioning of all circuit breaker and switch set
tings and stowage items. All cameras were checked for proper operation.
The crew was awakened according to the flight plan schedule. The
liquid cooling garment and biomedical harnesses were donned. In antici
pation, these items had been unstowed and prepositioned the evening be
fore. Following a hearty breakfast, the Lunar Module Pilot transferred
into the lunar module to accomplish initial activation before returning
to the command module for suiting. This staggered suiting sequence
served to expedite the final checkout and resulted in only two crew
members in the command module during each suiting operation.
4-5
The primary glycol loo p was activate d about 30 minutes early, with
a slow but immediate decrease in glycol temp erature . The activation con
tinued to progress smoothly 30 to 40 minutes ahead of schedule . With the
Commander entering the lunar module early, the Lunar Module Pilot had
more than twice the normally allotted time to don his pressure suit in
the command module .
The early powerup of the lunar module computer and inertial measure
ment unit enabled the ground to calculate the fine gyro torquing angles
for aligning the lunar module platform to the command module platform
before the loss of communications on the lunar far side. This e arly
alignment added over an ho ur to the planned time available for analyzing
the drift of the lunar module guidance syste m .
After suiting, the Lunar Module Pilot entered the lunar module, the
·drogue and probe were installed, and the hatch was closed . During the
ascent-battery checkout, the variations in voltage produced a noticeable
pitch and intensity variation in the already loud noise of the glycol
pump. Suit-loop pressure integrity and cabin regulator re pressurization
checks were accomplished without difficulty . Activation of the abort
guidance system produced only one minor anomaly, An illuminated portion
of one of the data readout numerics failed, and this resulted in some
ambiguity in data readout ( see section 16.2.7).
The probe, drogue, and hatch all functioned perfectly, and the
operation of closing out the tunnel, preloading the probe, and cocking
the latches was done routinely. Previous practice with installation and
removal of the probe and drogue during translunar coast was most helpful.
Two periods of orbital navigation (P22) were scheduled with the lu
nar module attached. The first, at 8 3 hours, consisted of five marks on
the Crater Kamp in the Foaming Sea. The technique used was to approach
the target area in an inertial attitude hold mode, with the X-axis being
roughly horizontal when the spacecraft reached an elevation angle of
35 degrees from the target, at which point a pitch down of approximately
0.3 deg/sec was begun. This technique was necessary to assure a 2-l/2
minute mark period evenly distributed near the zenith and was performed
wi·thout difficulty.
To prevent the two vehicles from slipping and hence upsetting the
docked lunar module platform alignment, roll thruster firings were in
hibited after probe preload until the tunnel had been vented to approxi
mately l psi. Only single roll jet authority was used after the l psi
point was reached and until the tunnel pressure was zero.
4 -T
The des cent orbit ins ertion maneuver was performed with the descent
engine in the manual throttle configuration . Igniti on at the minimum
throttle s etting was smooth , with no nois e or sensation of acceleration .
After 15 s econds , the thrust level was advanced to 4 0 percent , as planned .
Throttle response was smooth an d free o f oscillations . The guided cutoff
left residuals of less than 1 ft /sec in each axis . The X- and Z-axis
res i duals were reduced to zero us ing the react i on control system . The
computer-determined ephemeris was 9 . 1 by 57 . 2 mile s , as compared with the
4-8
Just prior to powere d des cent , the angle b etween the line of s ight
to the sun and a sele cte d axis of the inert i al plat form was compared with
the onboard computer predict ion of that angle and this provi ded a check
on inertial platform dri ft . Three such measurements were all within the
specified toleran ce , but the 0. 08-degree spread between them was somewhat
larger than expect e d .
Visual checks o f downrange and cro s srange pos i tion indi cat e d that
ignition for the powered des cent firing would oc cur at approximatelY the
correct loc at i on over the lunar surface . Bas ed on measurements of the
line-of-sight rate of landmarks , the estimates of alt itudes converged on
a predi cted altitude at ignition of 5 2 000 feet above the surface . These
me asurements were s ligh tly degrade d becaus e of a 10- to 15-degree yaw b i as
maintaine d to improve communications margins .
Ign ition for powered des cent oc curred on t ime at the ffil nlmurn thrus t
leve l , and the engine was automat i c ally advanced to the fixe d throttle
point ( maximum thrus t ) after 26 s econds . Vi sual pos ition che cks indi
cat e d the space craft was 2 or 3 s econds early over a known landmark , but
with very little crossrange e rror . A yaw maneuver t o a face-up position
was init iated at an altitude of ab out 45 900 feet approximate lY 4 minutes
after ignition . The landing radar began receiving alti tude dat a immedi
ately . The alt itude difference , as displayed from the radar and the com
put e r , was approximately 2 800 feet .
Att i tude thrus ter firings were heard during e ach maj or attitude
maneuver and intermittently at other t ime s . Thrus t reduction of the
des cent propuls ion system occurred nearly on t ime ( planned at 6 minutes
24 s econds after ignition ) , contribut ing to the pre di ct i on that the
i.; -9
The method used for target acquisition (program P22) while the lunar
module was on the surface varied considerably from the docked case. The
optical alignment sight reticle was placed on the horizon image, and the
resulting spacecraft attitude was maintained at the orbital rate manually
in the minimum impulse control mode. Once stabilized, the vehicle main
tained this attitude long enough to allow the Command Module Pilot to
4-10
move to the lower equipment bay and take marks. He could also move from
the equipment bay to the hatch window in a few seconds to cross-check
attitude. This method of operation in general was very satisfactory.
Despite the fact that the Command Module Pilot had several uninter
rupted minutes each time he passed over the lunar module, he could never
see the spacecraft on the surface. He was able to scan an area of approx
imat e ly l square mile on each pass , and ground estimates of lunar module
position varied by several miles from pass to pass . It is doubtful that
the Com mand Module Pilot was ever looking precisely at the lunar module
and more likely was observing an adjacent area. Although it was not pos
sible to assess the ability to see the lunar module from 60 miles, it was
apparent there were no flashes of specular light with which to attract
his attention.
The visibility through the sextant was good enough to allow the
Command Module Pilot to acquire the lunar module ( in flight) at distances
of over 100 miles. However, the lunar module was lost in the sextant
field of view just prior to powered descent initiation ( 120-mile range)
and was not regained until after as cent insertion ( at an approximate range
of 250 miles), when it appeared as a blinking light in the night sky.
ground for no obvi ous reas on . Out s ide the vehi cle , there were no appre c i
ab le communi cat ion problems . Upon i ngre s s from the surface , these diffi
cult ies re curre d , but under di fferent condi tions . That i s , the voice
dropout s to the ground were not repeatable in the s ame manner .
Depre ssuri zat ion of the lunar module was one aspect of the mi s s ion
that had never been completely performe d on the ground. In the vari ous
alt itude chamber t e s ts of the spacecraft and the extravehi cular mobility
un i t , a complete s et of auth ent i c conditions was never pres ent . The de
pressuri zation of the lunar module through the bacteria filter took much
longer than had been anticipat e d . The indi cated cabin pre s sure did not
go below 0 . 1 psi , and some concern was experienced in opening the forward
hat ch against thi s res idual pres sure . The hat ch appeared t o bend on ini
t ial opening , and small p art i cles appe ared t o be b lown out around the
hat ch when the seal was broken ( s ee s ection 16 . 2 . 6 ) .
Simulat ion work i n both the water immersion fac ility an d the 1/6-g
environm�nt in an airplane was reasonab ly ac curate in preparing the crew
for lunar n[odule egre s s . Body pos i tioning and arching-the-back t e chniques
that were requir�d to exit the hat ch were performe d , and no unexpected
problems were experienced. The forward plat form was more than adequat e
to allow changing the b ody pos ition from that used in egres sing the hat ch
to that required for getting on the ladde r . The first ladder step was
s omewhat di fficult to see and require d caution and forethough t . In gen
eral , the hatch , porch , and ladder ope1•at ion was not part icularly diffi
cult and caused little concern . Operat i ons on the plat form could b e
performe d without los ing b o dy balance , and there was adequate room for
maneuvering .
4 . 12 . 4 Surface Exploration
The s olar wind experiment was easily deployed . As with the other
operations involving lunar surface penetration , it was only poss ible to
penetrate the lunar surface material about 4 or 5 inches . The experiment
mount was not quite as stable as desire d , but it stayed erect .
Collecting the bulk s ample required more time than anticipated be
cause the modular equipment stowage assembly table was in deep shadow ,
and collecting s amples in that area was far less desirable than taking
those in the sunlight . It was also desirable to take s amples as far from
the exhaust plume ·and propellant contamination as pos sible . An attempt
was made to include a hard rock in each s ample , and a total of about
twenty trips were required to fill the box . As in simulations , the dif
ficulty of s cooping up the material without throwing it out as the s coop
4-14
became free creat e d s ome problem. It was almost impos s ible to collect a
full scoop of material , and the task required about double the planned
time .
No abnormal condit ions were noted during the lunar module inspection .
The insulation on the s econdary struts had been damaged from the heat ,
but the primary struts were only singed or covered with soot . There was
much less damage than on the examples that had been seen before flight .
Obt aining the core tube s amples presented s ome diffi culty . It was
impos s ible to force the tube more than 4 or 5 inches into the surface ma
terial , yet the material provided insufficient resistance to hold the ex
t ens ion handle in the upright pos ition . Since the handle had to be held
upright , this pre cluded using both hands on the hamme r . In addition , the
res istance of the suit made it diffi cult to steady the core tube and still
swing with any great force . The hamme r actually missed several times .
Sufficient force was obtained to make dents in the handle , but the tube
could only be driven to a depth of about 6 inche s . Extraction offered
little or virtually no resistance . Two s amples were taken .
s ublimat or st artup and operated at maximum pos ition for 42 minutes b e fore
switching to the intermediate position . The switch remained in the inter
mediate position for the durat ion of the extravehicular activity . The
thermal effect of shadowed areas versus those areas in sunlight was not
detectable ins ide the suit .
The crewmen were kept phys i cally cool and comfortable and the ease
of performing in the 1/6-g environment indicate that tasks requiring
greater physi cal exertion m8lf be undertaken on future flights . The Com
mander experienced s ome physi cal exertion while transporting the s ample
return container to the lunar module , but his physi cal limit had not been
approached .
Because of the bulk of the extravehi cular mobility unit , caution had
to be exercised to avoid bumping into switches , circuit breakers , and
other controls while moving around the cockpit . One circuit breaker was
in fact broken as a result of contact ( see section 16 . 2 . 11 ) .
Equipment j ettison was performed as planned , and the time taken before
flight in determining the items not required for lift-off was well spent .
Consi derable weight reduction and increase in space was realized . Dis
carding the equipment through the hatch was not diffi cult , and only one
item remained on the platform . The post-ingress checklist procedures were
performed without difficulty ; the checklist was well planned and was fol
lowed precisely .
4-16
The rest period was almos t a complete los s . The helmet and gloves
were worn to relieve any sub concious anxi ety about a los s of cabin pres
sure and pre s ented no problem. But noi s e , lighting , and a lower-than
des ired temperature were annoying . It was uncomfort ab ly cool in the suits ,
even with wat er-flow dis connect e d . Oxygen flow was finally cut off , an d
the helmets were removed , but the noi s e from the glycol pumps was then
loud enough to interrupt sleep . The window shades did not completely
block out light , and the cabin was illuminated by a comb inat i on of light
through the shade s , warning lights , and display lighting . The Commander
was resting on the as cent engine cover and was b othere d by the light enter
ing through the telescop e . The Lunar Module Pilot estimat e d he slept fit
fully for perhaps 2 hours and the Commander did not sleep at all , even
though body positioning was not a prob lem. Becaus e of the re duce d gravi ty ,
the posit ions on the floor and on the engine cover were b oth quite comfort
able .
4 . 13 LAUNCH PREPARATION
Aligning the plat form b e fore lift-off was compli cat e d by the limited
number of stars availab le . Becaus e of sun an d earth interference , only
two detents effect ively remained from which to s elect stars . Ac curacy is
gr·eat er for stars clo s e to the center of the field , but none were avail
ab le at this loc at i on . A gravity/one-star alignment was succes s fully per
formed . A manual averaging te chnique was used t o sample five succes sive
cursor readings and then five spiral readings . The re sult was then enter
ed into the computer . This te chni que appeared to be easier than t ak ing
and entering five separate re adings . Torquing angles were clos e to
0 . 7 degree in all three axes and indi cat e d that the platform did drift .
( Editor ' s note : Plat form drift was within specifi cat i on limits . )
4 . 14 ASCENT
4 . 15 RENDEZVOUS
All four sources for the coellipti c sequence initi ation s olution
agree d to within 0 . 2 ft /sec , an accuracy that had never been observed
be fore . The Commander elected to use the primary gui dance solution with
out any out-of-plane thrusting .
During the coellipt i c phas e , radar tracking data were inserted into
the abort guidance system to obtain an independent intercept guidance
s olut i on . The primary guidance s olution was 6-l/2 minutes later than
planne d . Howeve r , the intercept traj ectory was quite nominal , with only
two small midcourse corrections of 1 . 0 and 1. 5 ft /sec . The line-of
sight rates were low , and the planned braking s chedule was used to reach
a ·station-keeping position .
this maneuver was in progres s , all twelve docking latches fired and
docking was completed s ucces s fully . ( See section 8 . 6 . 1 for further dis
cussion . )
Following docking , the tunnel was cleared and the probe and drogue
were stowed in the lunar module . The items to be trans ferred t o the
comman d module were cleaned using a vacuum brush attached to the lunar
module suit return hose . The suction was low and made the process
rather tedi ous . The s ample return containers and film magazines were
placed in appropriate b ags to complete the trans fer , and the lunar
module was configured for jettison according to the checklist procedure .
4 . 17 TRANSEARTH INJECTION
The time between docking and transearth injection was more than
adequate to clean all equipment contaminated with lunar surface material
and return it to the command module for stowage so that the necess ary
preparations for transe arth inj e ction could be made . The trans earth in
jection maneuver , the last service propulsion engine firing of the flight ,
was nominal . The only difference between it and previous firings was
that without the docked lunar module the start transient was apparent .
4 . 18 TRANSEARTH COAST
4 . 19 ENTRY
4 . 20 RECOVERY
On the landing , the 18-knot surface wind filled the parachutes and
immediately rotated the command module into the apex down ( stable I I )
flotation position prior t o parachute release . Moderate wave-induced
oscillations accelerated the uprighting sequence , which was completed in
less than 8 minutes . No difficulties were encountered in completing the
postlanding checklist .
The powered descent traj e ctory was designed consi dering such factors
as opt imum propellant us age , navigat i on uncertainties , landing radar p er-·
formance , terrain uncertaint ies , and crew visibility restrictions . The
basic premi se during traj ectory des ign was to maintain near-optimum us e
of propellant during initial braking and t o provi de a standard final
approach from which the landing area can be ass es s ed and a desirable
landing locat ion s elected. The onboard guidance capability allows the
crew to re-des ignate the desired landing pos ition in the computer for
automat ic execution or, i:f late in the traj e ct ory , to take over manually
and fly the lunar module to the desired point . To provi de these des cent
characteristics , compat ibility between the automat i c and manually con
trolled trajectories was required , as well as acceptable flying quality
under manual control . Because of guidance dispersions , s ite-select ion
Uncertainties , vi s ibili ty restri ct ion , and unde fined surface i rregulari
ties , adequat e flexibility in the terminal-approach te chnique was pro
vided the crew , with the principal limitation b eing des cent propellant
quantity .
The maj or phases of powered des cent are the braking phase (wh i ch
terminates at 7700 feet alt itude ) , the approach or vis ibility phas e (to
approximat ely 500 feet alt itude ) , and the final landing phas e . Three
separate computer programs , one for each phas e , in the primary gui dance
system execute the desired traj e ctory s uch that the various pos ition ,
velocity , accelerat ion , and vi sibility constraints are satisfied. These
programs provide an automat i c gui dance and control capab ility for the
lunar module from powered des cent initiat ion to landing . The braking
phase program ( P63 ) is initiat ed at approximately 40 minutes be fore de
s cent engine ignition and controls the lunar module until the final ap
proach phas e program ( P64 ) i s automat i c ally entered to provide traj ectory
condit ions and landing site vis ibility .
If des ired during a nominal des cent , the crew may s elect the manual
landing phas e program ( P66 ) prior to the completion of final approach
phase program P64 . If the manual landing phas e program P66 is not entere d ,
the automat ic landing program ( P6 5 ) would b e entered automati c ally when
5 -2
Throughout the des cent , maximum use was made onboard , as well as on
the ground , of all dat a , system responses , and cues , bas ed on vehicle
position with respect to des ignated lunar features , to assure proper
operation of the onboard systems . The two onboard guidance systems pro
vided the crew with a continuous check of selected navigation parameters .
Comparis ons were made on the ground between data from each o f the onboard
syst ems and comparable information derived from tracking dat a . A powered
flight proces s or was used to simultaneous ly reduce Doppler tracking data
from three or more ground stations and calculate the required parameters .
A filtering t echnique was use d to compute corrections to the Doppler
tracking data and thereby define an accurate vehicle state vector . The
ground data were use d as a voting s ource in case of a slow divergenc e be
tween the two onboard syst ems .
The crew entered and began activation of the lunar module following
the first s leep period in lunar orbit ( see section 4 . 8 ) . A list ing o f
s igni fic ant events for lunar module descent i s presented i n t able 5-I .
Following des c ent orbit insertion , rendezvous radar data were recorded
by the Lunar Module Pilot and used to predict that the pericynthion point
would be at approximately 50 000 feet altitude . Initial checks using the
landing point des ignator capability produced close agreement by indi c ating
52 000 feet . The crew also reported that a s olar sighting , performed
following des cent orbit insertion and using the alignment telescope , was
well within the powered des cent initiation go/no-go criterion of 0 . 25 de
gree . The s olar s ighting consisted of acquiring the sun through the tele
s cope and comparing the actual gimbal angles to those theoretically re
quired and computed by the onboard computer for this observation . This
check is an even more accurate indication of platform performance if the
0 . 07-degree bias correction for the teles cope rear detent position is
s ubtracted from the recorded dat a .
The powered des cent maneuver began with a 26-s e c ond thrusting period
at minimum throttle . Immediately after ignition , S-b and communi cations
were interrupted moment arily but were reestablished when the antenna was
switched from the automati c to the slew pos ition . The des cent maneuver
was initiated in a �ace-down attitude to permit the crew to make time
marks on s elected landmarks . A landing-point-designator s i ghting on the
crater Maskelyne W was approximately 3 s e conds early , con�irming the sus
pected downrange error . A yaw maneuver to �ace-up att it ude was initiated
�allowing the landmark s i ghtings at an indicated alt itude o� about
4 5 900 �eet . The maneuver took longer than expe cted because o� an incor
rect setting o� a rate displ� switch .
Landing radar lock-on occurred be�ore the end o � the yaw maneuver ,
with the spacecraft rotating at approximately 4 deg/s e c . The altitude
difference between that calculate d by the onboard computer and that deter
mined by the landing radar was approximately 2800 feet , which agrees with
the altitude error suspected from the Doppler res i dual comparis on . Radar
altitude updates of the onboard computer were enabled at 102 : 38 : 45 , and
the differences converged within 30 seconds . Velocity updates began auto
matically 4 se conds after enabling the altitude update . Two altitude
difference transients oc curred during computer alarms and were apparently
ass oc i ated with incomplete radar data readout operations ( see s ecti on 16 . 2 . 5 ) .
The reduction in throttle setting was predi cted to occur 384 se conds
after ignition ; actual throttle reduction occurred at 386 se conds , indi
cating nominal performance of the des cent engine .
Arrival at high gate ( end of braking phas e ) and the automatic switch
to final approach phase program P64 occurred at 7129 feet at a descent rate
o� 125 ft /sec . Thes e values are s li ght ly lower than predi cted but within
acceptab le b oundaries . At about 5000 feet , the Commander switched his
control mode from automati c to attitude-hold to check manual control in
anticipation of the final des cent .
After the pit chover at high gate , the landing point des ignator indi
cated that the approach path was leading into a large crater . An unplan
ned redesignation was introduced at this time . To avoid the crater , the
5-5
Commander again swit ched from automat ic to att it ude-hold control an d man
ually incre as e d the flight-path angle by pitching to a nearly verti cal
attitude for range extension . Manual c ontrol began at an altitude of
approximately 600 feet . Ten s e conds later , at approximately 400 feet ,
the rate-of-des cent mode was activated to control des cent velocity . In
this manner , the spacecraft was gui ded approximately 1100 feet downrange
from the initial aim point .
The powered flight proces s or data refle ct both the altitude and down-·
range errors existing in the primary system at powered des cent init i ation .
The radial velocity error is directly proporti onal to the downrange posi
tion error such that a 1000-foot downrange error will cause a 1-ft /s e c
radial velocity error . Therefore , the 20 000-foot downrange error exi st
- ing at powered des cent initiation was also reflected as a 20-ft/s e c radial
velocity res i dual . This error is apparent on the figure in the altitude
region near 27 000 feet , where an error o f approximately 20 ft /sec is evi-·
dent . The primary-system altitude error in exi stence at powered des cent
initiation mani fests it s elf at touchdown when the powered fli ght proces
s or indicates a landing altit ude below the lunar surface . Figure 5-4
c ontains a similar c omparis on of lateral velocity from the three sources .
Again , the divergence noted in the final phas es in the abort gui dance
system data was caus e d by a lack of radar updates .
data and known surface features . The coordinates o f the landing point ,
as obtained from the various real-time and postflight s ources , are shown
in table 5-IV . The actual landing point is 0 degree 41 minutes 15 sec
onds north latitude and 23 degrees 26 minutes east longitude , as compared
with the targeted landing point of 0 degree 43 minutes 53 seconds north
latitude and 23 degrees 38 minutes 51 s econds east longitude as shown in
figure 5-10. Figure 5-10 is the b as i c reference map for location of the
landing point in this report . As noted , the landing point dispersion was
caus ed primarily by errors in the onboard stat e vector prior to powered
des c ent initiation .
The inertial measurement unit was aligned three times during this
period using each of the three available lunar surface alignment options .
The alignments were s atis factory , and the results provided confidence in
the technique . The s imulated countdown was terminated at 104-1/2 hours ,
and a partial power-down of the lunar module was initiated.
us ed for the sightings . It can be seen that the actual landing site , as
determined from films taken during the des cent , did not lie near the cen
ter of the sextant field of view for any of the coordinates used ; there
fore , the ability to acquire the lunar module from a 60-mile orbit can
neither be est ablished nor deni e d . The Command Module Pilot reported i t
was pos s ible to s can only one grid square during a single pas s .
Bec ause of the unsuccess ful attempts to sight the lunar module from
the command module , the decision was made to track the command module from
the lunar module using the rendezvous radar . The command module was ac
qui red at a range of 79 . 9 miles and a closing rate of 3236 ft /sec , and
los s of track occurred at 85 . 3 miles with a receding range-rate of
3 5 31 ft /sec ( fig . 5-15 ) .
The inertial meas urement unit was success fully aligned two more times
prior to li ft-off , once to obtain a dri ft check and once to establish the
proper inertial orientation for lift-off . The dri ft check indi c ated nor
mal system operation , as dis cussed in section 9 . 6 . An abort guidance sys
tem alignment was als o performed prior to li ft-off ; however , a procedural
error caus ed an azimuth mis alignment which resulted in the out-of-plane
velocity error di s cus s e d in section 9 . 6 . 2 .
5.6 ASCENT
Preparati ons for as cent began after the end of the crew rest period
at 121 hours . The command module state vector was updated from the ground ,
with coordinates provided for crater 130 , a planned landmark . This cra
ter was tracked using the command module sextant on the revolution prior
to lift-off to establish the t arget orbit plane . During this s ame revo
lution , the rendezvous radar was used to track the command module , as
previ ously menti oned , and the lunar surface navigation program ( P22 ) was
exercised to establish the location of the lunar module relative to the
orbit plane . Crew activities during the preparation for launch were con
ducted as planned , and lift-off occurred on time .
The crew reported that the ascent was smooth , with normal reaction
c ontrol thruster activity . The ascent st age appeared to "wallow , " or
travers e the attitude deadbands , as expe cted . Figure 5-17 contains a
time history of s elected control system parameters during the ascent ma
neuve r . A dat a dropout occurred immediately after li ft-off , making it
diffi cult to determine accurately the fire-in-the-hole forces . The body
rates recorded just prior to the dat a dropout were small ( less than 5 deg/
s e c ) , but were increasing in magnitude at the time of the dropout . How
ever , crew reports and ass ociated dynami c information during the data
loss period do not indicate that any rates exceeded the expected ranges .
The predominant disturbance torque during ascent was about the pitch
axis and appe ars to have been caused by thrust vector offs et . Fi gure 5-18
contains an expanded view of control system parameters during a selected
period of the ascent phase . The digital autopilot was designed to con
trol about axes offset approximately 45 degrees from the spacecraft body
axes and normally to fire only plus X thrusters during powered as cent .
There fore , down-firing thrusters 2 and 3 were used almost exclusively
during the early phases of the ascent and were fired alternately to con
trol the pitch disturbance torque . These j ets induce d a roll rate while
counteracting the pitch disturbance ; therefore , the accompanying roll
motion contributed to the wallowing sens ation reported by the crew . As
the maneuver progressed , the center of gravity moved toward the thrust
vector , and the resulting pitch disturbance torque and required thruster
activity decreased until almost no disturbance was present . Near the end
of the maneuver , the center of gravity moved to the opposite side of the
thrust vector , and proper thruster activity to correct for this oppos ite
disturbance torque can be observed in figure 5-17 .
5.7 RENDEZVOUS
Soon after the terminal phase initiation maneuver , the vehi cles
pas s e d behind the moon . At the next acqui s it i on , the vehi cles were fly
ing formation in preparation for docking . The crew reported that the
rendezvous was nominal , with the first midcourse maneuver les s than 1 ft /
s e c and the se cond about 1 . 5 ft /sec . The midcourse maneuvers were per
formed by thrusting the body axis components to zero while the lunar mod
ule plus Z axis remained pointed at the command module . It was als o re
ported that line-of-sight rates were small , and the planned braking was
us ed for the approach to station-keeping . The lunar module and command
module maneuver s oluti ons are summari zed in tables 5-VI and 5-VII , respec
tively .
During the docking maneuver , two unexpe cted events occurred . In the
alignment procedure for docking , the lunar module was maneuvered through
the platform gimbal-lock attitude and the docking had to be completed
using the abort gui dance system for attitude control . The off-nominal
attitude resulted from an added rot ation to avoid sunlight interference
in the forward windows . The sun elevation was about 20 degrees higher
than planned because the angle for initiation of the terminal phase was
reached about 6 minutes late .
damped to within plus or minus 3 degree s , and then initi ate retract to
achieve hard docking . The Commander detected the relatively low velocity
at init i al contact and applied plus X thrusting ; however , the thrusting
was continued unti l after the mis alignment excursi on had developed , since
the Commander had received no indication of the capture event . To further
complicate the dynamics , the Command Module Pilot also noticed the excur
s i ons and reversed the command module control mode from CMC FREE to CMC
AUTO . At this time , both the lunar module and the command module were in
minimum-deadband attitude-hold , thereby causing con s i derable thrust er fir
ing until the lunar module was placed in maximum deadband . The vehicles
were stabi li zed using manual control just prior to achieving a succes s ful
hard dock . The initial observed mis alignment excursion is cons idered to
have been caused by the continued lunar module thrusting following c ap
ture , s ince the thrust vector does not pass through the center of gravity
of the command and service modules .
The rendezvous was succes s ful and similar to that for Apollo 10 ,
with all guidance and control systems operating s atis factorily . The
Command Module Pilot reported that the VHF ranging broke lock about 25
times following as cent insertion ; however , lock-on was reestablished
e ach time , and navigation updates were successful . The lunar module
reaction control propellant us age was nearly nominal .
5 -13
Time ,
Event
hr :min : s ec
X -0 . 1 0.0
y -0 . 4 -0 . 4
z -0 . 1 0.0
a
TABLE 5 -IV . - LUNAR LANDING COORDINATES
b Radius of
Lat itude , Longi tude ,
Data source for s olut ion Landing Site 2 ,
deg north deg east
miles
Photography 0 . 647 or 23 . 5 05 or
c c
0°41' 15" 2 3°26 ' 00"
a
Following the Apollo 10 mi ssion , a difference was noted ( from the
landmark tracking result s ) between the t rajectory coordinate system and
the coordinate system on the reference map . In order to re ference tra
j e ctory values to the 1 : 100 000 s cale Lunar Map ORB-II-6 ( 100 ) , dated
December 1967 , correction factors of plus 2 ' 2 5 " in latitude and minus
4 ' 17" in longitude must be applied t o the traj ectory values .
b
All lat itude values are correct e d for the estimated out-of-plane
position error at powered des ce nt initiat ion .
c
These coordinate values are referenced to the map and include the
correction factors .
5-16
Radial Downrange
Alt it ude ,
Source velocity , velo city ,
ft
ft / s e c ft/sec
X -2 . 1 ft / s e c
Y = -0 . 1 ft/ sec
Z + 1 . 8 ft/ s e c
The orbit result ing aft e r res iduals were trimmed vas :
Maneuver
Time, Velocity , Time , Velocity , Time , Velocity , Time , Velocity ,
Solut ion
hr:min:sec t:t/sec hr:min:sec t:t/sec hr:min:sec ft/sec hr:min:sec ft/sec
8.1 retrograde
Initial 126 ,17,46 . 36 1.8 south
8.0 retrograde
Constant differential 17. 7 up 5-1 retrograde
(a) (a) 126 : 17,42 126,17 ,50 1.7 south
height 11.0 up
18 . 1 up
8.1 retrograde
Final 126 , 17 : 46 . 36
18.2 up
25 . 2 forward
Initial 127 , 0 3 , 16 . 12 1.9 Tight
Te�inal phase 0.4 down 22 . 4 posigrade 22 . 9 pos igrade
c 127 ,03,39 23.4 total 12 6,57 ,00 0.2 north 127 :03:52 1. 4 north
initiationb ,
25 . 0 forward 11.7 up 11.0 up
Final 127 ,03,31.60 2.0 right
0.7 down
0.0 forward
First m.idcourse
Final 127 ,18,30 . 8 0.4 right (a) (a) 127,12,00 0.0 ( d) ( d)
correction
0.9 down
0.1 forward
Second midcourse
Final 127 ,33,30.8 1.2 right (a) (a) 127 , 2 7 ,00 0.0 (d)
correction (d)
0.5 down
bBo�-axis reference frame ; all other solutions for locea-vertical reference frame .
c
For comparing the primary guidance solution for terminal phase initiation with the real-time nominal and actual values , the following components are
equivalent to those listed but with a correction to a local-vertical reference frame : 22 . 7 posigrade , 1.5 north , and 10.6 up.
d
Data not available because of moon occultation.
\.n
I
.....
-.J
5 -18
a
initial comput ed t ime of igni tion using nominal elevat ion angle
of 208 . 3 degree s for terminal phas e initi at ion .
b
Final solution us ing lunar module t ime of igni t i on .
NOTE : All s olut ions in local hori zontal coordi nat e frame .
NASA-S-69-3 7 09
Undocking
--a.a. Sun
Powered
descent
Earth
I
Figure 5-l . - Lunar descent orbital events .
V1
t)
I
5-20
NASA-S-69-lllO
�j_AX------ - +X
Z
�
�.�'-�� �\z\ +
,z
\ '•
. .y 7·
�X
� ·
I
I
I
:
T
·
I I
:
..-::]" ""'" ·: /
�- �·· :
7200 It ;7 I
I
I
I
I I
4 Radar Crater Maskelyne W 260
High gate acquisition
Approxi mate range to landing point, mi
I T
+Z�
5000 14 000 26 000
+X
f- !
T
--
+Z - fc
'I
-------
�
fif�
,z __:,:\(.
L'!' c\
400 ft
I
Landing 2500 5000
A utomatic landing program Attitude hold
Approxi mate range to landing point, ft
90 X 103
- �-- � -
� :�:_::.1-"-
-- -
--
1
-. L
-
--
50 Pri ry gu -
�-
-...
__ -· - -
-
--·
2 - --
-- -·-
-·
= ...-·
-· -- -·
·-·
--
�-
.;
:::J 40
.,
--
-
�ro;:
""
� 1'.
�
p
� 3
4
,.- ;:;
-f" k-r- --
30 5
- :;;.�
.
--s: -
f'-. .._
�--�!---
·-·
"'
--
- --- . �
6
��
20
--
_..,�--
- -- -
�_ {
_...
-.....
<---, r Primary guidance _r Abort guidance
r- l£7 �---
10
--
·:��.\..
.., --
t---f- - - - - t--- tr 8
.
-- ,/
--
'- ..
\(
-
-
-
--
[.-
·- · -
- -
--- -
etrork v
- -- ·-· -· - -
PL
- --� --
-- · · --
0 t-
r--
- ·-
N
·-· -·-
--- - --- - -
9-
-10
-150 - 140 - 130 -120 -no -100 -Ill -70 -60 -50 -40 -Jl -20 - 10 0 10 20
Mitude rate, It/sec
\.11
I
Figure 5-3. - Comparison of a�itude and altitude rate during descent. 1\)
1-'
NASA-S-69-3712 \.J1
I
1\)
1\)
60
•
50
•
• .•
lo
-• /
• •••
f- Network
. .. . . . .
.. - . •.
•
40
•
. • .•
.
• • ,
�"'
..
•
"
" 30
•/ . ;-. Approach phase Manual landing
K
"'
? /
program (P64) phase program (P66)
,
"
"
.£
. ··"'
I Land ing
·� .
" 20
>
I
• i
\. • l,r-Abort gu idance
I
� .
10
. � r )\ 1::-• -'
. ./'
•. . .,... ...... . .
\•· •• . . ....
•' �'-"-I,. 1\....�4\• •
- ,-
0
... =-- "
v .. .
.
l . .
.... ,
•
., .
-/
Primary guidance .
.
•
• I
!
I
I
-10
1 0 2:34 1 0 2 :36 102:38 10 2:40 1 0 2:42 1 0 2 :44 1 0 2:46 1 0 2:48
T ime/ hr:min
NASA-S-69-3713
14 0
120
10 0
I""
�
'
....
� Primary guidance .
80
�
"' Abort guidance -f.'
"'
"0
� I
I P-64 Approach phase program
I
•'1.;
�
"'
c
-"'
"'
Ia:. I
60
N-
'
-;;;
E
"'
Manual landing
.<::
I
u
::!
c..
40
\ P-66
phase program
I
\ I
20
'I I
�: I Landi �g
LJ.
0 -
"\t A
v
340
1 0 2 :3 2 102:34 102:36 102:38 1 0 2: 4 0 1 0 2 : 42 102:44 1 0 2:46
T ime , hr:min
30
en
m 20
�
P -66 Manual landing phase
2!.- I
"' �
en
c:
10
I
II i
i Lan ing
V'� t...
.l!!
""',,..'-<-,i I;'
E P r i mary guidance I
1 -
"0'> I
""
0
\� 4� v ·y )' .j
I
_/\ /
-��
0
I
Abort guidance I
i
350 i
102:43:00 43:20 43:40 44:00 44:20 44:40 45:00 45:20 45:40 46:00
Time, hr:min:sec
I
4
-� !.
Lan ing �
�
3
i
i i i
ii
I
I
i
r
I
2
i !
! !
I
.!/\ f'\,. i
P-66 Manual landing phase i . i i
I , "
i ! .
I
"' \�
I • I
]J i I •
i
i
I
J
""""
m
I . !
� I
( \
!
�
"'
! i i.
"0'>
c:
"'
0
1
I i
I
i
; i
i
I
\\ �
E I I
�
i
i
"0'> P rimary g uidance
I
.c:
i
"-
i
359
�
\\, l ! I
'
i
I
I I
/ �
Abor guid nce --\ �!\ ! ;
I
358
I
i ! ! I
I
I
I
i
i i I
I I
JI
I I
357
i
I
\d
'
i t�
),;'
!
/! I
i
356
'
I I
I
355
102:43:00 43:20 43:40 44:00 44:20 44:40 45:00 45:20 45:40 46:00
Time, hr:min:sec
Figure 5-6. - Expanded pitch and roll attitude time histories near landing.
/ ػՕ
խȉሤ
ĝሤ Ğሤ ᇮሤ ğሤ
ě ̔̕ נ ҫ
ߦ߷
ݕݔҀ߷
Օ߷
ѯχђѓ҃
ᱎᱎ ᱎ
Ìᱎ ᤂᯍᱎ
ॕ
,&/
֗˽
- '/ ˾ & غ/ #!)! ' $/ ! / / ( */ %(/ ( ". / &",/ +& //
ŃƒƈŀƓƚɳɴɵƐłʣ Ê
ןի ሤ
# ! ( ! ( + +
ᇭ Ȉሤ
ሤ
ညᱎ
ഽᱎ
ᄮሤ
߷ݓ
ѿ߷
যሤ ęሤ
ࡦ
Òז
ഘቷቷᤁᱎ
ഘ ᐮ፴ុቷᱎ ੳ ̂ ̃ ೳ ᘧ ᔅ ႀុ፴ቷᇱᱎ ᗁႀᝤᱎ ᚹጢᱎ ᘨႀុᱎ ᔆႀᘩᇱ ᐮ ᘪ፴ᱎ ႀុቷႀ ᱎ
ن֖ ˼ ࠒࠟݭ۽ ۩܍Ħࠓ ܩĈߗ۞ ܪĈ Ħ ߋ ࠄ ࠌࠫ ۼ܌߮ߖݬܣФ
Æᱎ ȓ ÄᱎQז
֑ ď߷
ߤ߷ ߷ݏ ߢ߷
˺˻˹̎ У ሤ Ĕሤ ĕ Ė ե ሤ
Т ֲז ᬟᱎ
˸ өӪӫ
ߕ АБժВ ሤᇪ̓ᇫͬ
ՠ դաሤ Qז Ãᱎ Ȕז ۛʩ ʪ
߷
b̏̐̑
٧߷ ߷ݐ
ଚ ଛ
ᄫሤ
ȟᇩሤ հ߷
ƄȒ
߷ݑ
٫ᱎÇ Èᱎ
Ӄ߷ ᯈᱎG ᱎߥ߷
üᱎ ȅሤÅᱎ߷ ඵሤ ߡ߷Ñז
ᄭėሤ Ę Ȇ ሤ
ēሤ
ѻՔ Đ߷
ѷ߷ߣ߷ Ѹ
ѹ߷
բ գ ሤ
ۑېۏա զ ሤ
ሤ
ȕ זµִזɨɩȖז
םሤ
ݴ
ޏ
රᱎ
ֱȉȊఝ ڡ$ " $
ʱ ʲ ʳ
යᱎ
ͱᇧ w ʴ ʵ ʶ w ˼ᯌ˷˽
ᱎ ˽˸˹˺˻˾˽˽˽˽˼ I
߷ Ξ҃
2 2 2
I I 2
߷
ሤ Ѿ
֣߷ v ଡ଼vଢ଼ሤH߷ō ΝΞΟΠΡ߷
ሠሡሢሤ
ࢾ߷ঋÊਸ߸Ëᱎ༎ᱎ &&
& &
$
$
ฏሤÍ ᱎ
ᇇᱎ
ᒝᱎ
֟ሤ
ְז
ሤ
ሤ
Ɓ
˖ ː ۾՚ =՛ ՜ ՝ᄩ ՞ ̍
Ӥזcdՙ ̚
Ř ·҃ ~ vŘŘ
ሤ
ēʣ
ʤ ۚ ߷ ʥ
$ " )" ( * +
ٛ߷ؼ
ýሤ
ćߘÃ߷
dzሤᇕ ૉሤ ߓ ݈߷
Ԫەѳ ߷ۖوه ڄȕՔ֕æ
Ѥ߷ Ĺ߷ ߖ ѥ ߷
ǶǷሤ
ථࣈሤ ࢊބबሤ
Ʃןׯᱎ ֢
M߷
ĈႢ
དྷ[ሤ
עᱎ ࢋƆमሤ
ࢍ ߲ ᄟሤ
҃˲Ƈ ьÇ҃
˚ ƥ ᱎ ᱎ £҃ Xሤ
ɰāĘƖʣ
ċ Ūࣇ ߮ Ƅ ሤ
Ⴆ
ᇡ ̈ ྒሤ
Ć זƉÈя҃ ᇓሤ ఱ উঊडᱎ
ŭࢯ߶ވሤ
̇௶ሤ
כሤ๔ ௹ চ ሤ ᄡ ๕ ฎ ሤ
߳ ࡙ޅঙሤ
¨џ߷
ȗȘ
ăሤ
ຟᱎץᱎ ߰ሤ௵ሤ ֏ ֮֯ ĉ Ċٴ
๒ ā ሤ߱ࢢƅሤ
ʗ ʘʙ
ױװᱎǝᱎ ཥ ࢥࢦ̆
ǵ קՇሤ ɾࠧæȦ ƁƂƃ ĺ ̉ƄʖʗЩʙז
ĉሤདྷሤᇝǻႣሤ ༙ ᔃ ၻᘡᘢቴᇰᱎ ඵ ၼᘣᇰ ᐪ ᘤᱎ ᢕ ᐩ ᫍ ᱎ
ǿŮࢷ ࡚ Ⴄሤ ᄝඊໄሤ
๑దÿሤ ֭ࡷࡄሤ Ą ࣊ ߵƇሤ
ရᱎ ¿ᱎ ࢉࢉሤ
๓ ᄠሤ \ݐƉ֗ņሤ
ȥ ᔇ ႁ ᘫᇲ ᐱ ᘬ፵ᱎ ᢗ ᐰ ᤃ ቸ ᱎ
Ðז
ͥƈ юτ҃ ᄆሤ ƐҤᅯሤ
ᄢ֦ଶஶሤ
߸ Ċ ܗމႥ ሤ ݆߷ ሤ
íऻᱎ
Ǽǽሤ
དྷപሤ
Տ߷ དനሤ ঘሤ
ଳᇗሤ
๖൨ ᄤሤ Ն Ǵሤ
Ւ ՙ ՓՃ߷
ՊǸሤ Ă ࡴ ࢌሤ
üሤ ֜ą߷
˙s חሤ ԤԥG߷ ߷ ؞ʖ ܱܲሤ G )* ߷
Ⴁ ௸హሤ
ᇖ၍ሤ֛߷߷ַل
ᄘሤР ࢆ ᆼ ሤ ڃРڰజ
၏ᇟƊሤ
ѣ߷ ֶ߷ ߔ߷
Qٜ֭߷
Ց՚Ծ߷
Ðԇஏᱎ ᤄႂ༼ᱎ
ĀՉሤ ރሤ ߷ࡶ࣋ᇜሤ
ņ ᯇ༺ᱎ Ոഩሤ
Ć य ƈሤ ੭ોćŬሤ ӣՌז
ߑ߷ ၀τ˟ˠ ˡ À ײÁ' ׳ᱎ Ӯˢ
Ր ߷ ْ߷
ࡆ࡛ሤċ ފሤ ࡇሤ ৪ࢶሤ
མಟ ో ೲ ᱎ ࢼ ࢼ ߶߶ƪ߶߶߶ᱎ
ؒٙӦӧ߷
*, , ֨ծ߷ ջ߷ذدሤ֚߷ زرሤਖ਼ න ௺ ሤ
ےۑᱎ & '($,
ռ ىγδ߷ , %, +"#,
) !,
Î Ïᱎ
༂ໞྒྷ༂࿀ཀᱎ ෦ໟྔംཀ࿕ౙᱎ ෨ᱎ೦ᱎ ັംྒྷംཁᰪ
າംཁౙྒྷ༂ཁᱎ ༡ཀ༂๖ംྒྷ෦༂ໟᱎ
ߔ࠻ࠃćī ߭ćࠑ
ࠋī ˶ ˷ ءõזՓ VV Ĉז ćז
ĸ߷ Ҿ߷ճԬ߷ Á߷ҿ߷ե߷ ۗۘߛߜÊ߷Ѵѵ߷ पሤ ߷ٱ ߷ߗۓ ݅߷ѝķ߷Ҽ߷ߐ߷ Ąߒ߷ ௗᱎ ʔ߷ے
߯ሤ ཨጼᱎ ഀ෧ྕ෩༁ໝᱎ ഁᱎ ߴ ࣉࢣ ࢮ ሤ ציᱎ Č߷Ӣӣ߷ߝʣѶۙߞε߷
߷ ߷ݍ קᱎ ݇߷ҽў߷դ߷רᱎ ʚ © M مՠ߷ת שشسᱎ ሤ
֔˲٠
́̅̄ ᇏ̂̃̄ᇐࡃࡌሤ
߷طضʓ
Կሤ ˏ ː ˑˏ ˒ ᯄᱎ ɡɢ˯˰ ப˳ǷځڂƋԾሤ
ࡘ ó ߭ Ǯ ሤ ࠽ޕ ាᱎ ä O ญ ሤ
ેǭሤ
ࡲ Ղ Ş ö÷ሤ
î ሤ ྐധሤ ਸ਼ᯅߵô˓˔ᯆ
Mãሤ
ᄙ ïଯ Ɓሤ ˕˖˗Ӥӥ̆ېۏᱎ
ລሤ ࣅݼሤ §߷ۏ ሤ ᇉᇊ༜ሤ
Ũࣆ ߪݿՁሤ
ñሤ
๐ຟ
ũණòƂሤ
ࡕܔܓሤ
é ê ޘëሤ
ì Ƒ ሤ
ᄥ ࡸ ౝሤ ཏඬ னபƃሤ
࣌߹ሤ
ਇሤ ࣄሤ
ే ұҲᚸҳ၏ ဿᱎ
í ࡗȱሤ ƀሤ
߷֦ع
ߍߎ
Դ
ǯ Ճǰሤ
ႚ DZ ड़ ŒDzሤ
ᄚᄛᄜሤ ſሤ
Qøሤ ૅሤ
Ŧሤ
ݏሤ
ሤ
åଭሤ
ࡖ ݓæሤ
ࢠ௲ሤ
ç ᇋநௗሤ
Öז Pèሤ
ଲሤ
ߨ࠾ࢇݽሤ
þሤ
ሤ ཎሤ
Խሏሐሤ ̈́ ߫ ߫ ࡰ߬ሤ
ĸԆ²ז ݀ћ߷
ࡀ ࡱ ô ࡁ ࡱሤ
Ƭʣ ࡂ ࡱ õ ࡘࢡሤ
བྷ ಞ ొ ೱ ᱎ ࢷ ߳߳ ࢸ מƤ ߳߳ߴᱎ
ອ༻ಡౌྈ༻ᱎ ༚༻็ಢྉශ໒ᱎ
lሤఖ ᇚ ᄠ ūሤ ݃ ֖ ݄ ߷ È߷
ᇑੋ࿊ ࿋ཐ௴ሤ צఝሤ
ۊ ۉ ۈÿ߷
ؼᬞ ½ ¾ ᯃ ᫈ᱎ
! т 'р . с߷ ֥߷!"DEZ
Ѭ߷ ѭ߷
Ӏ ߷ӁӂE
Ѫ߷Qѫ
߷:; <=Z ѯԩѧ Q ʕ ӞӟE ѮӠӡ # Ѩ ѩE Ӱӱ
қ߷Ҷ э ю ߷ ġ ߈ ߷/ $ #%0Z߆߷шщ Ӯӯу ұ ф х ц ч߷Ӓӓ߷
iʣ
ý߷ёђяѐ߷ ʐʑ߉߷
& ' ( 1 ZQRS! 2 Ҳ
Z ъ ыҳ ӔӕҴ߷ٖ߷ҵ߷
ܹ߷м Ķ ߷
߷ ߷
ܻ ܼ ߄߷
֢߷ܺ߷ 3 4Z߇ ь ߷
Z оп ! ӎӏн߷߅߷Ӑӑ߷
ҷ Ҹ ҹ Һ߷ʒ
Ę߷-ZĀā ܽ·ێĂ߷ є ߋ ѕ Ł ߷
ˊˋ᫉ˌ᫈ˍ᫊᫋ᫌ ˎ 89@AZ ߷NOP BCZ
6I
" &&I 6I
Z
" &&I 6I
ᱎ
" &'I W"߷
6I
" &&
Z
©Ñ
FGZ
ᎋᱎ
֪߷ ධᱎ כᱎ )Z
κ TU %D"%6EI ᝡ឵ᚵᝢᔂᢓᚶᘞ ᱎታᘟ፲ᤀᘟታᱎໃሤ
߷
ධᱎ
ƩȒ
7߷
5Z
නᱎ
ᢑᱎ Ɛ
ˇ ᯂˈ ˉyᱎ٩ᱎ ®߷
݆݇ሤ
߷
ݷሤ
्ᱎ
ז
ௐሤ "
ᯁᱎ
ሤ ̭̮ߊ̯̰̣̤̅
ॶሤ Z
S߷
ဉሤ
6I Z ೯ᱎ ˆ ਵঈᱎᢔታᆛᚷᘠᇯᢓᱎᚶጡᱎ $FI4+DD+7(I
ࠊᱎ ͟
" &&I 6I ̈́ז
ੈሤ
֯ז Z
݈݉ሤ
" &&I 6I ᫈ეᱎ ػغᱎ
݄݅ሤ ࠽࠾ᱎ ᱎ
" &&I 6I
Ñ
ᯁᱎ
" '&IX Y ೖሤ " "
" ᎌᱎ +>?Z
7߷ ᯀᱎ
Z
Z
ইᱎ ֫߷
8߷
Z Z
É߷
ᢒᱎ ቲᱎ ࠊᱎ
߷
ᰨᱎ
ݹሤ
ࠩሤ Ԅז ҂ז
२३ሤ Z " ്ൎሤ
ဈሤ Z T
6Z " 8߷
ĽȒ
ሤ 7Z
^ x
6I ႗ሤ
½ÇÇ ӞӟӠ½
ᱎ ሤ
6I ሤ ോൌሤ ďщы҃
ਘᱎ &&I 8I ಐԼሤةب՜VZ
߷
ൊሤ ˅ંᱎ
Ċᱎ &&I "
6I Z Z
tȇ ;&&
I
" ੇሤ
Z
8߷
JKZ
Z
ݷሤ ుሤ
LMZ
Z ઁᱎ ইᱎ Z
Z
Z S߷
ఖ ᢒᱎ
®߷
ז
ᜟᱎ "
Ɛ *
ඡᱎ
٫߷
" ᇇሤ
֩߷
ඡᱎ
ъ҃
ᱎ
߷
7߷
ৠ
7Z
6I ॶሤ
" &&I 6I বሤ " 7Z
ᇇሤ
" &&I
9I 7Z
ৠ
ৄᱎ ;&&
I
CI "
" ''I
ඡᱎ
"
"
"
" "
Ȯʲ Ȩʻ҃ Ͳ͝
Ϯ ԅז ĕȒ
x¬ <¼
ɽɾ-??- ɻ ? ɿ ߃ɼ - ɴɵɳ
ᇆͫ лʀ rΩΪΫάʁҠʂ ᱎ ԣ߁̃ɺ߂߷
ێۍԢ
4 njƊƁƽƸƆȇ 6zȇddȇ ƳƟǍfW§ȇ ö 4ȇ LƂ6ȇ ᚴጠᱎƈǐȇ .fLLUWȇ
ōddȇ W6fLºȇ
R ႛ ႚ ùúႜႝ႞Մ႟ሤ
Ϛϛ
@@ʈ
ɰ ̫̬ɱɲ ̴̵̶
ΑΒ̖ʜ҃ ͳ͜ ϭͮ זʝ҃
¼<
-G#*I).5 2I?D+E/<6I
# 4 ᢏቱᆚᚳᘝᇮᢐᱎᱎ$GI 5+DD,:)I
ુሤ
ඡᱎ
ඦᱎ
Ɯ ૅ =33I AG%I
ಳವಶಷಸಹಥሤۨ۩۪ۭ۫۬ۂہۀ߷ۮሤk5W l5\234m] ^ n
RSD,oXKLUAH !p EZ_6`01#$7-[.a8bQ =>?+" T9 :OF;%Y@GJP<cCdqI/r*()s
B Mt&'Nefghiju
I <32I)05!3I@D.E+>:I
˯ ˰ ˱˲
1)HB%I$
MW\WQc Z©QzmX Z¥meZ\;]
ͯ ϲυ ͪϳ ЅЁܢ ז
˺
ࠞ ߬ ۉଙ ˻ ʍז
ૢሤ
ȗȘ șȚȑ ߷ۀȒȞȟȠjț ȓȔ j ȕ ȜȝkkȡȖ ˚ Ȥ
ȵ ȶ ȷȸȹ ˚˚ Ⱥ ˛ =, ȳ
¬" " Ȩ «ȴ "«ȩ Ȫ ȫ Ȭ ȭȱȮ Ȳ " ¬"
Ȱ mȯ m
ીሤ
ȏߴȍ Ȏ ˜˝ ͻ
ȏ
;I)Eb"BCE="$,b B,#G)b ) 3)%Kb "JK.JL')b RO/'b 5A)b Λז ˪ ੋ܋ଘ ऴ
ࢭᱎ %>6BLJ)Fb BF>+F# 7b ᱎ ᱎ %>8
ᱎ
ᱎ ᱎ
ᱎ $ ඳᱎ
Êȇ cȇ
לᱎ
ߢሤ ް ޱȌ ٬ ߷ ц֭֮ хфут ֬
˂˃
ᱎ
ࢶᱎ
ᱎ /$ʬͣ ҃ᱎࢶᱎ
0J%- #JJ1JM()b) E E&Fb
͓͔͕͖͐͑͒ ߤಌזಎಯರಏЄЅᇃ
ІЇ ᇄ΅Ά ᱎ
рɯ Țпо ș Լזҁז
ہȻȼخحԝ
د Ԟԟ
ΨȽ ߷ائ ޯ߷ üۂɎҰɗɓɔɕɖp ̛̜ p oɏ o ɍ ɐ ɑ ɒ
ᮾᱎ VWXYZ[
ল
PS\ с ˤ˥ ^!_T`
a
U
D 0
ģሤ Ĥሤ
] ૣሤ
ૢሤ
˞ ˟ˠ ˦˧
,Ɉɉ,,Ɍ̙̚ й߷
=
ࢮᱎ ඬᱎ ܸ߷
κλȋ
ȉȊ
ٵיט ˡ Ϳ ז
к߷
ીሤ ᒳᱎ
̧ז
߷؛״׳ ߷ת ʥ҃ ߷ש߷ײױװׯ
ʸʹʾʿʽᮿ ʺʻʼ
Ԓחᱎ NJ?9"J0%bJ - F>JJ4)b ͉ז ᇂࢄ
ѱ/$ ʭͤ҃ Í
ඳᱎ
ˢ
ඳᱎ
ᐥᱎ âሤ Μτ ҧϱЄЀ זȱȫ҃
˻Ŷ ˩.ז
ǽ
Ǿ
ǿ
Ȁ
hȁȂ
ȃ
ǺǻǼ Ȅ
ȅ
Ȇ + +
+ +
< <
i ȇ
Ȉ
;
; *
*
ǵޮ߷
зǶ
-jә֦͇͈͆זJHк 0Y 0Gй҃λ] ז
˽
ૂሤ
ઽሤ
ાሤ
ಋઽሤ
*
]
H17j XW2X^ %j&MMGMj :j ʩז
H>;j)/@;jKGS/X3GBj ̦ז ȋ£
߷
]
˭ ˫
߷
ͺ˒˓
॓ bjP X(j
eXGf[4g\-hW]i Ĉʺ ઽሤ
f g̖̕ǹǸgǷ
ඩᱎ *
D ;!';D 9%': D
0! :D 1&1 D
D = H = H ")%}
}
K}
ٚܶ666ܷ6߷ K K L L K K LKÊ
} И ЙК K ÊKLK Ê
"}
௴ᱎ
ሤ
E pioS c bMgt}Mt} bMgRBY}
সሤ
ሤ }
ൃሤ J rMObS}opioSbbMgt}
ဲሤ
ޒሤ pSeM ] g ] g Y } "(&} bO}
m҃
ৢሤ
? ^ p ] g Y }t ] eS}
ᆘᱎ
֒֓ሤ pSeM ] g _ g Y } }rSP}
།ᱎ &}
ޓሤ
ݶሤ
܉ሤ
ഷസሤ
}
߮ᱎ
M} > SrPTgt} opio v b r ] j g }rzrtSe}
}
@MY^ gY}rzrtSe}
ࢪ࠙ᱎ
௴ᱎ &}
҃
кLሤ
ሤ
҃
ൄሤ
ޓሤ
ৣሤ
হሤ "}
ᆘᱎ
֒֓ሤ
ޔሤ
ݶሤ
܊ሤ
ഹഺሤ
}
g{} ;zrtSe} iySpr}[iit} E$} o[MrS}
}
Ä[Ê = oopnP[}
઼ሤ
opiYpMe}
}
ɭ
W p}
I^eS} [p/e]g}
? ^ Y v pS}
E p ioS b bMgt}Qlg r x eot]ig}
5 - 33
NASA-S -69-3722
I l11
Tl 1/ \
I
u
�
g
"'
'"
/t--J
�
.c
u
,..._ I .....-.. 1-. "" I "-=
I _k-Vf I I
""
o._
I\.
tttD+++++t-t±=ttd�
'"
= 8
�
..,
"'
c
ro
.0
ll /
-;;; 4
�
-4
E
·o;,
.c
�
o._ 0
� X
J t.anding
I
�
�
ar
"' ·
.:>:"
E
1 11l!i1+t MtH 1\ I I n
h
VI .....
r---
-4 1 I I I I I I I I I I I \!1; 1
0
1 1 1 1 1
"'
"'
'"
"'
4 8 DIP-+--H-t-+++H--tt-tttr I I
..,-
"'
c
ro :J/1 11
v
0
'K:
ro
E
.0
·o;,
_7
g -4 4 :1 I I I I I I I 1
� I
� I
� -
'\. 1 I I__._ 1.... I I I
��
[
�- o
l I I I I
�
"
.
�
I "'-/
1 I I I I I 1
4 T I T '
ar I
I
"'
r
6
�
"' 12
11 N l o l l a l l o-!
I
" 102:45:34 102:45:36 102:45:38 102:45:40 102:45:42 102:45:44
;'.
Time, hr: min:sec
't-
NASA-5-69-3 7 2 4
v
/
/
v I
'
/ v
I I
I\
80 1000
<> /
I
QJ
Range
"' "'
1 J
.!!!
d
E
Qj
"'
76
Qj
"' 0 \ v v
I
�
�
!-- Range
<::
"'
0::
QJ
"'
<::
1\ rate L
I
\ 1/
"'
J
0::
72 -1000
\ 1/ /
\ 1/ 1/
/� I
'\ /
v 1\.. /
I/ "'
...... v
.....
60 -4000
1 2 2 : 2 1:40 : 2 2 : 0 0 :20 :40 :23:00 :20 :40 :24:00
Time , hr:m in:sec
·
Figure 5- 15.- Rendezvous radar tracking of the command
module wh i le LM was on lunar surface .
NA SA -5 -69-3725
-- -
90
- - - - P lan ned -
Actual
80
5600 70
�
ll'�
"' �
4800 60
Q) / F l ight path angle
v
"0
L.
Q)- .�K
"' �....
4000 <7> 50
c:
f
-"'
['-Velocity
J::
<..>
Q)
Vl
c.
40
v
:E
3200
� v
.SO' /
u..
li
<..>
\ v
v
E. 2400 30
Q)
> \ ../
1\ v
"
20
v
1600
v
'J"-...k'v
r-1--t-
L.-<v / Flight path a n g le
800 10
0 0
_v v !'-Velocity �
124:22 124:23 124:24 124:25 124:26 124:27 124:28 124:29 124:30 124:31
Time, hr: m i n
g
g
બᱎ ʏ
ᱎ
Ă Ӡӡ
Z
&gM - \g4;>g
ᱎ Ǫሤ
g
ҋ˙˘ז
gÑÊIJÊ
Վ߷
ߡሤ
টᱎ ࢦᱎ
g
g řࠊ &8\ga7gᱎ
ʌʍʌ 1 L7 S S >;g
g g
ɠ҃
" g #g # g $g $g f%g % g %!g " g " g
3 D R> g B X $ R E S g
Щ҃
ǃ
ᚰᱎ
$ ßזভ
Нזᱎ ంГ
Ȗʣȹ
;ʣĠʣġʣா ࡴ
Hሤ ୷ᱎ ୶ᱎ ᫂ޠᱎ
Ƿgʣ
Ê
ሤ ڡڠڟሤ
ݳሤ Əˑ ࢤᱎ ఃᱎ i 3È}Ò ·xÈÒ
ҟҠ
ҡ धሤ ᱎ ǧሤ ᄗሤ ඔᱎ n׆ᱎ
Ⱥƃʣˬ҃ ᫂ᱎȌ҃
$ ᱎ
ృሤ
ሤ
ڼڻںሤ
ਯሤ
ਰሤ /Ò
ᮝᱎ
Ш҃ ࢥ ᫃ ᱎ ĩʣ љז
ፗᱎ
ʣ ᬛᱎ
ᱎ ሤ
ॲሤ
$ ࿒ᱎ ᫂ᱎ
ڛښڙሤ ഻഼ሤ ᆷ Ⴊմሤ i È}ÒxÈÈÈ Ê~Ò··© ³Ò
Z_k Ò
ݲሤ
ٲٱሤ ሤ
*+#Ò မᱎ
"Ò ¸ i È{ xÈÈÈË~ ··©³
_ g d Ò
ٲٱሤ ~ሤ :CÒ
˒ ДǴ ః
JÒ /Ò ڀ
w h¤Ò
A0Ò hÒ s Ȉ
h ¤Ò ү ӌӍ߷ פ ¨ ʿ ¨ ھ ˀ ʉ ˁ
e e
º
@[Ò
:0Ò
[Ò
hÒ t
hÒ
h Ò
`
5 :Ò
ሤ
ܴܳሤ
഻഼ሤ
kሤ
ሤ
ሤ
ᢍᱎ /Ò ɢĭ ´ʣ
ষሤ ڧڦڥሤ
ᱎ
ݳሤ
ڞڝڜሤ -
Əː
ሤ
ڞڝڜሤ ڞڝڜሤ
ڳڲڱሤ ਨሤ
ݴሤ ሤ
ڶڵڴሤ ሤ
ҥҦ
ҧ жሤ ሤ
AjÒ hÒ
e h¤Ò ਧሤ
ुूሤ
иሤ
/Ò ఀ
ሤ
য়
Q҃ ߷ڽ pzÑKÒxÈÈ ÈË~Ò ¶ ·©·Ò
Z_ k dÒ
॰ሤ ሤ
@jÒ ˋ
ࡡሤ :CÒ
ሤ
֬ hÒ
ࡠሤ
য়
ř৶ hÒ ఁ C/Ò
7 :AL ::O1Ò 4:AL::M@/Ò 7:AL:@NÒ
4:Ò
qxÒ 7:AP ::Q//ÒÉ©Ò 7:AR:CP@/"Ò
^ ʹÒC 7 E $ k±||¶xÉÒ~Фx }ÁÒ~ Ê ¶ ¤ÒxÂ{¤È"Ò
̀
́
!
́
!
͂ 3 3Ⴕሤ
! !
̓Ⴗሤ
!
̓
ˊ
)
) )
3
3
́
́
DZ
ˋ
:
́
Dz :
d
:
͂
ddz
)
ƀ҃
Ɓ
Ǵ
Ջ߷ Ռ߷ Ս߷ ʆ жެ߷
ǃ҃ûƂƃ҃ ]҃ǂ҃¤ƅ҃ Ԍ
ᚯឭᱎ s< @ H ` }
ឭᱎ @ D a}
55} **} 6+}
྅ ᐘ ᖽቬׅᱎ Ꭵ ឮ୴ ᖾ ᐙ ᘗ ୵ ᢌቭᆗᱎ
ǯ де߷ ǰ ߚ߷ ʛ M߷ ʜ ɺ ɺ ɺ
ɼɽ ɾɿʀɻʁɿɼɻʂɽ ʃɻɼɻ న ప ᮗᱎ ே ԏ Ԉ ԉ Ԋ ᱎ
ӱ ᱎ٨ᱎ
Ӳ Ӻӻ
ᱎ Ӽӽ ӳ ӷ ை ᮘᮙ ᫁ ʄ ׄʅ ఫ ᮚᱎ ᮖధᱎ
ʟ ɻ # t # # # $ ʠ $ ᮓ $ $ ʡ ʢ $
ɻ ɻ
ʝ # ʞ t ᮔ ᮕ ߷
ᐗᱎ# -
- !- !"-
ȑ ͋Ǜˌז ¶ ڮ
Ժ߷
זRSᪿᱎΎז
୬ᱎ ߷
߷ ֝ሤ
ᪿƚᱎ ુ ᬙᱎ % °Ȓ
߷ ݎ
߷
ࢽஎᱎ ᐫ É
߷
Ԫ ԫ ǂìȒ
Ȑ ז᠏ ୭ ୧ ᱎ µ ᱎ
ᱎ
Ջ זɞ ef - VŘ ᬚᱎ ᫀƛ Ǔǔᱎ ᪿ׀ᱎ ᪿᱎ ֿᱎ Ǘǘᱎ
ሤ ÞԳሤ 8Ř
ᫀ׀ᱎ ᱎ Ř
ᰰᱎ Íז ᦒᱎԼԽ߷
ܳ߷
Զ ڬ ז-Ř҃Ɠᐓᱎᱎ ҃
୨ ୩ ᱎӢӣᐕᱎ 3
᪾ఁᱎ Ûሤ 3Ř ׂᱎ ³òᱎ ୪ᱎ ᱎ ڟڞᱎ
Ɨ୮ᱎ ဘᱎ
߷
3ŋUŘ Ćᱎ ćᱎ ᆶሤ
߷
ǀ҃ ӡ՞זȇ پƴࡢ Ό҃ Ȝ᫂ơᱎ eƦåʣƔᱎЙזԜ זƴȒ҃ˡ҃
ுᱎ
ƾ҃ǁ҃ Ƴ҃÷ᱎ ৗ ͉ࡡ ֊߷
ȝ
߷
lᱎ ڭ Өᱎ ߞሤ ᪽ݳᱎ ౪ሤ အ
Ǩ ' ު ˴ ǩ ǭ ' ( ' ` a ( ' ( b ` Ǫ ǫ b Ǯ Ǭ a c ( c ٳ
߷ ߷ D߷ в߷ Ү߷ D߷ Ǟ D߷ ǟ Ǡ & г߷ & & ߷ڻں ǡ Ǣ ǣ ɱ & ԭሤ ɲ ɳ ɱ ԭ Ԯ ሤԯሤ ɴ ɵ ɶ ɷ ɸו ɹሤ Աሤ Բ
Ǥ \ ] ^ 9 ǥ ^ _ ] \ ڼǦ _ 9 ǧ 9
ܴ߷ ֤߷ ߷ ީ߷ Ի߷
7 2} ."8} 9"8} 8 %8}
D=H=}H}
-
ɚ ɯɛ ᮒᱎ ɝɠɞɟᮍ ᮎᱎӔ ᮏ ᮐɡᮑ ɢɦɣɤɤΙּ ɥ
&-
'
-
(
!-
'-
I ^ fS} [ q : f ] h : rUQ}
ēᅱģᱎ ᩏ
ഓ ᐒ ፭ ᧄ ឬቫ
ָ ˅ٟ
+ ) , * ,
-
Ǜǜֵᱎᱎ
ě
'[9>T_g WXVW b Q[ HUSg [e[_>Rg9a`U@@g ¦ȁʣ
ָֹᱎ ֶַᱎ
ᱎ
$
%- 8:BS:GT@/Ò 8 :BUÒ 9 ;BV>V<Ò
Ēᅰᓹᱎ 8:BW:CX@/Ò_Ug 8:BU ?Y=(Ò
, H Ab X>g C8 F ) )US 9 Q c;>;g
໓ ్ ཙ్̓̔ઢତ̓ৰঐତᱎ ౬ሤ
રሤ
ļᱎ
ߩᱎ
ॾ੩ᱎ
ī҃
߷
^ᱎ
ֱ߷ վտ .7 ॾߪᱎ 1--7 #1-7 (/7 )/7
±ɒ #7 (1-7 )&7V */7
_ᱎ
^ᱎ බᱎ ז
R߷
ඵᱎ
ॽᱎ ߫ᱎ V
ᮈᱎ
Ԯԯ߷ භඹᱎ
ಌᱎ ඵ ඵᱎ
ᮇᱎ ऐऑ /7 ১߬ᱎ 2--7 #1-7 (.7 )/7
͟͠Տ
" 7 ҃ ᱎ
V "7 ! $3-7 ) ' 7 )/7
/ 7 ਯ߬ᱎ
7
ඐᱎ
Яа
#7 ז
ɕॿ੪ᱎ
. 7 ১߬ᱎ
7 #
ݚ ˇז #
࢚
#7 ඵᱎ
.7 ॿ߬ᱎ
7 ॿ੪ᱎ ࢹ ࢺᱎ યሤ
ඵᱎ
࢙
#7 ඵ ඵᱎ ඵᱎ
. 7 ࢢ߬ᱎ
7
߷
ࢹ ࢻᱎ
`ᱎ ඵ මᱎ
ǓǔǕז
ֲ߷
ոչխ߷ ඵ ඵᱎ
`ᱎ
R߷
ॽᱎ
߫ᱎ ɠɡʣ
ᮈᱎ
ġ
ඒᱎ
ޜޛሤ
ɓॾ੩ᱎ
ࢢ ॿ ਯୡॿ১ୢ߫੪ᱎ ࢢ ॿ ਯୣ ॿ১ ࢢ ߫ ᱎ ࢢ ॿਯ ॿ২ ࢣ ੫ ᱎ
! 7 )!"-7
4+7 ੪ɔࢢ ִ ɔ 5(#7 . 7 -.')67 '7- ()07 ,
.-7 1)%7 -%/7
5-41
N A SA-S-69-3 7 3 0
10 - Sun
�
Earth
Event T i me
1 L i ft-off 1 2 4 : 2 2 :00 . 8
9 Begin stationkeep i ng 1 27 : 5 2 :0 5 . 3
10 Docking 1 28 :03 :0 0 . 0
10 I
-
--
[ Nominal trajectory
1.!
�� Nominal k 36 = 109 . 8
""' K
......
Constant d ifferential
"-..,. "'-.
'-
height maneuver
....... .......
ro....
.......
!'---
ll' ....1--.
...
k36 = Range rate 36 minutes Computed trajectory � !'--.... ....._
40 prior to constant differential
height maneuver
I I I I I I
50
70 80 90 100 110 12 0 130 140 150 160 170 18 0
Displacement, mi
( Lunar module behind)
6.0 COMMUNICATIONS
Two-way phase lock with the command module S-band equipment was
maintained by the Merritt Island , Grand Bahama Island , Bermuda , and USNS
Vanguard stations through orbital insertion , except during S-IC /S-II
staging , interst age j ettison , and station-t o-station handovers . A com
plete loss of uplink lock and command capability was encountered between
6 and 6-1/2 minutes after earth li ft-off because the operator of the
ground transmitter at the Grand Bahama Island station terminate d trans
mis s i on 30 seconds early . Full S-band communi cations capability was re
stored at the s cheduled handover time when the Bermuda station established
two-way phase lock . During the Merritt I sland station ' s coverage of the
launch phas e , PM and FM receivers were us ed to demodulate the received
telemetry dat a . ( Normally , only the PM data link is used . ) The purpose
of this configuration was to provide additional data on the pos s ibility
of improving telemetry coverage during S-IC/S-II st aging and interstage
j ettison using the FM receive r . There was no los s o f data through the
FM receiver at staging . On the other hand , the s ame event cause d a 9-
second loss of data at the PM receiver output ( see fig . 6-1 ) . However ,
the loss of dat a at interstage j etti s on was approximately the s ame for
b oth types o f receivers .
The USNS Reds tone and Mercury ships and the Hawai i stat i on provi ded
adequat e coverage of translun ar i nj e ct i on . A late handover of the com
mand module and ins trument unit uplinks from the Redstone to the Mercury
and an early handover of both uplinks from the Mercury to Hawaii were
performed because of command computer problems at the Mercury . Approxi
mat e ly 58 s econds of command module dat a were los t during thes e handovers .
The loss of dat a during the handover from the Mercury to Hawai i was caus ed
by terrai n ob struct i ons .
After landing , the lunar module steerab le antenna was swi tched to
the slew ( manual ) mode and was used for all communi cat i ons during the
lunar surface stay . Als o , the Network was configured to relay voice
communicat i ons between the two space craft .
have supported the lunar surface acti viti es without deployment of the
erectable antenna with s li ghtly degraded dat a .
No frame
synchron i zation
-o 20 r----r--_,----+----+orn��--�..-+----+---�-- �----+---�
(/) 1::
.... 0
0
�
u
� 10 r----r--_,--+-
"'"' ....
co 2i
FM T E LEMETRY PERFORMANCE
0\
F i gure 6 - 1 . - Communicat i ons system performance (down I ink) during launch . I
V1
6-6
-60
A - Steerable antenna automatic mode
S = Steerable antenna slew (manual) mode
� r{; ce-up maneuver
-70
J
"
E
I A A
[ i ! I !
co
I
-80
"'
c.
1
2 1 0- oot antenna
I
l i VV\ �'-"""�....-ti'ii¥-i"'"'-"i'-
� �
"""'
I I
-100
lil 1 I I 85-foot antfnna
Jl
I I 1
"'"
�
I I
U i lf\' I
I ' �1-w\... �
>
0::
1
-120 1-- · -
�·-
...
7 - : :-+_ "- ::-"" _. -
0 -percent word mte 1 1 191 b1 l tty
-130
u
-140 [ DOWN L I N K POWER
No frame
synchronization
" �2 o �ty�--����r+--;n----���---w�rr---t--�-�+---+---�
m =
� 0
g
"' m
g
��
No frame
synchronization
"
m c:
� 0
g
"' m
g
05 �
Time , hr:min
F igure 6-2 . - Communications system (down link) performance during final descent.
7-l
7.0 TRAJECTORY
The earth an d moon models us ed for the traj e ctory analysi s are geo
metri c ally described as follows : ( l ) the e arth model is a modi fi e d
seventh-order expan s i on containing ge odet i c an d gravit ati onal constants
repre sentat ive of the Fi s cher ellips oid , and ( 2 ) the moon model i s a
spheri c al harmoni c expan s i on containing the R2 potent i al funct i on , whi ch
i s de fined in reference 2 . Table 7-I defines the traj e ctory and maneu
ver parameters .
The launch traj e ctory was e s s enti ally nominal and was ne arly i dent i
cal to that of Apollo 10 . A maximum dynami c pres sure o f 735 lb / ft 2 was
experi enced. The S-IC center and outboard engines and the S-IVB engine
cut off within l s e c ond of the planned time s , and S-II outboard engine
cutoff was 3 s econds e arly . At S-IVB cutoff , the alt i tude was high by
9 100 feet , the velo city was low by 6 . 0 ft /sec , and the flight-path angle
was high by 0 . 0 1 degre e all of which were within the expected di spersions .
The p arameters derived from the best estimat e d traj e ctory for e ach
spacecraft maneuver executed during the trans lunar , lunar orbit , and
transearth coast phas e s are presented i n t ab le 7-I I . Tab le s 7-III and
7-IV pres ent the re spective peri cynthion and free-return conditions after
e ach trans lunar maneuver . The free-return results indi c ate conditions at
entry interface produced by e ach maneuver , as suming no additi onal orbit
perturbat i ons . Tab le s 7-V and 7-VI present the respe ctive maneuver sum
maries for the lun ar orbit and the trans earth coast phas e s .
The pericynthion altitude result ing from translunar inject ion was
896 . 3 mi le s , as compared with the pre flight predi ct i on of 718 . 9 miles .
This altitude difference i s repres ent at ive of a 1 . 6 ft /sec accuracy in
the inj e ct i on maneuve r . The as soc i at e d free-return conditi ons show an
e arth capture of the space craft .
The command and s ervi ce modules s eparated from the S-IVB and suc
cess fully completed the transpositi on and docking sequence . The space
craft were e j e ct e d from the S-IVB at 4 hours 17 minutes . The e ffect of
the 0 . 7-ft /s e c e j e ct i on maneuver was a change in the predi cte d peri cyn
thion alt itude to 827 . 2 miles . The s eparat i on maneuver performed by the
s ervi ce propuls ion system was execut e d pre c i s ely and on time . The re
sult ing t raj e ctory conditions indi cate a peri cynthion altitude re ducti on
to 180 . 0 miles , as compare d to the planned value of 167 . 7 mi les . The
difference indi cates a 0 . 24-ft /s e c executi on error .
The computed midcours e corre cti on for the first option point was
only 17 . 1 ft /se c . A real-time de cision was therefore made t o de lay the
fi rst midcourse corre ction until the s e cond opti on point at trans lunar
inj e ction plus 24 hours becaus e of the small increase to only 21 . 2 ft / s e c
i n the corrective velocity require d. The first and only trans lunar mi d
cours e corre cti on was init i at e d on time and result e d in a pericynthion
altitude of 61 . 5 miles , as compared with the des ire d value of 60 . 0 mi le s .
Two other opportunities for midcours e corre ction were availab le during
the translunar phase , but the velocity changes require d to s at i s fy plan
ned peri cynthion altitude and nodal positi on targets were well below the
7-3
The lunar orb it ins ert i on and circulari zat i on t argeting philos ophy
for Apollo ll differed from that of Apollo 10 i n two way s . Firs t , t ar
geting for landing site lat itude was b i as ed to account for the orbit
pl ane regres s ion observed in Apollo 10; and s e condly , the circulari zation
maneuver was targeted for a noncircular orbit of 6 5 . 7 by 53 . 7 miles , as
compared with the 60-mi le-circular orb it t argeted for Apollo 10. A di s
cus sion of these considerat i ons is presented in s e ct i on 7 . 7 . The repre
sentat ive ground t rack of the space craft during the lunar orbit phase o f
the mi s sion i s shown in figure 7-2 .
The s equence of events for lunar orbit ins ert ion was initiat e d on
time , and the orbit achieved was 169 . 7 by 60. 0 miles . The firing dura
tion was 4 . 5 seconds les s than predi ct e d becaus e of higher than pre
dict ed thrus t ( see s e ct ion 8 . 8 ) .
The lunar module was undocked from the command module at ab out 100
hours during lunar revolut ion 13 . The command and s ervi ce modules then
performed a three-impuls e separat ion sequence , with an actual firing
t ime of 9 seconds and a velo city change of 2 . 7 ft / s e c . As report e d by
the crew , the lunar module t raj e ctory perturb at ions resulting from un
docking and stat i on-keeping were uncompensat e d for in the des cent orbit
insertion maneuver one-half revolut ion lat e r . These errors directly af
fe cte d the lunar module state ve ctor accuracy at the initiation o f pow
ered des cent .
7-4
The des cent orbit ins ertion maneuver was executed at 101-l/2 hours ,
and about 57 minutes later , the powered des cent s e quence began . The
detailed traj ectory analys i s for the lunar module des cent phase is pre
s ented in s e ction 5 . 1 . The trajectory parameters and maneuver results
are presented in tables 7-II and 7-V .
The lunar module as cent stage li fted off the lunar surface at
124 : 22 : 00 . 8 after staying on the surface for 21 hours 36 . 35 minutes .
Lunar orbit insertion and the rendezvous sequence were normal . The
terminal phase was completed by 128 hours . The detailed traj e ctory anal
ysis for as cent and rendezvous is presented in s ections 5 . 6 and 5 . 7 .
Tables 7-I I and 7-V present the traj e ctory parameters and maneuver re
sults for these phases .
The transearth inj ection maneuver was initiated on time and achieved
a velocity change of only 1 . 2 ft /sec les s than planned . This maneuver
exceeded the real-time planned duration by 3 . 4 s econds because of a
s lightly lower-than-expe cted thrust ( see s e ction 8 . 8 ) . The trans earth
inj ection would not have achieved acceptable earth entry conditions . The
resulting perigee altitude s olut i on was 69 . 4 miles , as compared with the
nominal value of 20 . 4 miles .
At the fifth midcourse-correction opt ion point , the first and only
transe arth midcourse correction of 4 . 8 ft /sec was made with the reaction
control system , whi ch corrected the trajectory to the predicted entry
flight-path angle of minus 6 . 51 degrees .
The best est imated trajectory for the command module during entry
was obt ained from a digital postflight reconstructi on . The onboard te
lemetry recorder was inoperative during entry , and s ince the spacecraft
experienced communicat i ons blackout during the first portion of entry ,
7-5
complete telemetry informat ion was not recorde d . A range ins trumenta
tion aircraft re ceived a small amount of dat a s oon afte r the entry inter
face was re ache d and again approximately 4 mi nutes into the entry . Thes e
dat a , combine d with the best estimat e d traj ect ory , produce d the postflight
dat a presented herein. Tab le 7-VII pres ents the actual conditions at
entry interface .
The flight -path angle at entry was 0 . 0 3-degree shallower than pre
di ct e d at the last mi dcours e corre cti on , c ausing a peak load fact or of
6 . 56g , whi ch was slightly higher than planned.
The s ervi ce module entry was recorded on film by aircraft . Thi s film
shows the s ervi ce module entering the earth ' s atmosphere and di sintegra
ting near the command module . Acc ording to preflight predict i ons , th e
s ervi ce module should have skipped out of the earth ' s atmosphere into a
highly elliptical orbit . The Apollo 11 crew ob served the servi ce module
ab out 5 minutes aft er s eparat i on and indi cat e d that its react i on control
thrusters were firing and the module was rot ating . A more complete di s
cus s ion of this anomaly is contained i n s ecti on 16 . 1 . 11 .
The t argeting philos ophy for the lunar orb it ins ertion maneuver di f
fered in two ways from that of Apollo 10 . Firs t , the landing s ite lat i
tude t argeting w as b ias ed in an attempt t o account for the orbit plane
regre s s i on noted in Apollo 10 . During Apollo 10 , the lunar module p as s ed
approximately 5 mi les s outh of the landing site on the low-altitude pas s
following des cent orbit insert i on . The Apollo 11 t arget bias of
minus 0 . 37 degree in latitude was bas ed on the Langley Research Center
13-degree , 13-order lunar gravity mode l . Of all gravi ty models investi
gated , this one came the clos es t to predi cting the orbit inclinat i on and
longitude of as cending node rates ob served from Apollo 10 dat a . During
the lunar landing ph as e in revolution 14 , the lunar module latitude was
0 . 0 78 degree north of the des ire d landing site latitude . A large p art
of thi s error resulte d becaus e the targete d orbit was not achieve d at
lunar orbit ins ertion . The difference b etween the predi cted and actual
values was approximately 0 . 0 5 degree , which repres ent s the predi cti on
error from the 13-degree , 13-order model over 14 revolut i ons . Howeve r ,
7-6
the amount of lunar module plane change required during des cent was re
duced from the 0 . 337 degree that would have been require d for a landing
during Apollo 10 to 0 . 0 7 8 degre e in Apollo 11 by b i asing the lunar orbit
ins ertion t arget ing . A compari s on b etween Apollo 10 and 11 latitude
t arget ing results is presented in t ab le 7-VII I .
The s e cond change from Apollo 10 t argeting was that the circulari z a
tion maneuve r was t argeted for a noncircular orb it of 53 . 7 by 6 5 . 7 mile s .
The R2 lunar potent i al model predi cted this orbit would de cay to a 60-mile
circular orbit at nominal time for rendezvous , thereby conserving as cent
stage propellants . Although the R2 model i s currently the best for pre
dicting in-plane orb it al elements , it cannot predict accurat ely over long
intervals . Figure 7-3 shows that the R2 pre di ct i ons , using the revolu
tion 3 ve ct or , mat che d the ob served altitudes for approximat ely 12 revo
lutions . It should b e not e d that the command and s ervi ce module s epara
t i on maneuver in lunar orbit was t aken into account for b oth the ci rcu
lariz at i on t argeting and the R2 predi ct i on . I f the spacecraft had been
placed into a ne arly circular orbit , as in Apollo 10 , estimat es show that
a degenerated orbit of 5 5 . 7 by 67 . 3 mi les would have result e d by the time
of rendezvous . The velocity penalty at the constant different i al height
maneuver for the Apollo 10 approach would h ave b een at least 23 ft / s e c ,
as compare d to the actual 8 ft /sec result i ng from the executed circular
i z ation t argeting s cheme . A comparis on between Apollo 11 and Apollo 10
circulariz ation result s is presented in t ab le 7-IX .
Seve ral unant i cipat e d problems s everely affe cted n avigat i on accuracy .
First , there was a gre ater incons istency and large r errors in the one-p as s
orbit plane e stimat e s than had b een ob s erved on any previ ous mi s si on
( fi g . 7-4 ) .
7-7
These errors were the result o f a known defi ciency in the R2 lunar
potent ia]_ model. This condition should not o c cur on future m i s s i ons
bec ause di fferent lunar inclinat i on angles will b e flown .
A second problem , clos ely relat e d to the first , was that the two
revolution prop agat ion errors for cros s track , or latitude , errors were
ext remely incons is tent. The average progagat i on error b as ed on five
samples at the end of revolution 10 was 2900 feet ; b ut the uncertainty
in this estimate was plus or minus 9000 feet. On the other hand , the
propagat ion errors for radial and downtrack , or longitude , errors were
within expect ed limits . No adj ustment was made for either latitude or
longitude propagation errors becaus e of the large uncertainty in the case
o f lat itude and the small corre cti on ( 800 feet ) required in the cas e of
longitude .
The third problem area was the large number of traj e ctory perturb a
t ion in revolut ions 11 through 1 3 because of uncouple d att itude m aneuvers ,
such as hot firing tests o f the lunar module thrusters , undocking impuls e ,
station-keeping act ivity , sublim at or operat ion and pos sibly t unnel and
cabin venting . The net effect of thes e perturbat i ons was a s i zeable down
r ange miss .
The coordinat e s us ed for ascent targeting were the best pre flight
estimat e of landing site radius and the onboard-guidance estimate of lat
itude and longitude at touchdown ( corrected for initia]_ state vector errors
from ground tracking ) . The e stimated errors in t argeting coordinates were
a radius 1500 feet less than des ired and a longitude 4400 feet to the wes t .
7- 8
Perigee Minimum alt itude ab ove the oblate earth model , miles
Peri cynth ion Minimum altitude above the moon model , referenced
to Landing S i te 2 , miles
Longitude of the as cending Longitude where the orb it plane crosses the ref
node erence b ody ' s equatorial plane from below , deg
7-9
deg E of l�
Event ve1oci ty, flight.path heading angle,
body hr :min :sec deg deg miles
ft/sec angle, deg
Trans1unar Phase
S·IVB second ignition Earth 2 : 44 : 16 . 2 5 . 038 l72 .55E 105 . 8 25 562 0.02 57.75
:3-I\l'.O :>econd cutoff Earth 2:50:03.2 9.52N 165 .61W 1T 3 . 3 35 567 6.91 59.93
Command module/S-IVB separation Eart h 3 : 17 :0 4 . 6 31.16N 88. 76W 4 110 . 9 24 4 5 6 . 8 46 .2i.J 95.10
Spacecraft/S-IVB separation (ejection} Earth 4:16:59.1 23.18N 67. 70W 3 506.5 1 6 o6o . 8 62.D1 l1J .90
Igni ti o n
First nidccurse ccrrection
109 4'(5 . 3
5 ClO . 0
Earth 26 : 44 : 58. 7 5 .99N U . 16H 5 025.0 77.05 12C .86
Cu-:;off Earth 26:45:01.8 6 . 00![ ll.l7W 109 4n . 2 76.88 120.87
:·bon
�U!:ar orb:.: c ircularization
170. 09E 61.8 477.3
8 0 : 11 : 53 . 5
lg':lit:'.on 80:11 : 3 6 . 8 0 . 023 -0.49 -t.6 . 5 5
Cutoff �·!con 0.023 169 . l6E Gl.6 338. 3 0 .32 -66.77
Sefarat i::m
31. 86E
5 332 . 2
Igniti cr, Moon 100:39 5 2 . 9 0 .99N G2. 7 5 332.7 -0.13 -106.89
C-\:.to:'f Moon 100:40 01.9 1. 05N 31. 41E 62.5 -0.16 ·l06.9C
?moered C.e3;:ent init:'.ation Moon 102 3 3 : 0 5 . 1 . oa; 3 9 . 39E 6.4 5 564 . 9 C.03 -l::l l . . 2 3
�una!' orbit engine cutoff Moon 12)1 : 2 9 : 15 o . 73?1 12 .99E 1C . 0 5 531.9 0.28 - 1:)8 . 15
;;oo r::tir.al fhase finalize /.loon 127 : 46 : 09 . 8 o . &lx 118, 61E 339 . 7 0 .42
.t,.sce:-.t s:agc je:-:;:.son Moon 130:09:31.2 1 , 10N 41. 85E 61.6 335 . 9 0.15 ·97 . 81
-52 . 06
.::.r.al :oeparation
20 .ssw
:g:-. ::.t:.cn Moon 130 :3::l : O l . C 0 .08if 20 . 19W 5 330.1 -0.05
::\.:.:o::: Moon 130 : 30 :08.1 0 , 1 9N 5 326.9 -0 .02 -52.73
135:23 42.3
:Tar.seartf_ �n�ection
:g:-.ition !>bon 0 . 168 164 . 02E 52.4 5 376.0 -0.03 -62.77
:;utoff Noon 135:26 13.1 0 . 50N 154.02E 58.1 8 589.0 5.13 -62.00
-00 . 34
Se;:,ond. n::.d;:,ourse correction
150:29:57.4 13.163 37. 19'� 169 0 8 7 . 2
-8o .41
Igniti�:-� Earth 4 075.0 129.30
Cutoff Earth 1 5 0 : 3 0 : 07 . 4 l3.16S 37 . 83W 169 o8o . 6 4 074 . 0 129 .3C
separati�r,
::: ::.=and :r:odu1e/ser..-i ;:,e n:.odule Earth 194 : 4 9 : 1 2 . '( 35. 09S 122. 54E 1 778. 3 2 9 615 . 5 -35.26 6 9 . 2'(
TABLE 7-III . - TRANSLUNAR MANEUVER SUMMARY
-.1
I
I-'
0
Resultant pericynthion conditions
Velocity
Ignition time , Firing time ,
Maneuver System change ,
hr:min:sec sec Altitude , Velocity , Latitude , Longitud e , Arrival time
ft /sec
miles ft/sec deg deg hr:min:sec
Translnnar injection S-IVB 2 , 44 , 16 . 2 347 . 3 10 441 . 0 896 . 3 6640 O . llS 174.1311 7 5 , 0 5 ,21
Command and service mod- React ion control 3 ' 17 , 04 . 6 7.1 0.7 827 . 2 6728 0 . 098 174. 89\1 75 , 07 , 4 7
ule/S -IVB separation
Spa.cecraft /S-IVB Service propulsion 4 , 4 0 , 01 . 8 2.9 19 . 7 180 . 8 7972 0 . 18N 175 . 97E 75 : 3 9 , 30
separation
First midcourse correc- Service propulsion 26 , 44 , 58 . 7 3.1 20.9 61. 5 8334 0 . 17N 173 . 5 7E 75 : 5 3 : 35
tion
After command and service mod- 4 : 40 : 0 1 . 0 36 079 -67 . 43 0 . 198 9 8 . 05E 160 : 32 : 27
ule/S-IVB separation
Ai"ter separat ion maneuver 11 : 2 8 : 0 0 . 0 36 139 -48 . 9 5 37 . 386 5 9 - 95E 146 : 3 9 ,27
After first midcourse correction 26 : 45 : 01 . 5 3 6 147 -10 . 25 18. 468 168. 10E 145 : 0 5 :28
Resultant orbit
Velocity
Ignition time , Firing t ime ,
Maneuver System change ,
hr:min:sec sec Apocynthion , Pericynthion ,
ft/sec
miles miles
Final separation Lunar module reaction 130 , 30 : 01 . 0 7.2 2.2 62.7 54.0
control
Transearth injection Service propul sion 135 : 2 3 , 42 . 3 151 . 4 3279 . 0 - 0 . 70 36 195 4 . 29N 180.15E 195 : 0 5 , 5 7
Second midcourse cor- Service module 150 : 2 9 : 57 . 4 11 . 2 4.8 -6 . 4 6 3 6 194 3 . 178 171. 99E 195 : 0 3 : 0 8
recti on react i on control
7-12
Time , hr :min : s ec 19 5 : 0 3 : 0 5 . 7
Geodeti c lat itude , deg s outh 3 . 19
Longi tude , deg e as t 171 . 96
Alt itude , mile s • • 65 . 8
Space-fixed velocity , ft /s ec 36 19 4 . 4
Space -fixed flight-path angle , deg -6 . 48
Space-fixe d heading angle , deg east of north 50 . 1 8
Velocity , ft /s e c ·' .
Acc elerat i on , g .
Drogue deployment
Time , hr :mi n : s e c 19 5 : 12 : 06 - 9
Geodetic l at itude , deg s outh
Recovery ship report 1 3 . 25
Onboard guidance 1 3 . 30
Target . . . • • • • 1 3 . 32
Longitude , deg wes t
Recovery ship report 169 . 15
Onboard guidance 16 9 . 15
Target 169 . 15
7-13
Apollo 10 Apollo ll
Actual 0 . 35 4 0 . 769
Apollo 10 Apollo ll
I
'ejj
F l ight-path angl
1000 36
v-"""'
/.v
�
�
900 f- 5 f- 34 �
cU
IJ Velocity
L
""C
�
"' .,
Q)
�
""C
� 0
QJ �
l/� 17
<II
<1:
800 � 4 32
.,
"' ':::;-
�
"'
�vv
..<::
� �
"'
�
�
c. u
700 1- I
3 0 ./
..<:: 30
� � 1/
a;
.!<' >
Altitude -
u.
.....
600 L 2 1- 28
,
v /
Velocity
� v v
�v /
,.
F l ight-path angle
26
� :,..""
� v j
�
,
-
24 ..
ࠂĆĭĆʳĭ ʵ؝ٜʴÛzÛ
§ࠦ Bý ;õ;B
Y Y
1 .2
1.1
/
k' 0 Pass solutions
D Optics solutions
1.0
/� R 2 predicted l:J. Premission target
v
u
0
"' <P o
<D o
"-
QJ 0
oo
"
"
.; - On
� oOO
:::1
0.9 - .. •v
� 0
'..0 0
� 0 ..... ....
0
0
o �P
.....
0 ,.. ,
.... ...
...
0.8 ·� LJ
D
',
�
'
l:J.
'
'
Langley-model
' predicted
1'.
0.7 '
Landing site _/ ",
'
',
0.6
0 5 10 15 20 25 30 35 40
Revolution
Figure 7-4 . - Se lenograph ic latitude estimates based on a one pass solution using R2 mode l .
8-1
The probe retract time for both events was between 6 and 8 seconds . Dur
ing the gas retract phase of lunar orbit docking , the crew detected a
relative yaw mis alignment that was estimated to have been as much as
15 degrees . See sections 4 . 15 and 5 . 7 for further discussion . The un
expected vehicle motions were not precipitated by the docking hardware
and did not prevent accomplishment of a success ful hard dock . Computer
simulations of the lunar orbit docking event indicate that the observed
vehicle mis alignments can be caused by lunar module plus X thrusting
after the command module is placed in an attitude-free control mode ( see
section 8 . 6 ) .
TABLE 8 . 1-I . - MAXIMUM SPACECRAFT LOADS DURING LAUNCH PHASE
Launch escape Bending moment , in-lb 520 000 l 000 000 136 000 310 000 110 000 173 000 230 000 110 0 0 0
36 DDD
system/ command
Axial force , lb -12 100 -11 ODD -22 200 -24 ODD -34 6DD - 5 DOD 8 ODD
module
Command module/ Bending moment , in-lb 680 DOD 1 320 000 166 000 470 000 340 000 590 000 300 000 140 ooo
service module
Axial force , lb -28 6oo -36 000 -88 200 -88 000 -81 6oo -89 600 ll 000 19 0 0 0
Servi c e module/ Bending moment , in-lb 696 000 l 620 000 2 000 000 2 790 000 l 220 000 540 000
adapter
Axial for c e , lb -193 300 -200 000 -271 000 -296 000 34 000 60 0 0 0
Adapter/instru- Bending moment , in-lb 2 263 000 4 620 000 2 600 000 5 060 000 l 400 000 440 000
ment unit
Axial forc e , lb -297 800 -300 000 -415 000 -441 000 51 000 90 000
The flight conditions at maximum qa were : The accelerations at the end of first-stage boost were :
c d
Condition Measured Predicted Acceleration Measured Predicted
a
Calculated from flight data.
b
Predicted Apollo 11 loads based on wind induced launch vehicle bending moment measured prior t o launch.
c
Predicted Apollo 11 loads based on measured winds aloft.
�redicted Apollo 11 loads for b lock II spacecraft design verifi cation conditions .
e
Predi cted Apollo 11 loads based on AS-506 static test thrust decay data.
8-4
8.2.1 Batteries
The bus voltages of the entry and pyrotechnic batteries were main
tained at normal levels , and battery charging was nominal . All three
entry batteries contained the cellophane separators , whereas only bat
tery B used this type of separator for Apollo 10 . The improved perform
ance of the cellophane separators is evident from voltage/current dat a ,
which show , at a 15-ampere load , that the cellophane type batteries main
tain an output 1 t o 2 volts higher than the Fermion-type batteries .
The only departure from expected performance was when battery A was
placed on main bus A for the translunar midcourse correction . During
this maneuver , normal current supplied by each battery is between 4 and
8 amperes , but current from battery A was initially 25 amperes and grad
ually declined to approximately 10 amperes j us t prior to removal from the
main bus . This occurrence can be explained by cons ideration of two con
ditions : ( 1 ) fuel cell 1 on main bus A had a lower ( 400° F ) than average
skin temperature , causing i"\; to deliver less current than usual ; and ( 2 )
b attery A had been fully charged just prior t o the maneuver . Both these
conditions , combined to result in the higher than usual current deli very
by battery A . Performance was normal thereafter.
The total b attery capacity was continuously maintained above 103 A-h
until separation of the command module from the s ervice module .
The fuel cells and radiators performed s atis factorily during the
prelaunch and flight phases . All three fuel cells were activated 68 hours
prior to launch , and after a 3-1/2-hour conditioning load , they were
placed on open-circuit inline heater operation unti l 3 hours prior t o
launch . After that time , the fuel cells provided full spacecraft powe r .
During the 195 hours of the mission , the fuel cells supplied approxi
mately 393 kW-h of energy at an average spacecraft current of 6 8 . 7 amperes
( 22 . 9 amperes per fuel cell ) and an average command module bus voltage of
29 . 4 volt s . The maximum deviation from equal load sharing between indi
vidual fuel cells was an acceptable 4 . 5 amperes .
During the flight , it was dis c overed that one heater in oxygen tank 2
was inoperative . Records show that it had failed between the times of the
countdown demonstration test and the actual countdown , and current meas
urements indicate that the element had an open circuit , This anomaly i s
dis cussed i n detail i n section 16 . 1 . 2 .
The operation of the VHF ranging system was nominal during des cent
and from lunar lift-off until orbital insertion . Following insertion ,
a number of tracking dropouts were experience d . These dropouts resulted
from negative circuit margins caused by use of the lunar module aft VHF
antenna instead of the forward antenna . After the antennas were switched ,
VHF ranging operat i on returned to normal . A maximum range of 246 miles
was measured , and a comparis on of the VHF ranging dat a with rendezvous
radar data and the predicted traj e ctory showed very close agreement .
8,5 INSTRUMENTATION
The data storage equipment did not operate during entry because the
circuit breaker was open . The c ircuit breaker which s upplies ac power to
the recorder also controls operation of the S-band FM transmitter . When
the television camera and as s oc i at e d monitor were to be powered without
transmitting to a ground station , the circuit breaker was opened to dis
able the S-band FM transmitter . This breaker was inadvertently left open
after the last television transmis s ion .
8-6
The other dis crepancy concerned the entry monitor system velocity
counter . The crew reported bias ing the counter to minus 100 ft /sec prior
to s eparation , thrust ing forward until the counter indi c ated 100 . 6 , then
thrusting aft unti l the counter indicated 100 . 5 . After the transposition
maneuver , the counter indicated 99 . 1 , rather than the expected 100 . 5 .
The cause of this apparent dis crepancy was also procedural . The trans
position maneuver was made at an average angular velocity of 1 . 7 5 deg/se c .
The entry monitor system i s mounte d approximately 12 feet from the center
of rotation . The resulting centripetal acceleration integrated over the
time necess ary to move 180 degrees yields a 1 . 2-ft /s e c velocity change
and accounts for the error observe d . The docking maneuver following
transpos ition was normal , with only small trans ients .
The crew was concerned with the durati on of the transearth injection
maneuver . When the firing appe ared to be approximately 3 seconds longer
than anticipated, the crew issued a manual engine-off command . Further
discuss ion of this problem is contained in s ection 8 . 8 . The data indi cate
that a computer engine-off dis crete appeared simultaneously with actual
engine shutdown . Therefore , the manual input , whi ch is not instrumented,
was either later than , or simultaneous with , the automati c command .
After entry into lunar orbit , and while still in the docked config
urat ion , the crew report ed a tendency of the space craft to pos ition its elf
along the loc al vert i cal with the lunar module pos itioned down . This ef
fect was apparently a gravity gradient torque , which can be as large as
0 . 86 ft -lb when the longitudinal axis of the vehicle is oriented 45 de
grees from the local vert i cal . A thruster duty cycle o f once every 15
to 18 seconds would b e con s i stent with a disturb ance torque of this mag
nitude .
8 . 6 .6 Landmark Tracking
8.6.7 Entry
The onboard calculat ions for inert ial veloc i ty and flight-path angle
at the entry interface were 36 195 ft / s ec and minus 6 . 4 88 degrees , respec
tively , and compare favorably with the 36 19 4 ft / s ec and minus 6 . 483 de
grees determined from tracking . Figure 13-1 shows a s ummary of landing
point dat a . The onboard computer indicat e d a landing at 169 degrees
9 minut e s west longitude and 1 3 degrees 18 minut e s north latitude , or
1 . 69 miles from the des ired target point . Since no telemetry nor radar
8-9
The gyro drift compensation updates were not as succes sful as ex
pect e d , probably because of the change in s ign of the comp ens ati on values .
With the change in the torquing current , a bias difference apparently
occurred as a result o f residual magnet i zat ion in the torquer winding .
The difference was small , however , and had no e ffect on the mis sion .
8.6.9 Computer
8 . 6 . 10 Optics
The sext ant and the s c anning teles cope performed normally throughout
the mis s ion . After the coelliptic s equence maneuver , the Command Module
Pilot reported that , after s elect ing the rende zvous tracking program ( P20 ) ,
8-10
Operation of the entry monitor system was normal , although one seg
ment on the electrolumines cent numerical display for the velocity counter
failed to operate during the mis sion ( see section 16 . 1 . 4 ) .
TABLE 8 . 6-I . - PLATFORM ALIGNMENT SUMMARY
79,10 3 33 Antares, 41 Dab ih + 0 . 100 +0.159 + 0 . 044 0 . 02 -1.2 -1 . 9 +0 . 5 Check star 3 3 Antares
103 , 00 3 10 Mirfak , 12 Rigel + 0 . 032 +0 . 009 + 0 . 001 0 . 02 -1. 2 -0 . 3 0.0 Check star 7 Menkar
134 , 3 4 3 1 Alpherat z , 11 Aldebaran +0 . 166 +0 . 212 -0 .019 0 . 01 -1.1 -1 . 4 -0.1 Check star 1 Alpherat z
149,19 3 14 Canopus , 16 Procyon +0 . 265 + 0 . 268 + 0 . 012 0 . 01 -1 . 5 -1 . 5 +0.1 Check star 11 Aldebaran
192,12 1 2 Diphda, 4 Achernar -1.166 -0 . 690 +0 . 456 0 . 00 -- -- -- Check stars 10 Mirfak, 1 Alpherat z ,
45 Fomalhaut , 3 Navi
Parameter
First midcourse Lunar orbit Lunar orbit Transearth
Separat ion
correction insertion circular! zation injection
.
Time
Ignition� hr :min :sec 4 : 40 : 01. 72 26 :44 : 5 8 . 6 4 75 : 49 : 5 0 . 37 80:11:36 .75 135 : 2 3 : 4 2 . 2 8
Cutoff , hr:min:sec 4 : 40 :0 4 . 65 26 :45 : 01 . 77 75 : 5 5 :47.90 80:11:53.63 135 : 2 6 : 1 3 . 6 9
!bration, sec 2 . 93 3.13 357-53 16 . 88 151.41
Velocity , ft/sec
(actual/desired )
X -9.76/-9.74 -14 . 19/-14 . 68 +327. 12/+327.09 +92. 53/+92. 51 + 9 32 - 77/+932 . 7 4
y +14 . 94/+14 . 86 +13.17/+13.14 +2361. 28/+2361 . 29 +118. 18/+118. 52 -2556 . 06/-25 55 .81
z +8. 56/+8. 74 +7 . 56/+7.66 +1681 . 85 /+1681 . 79 +51 . 61/+5 1 - 9 3 -1835 . 66/-1834 . 6 0
*Saturated.
NOTE: Velocities are in earth- or moon-centered inertial coordinates ; velocity residuals in body coordinates .
TABLE 8 . 6-III . - ENTRY MONITOR SYSTEM VELOCITY SUMMARY
Total velocity to be gained Velocity set into entry Planned Actual Corrected entry
Maneuver along X-axis , minus residual, monitor system counter, residual, res idual, monitor error,
ft/sec ft/sec ft/sec ft/sec ft /sec
Separation 19 . 8 15 . 2 -4.6 -4 . 0 +0 . 6
first midcourse correction 20 . 9 16 . 8 -4 . 1 -3.8 +0.3
Lunar orbit insertion 2917. 4 2910 . 8 -6 . 6 -6. 8 -0.2
Lunar orbit circularization 159 . 3 153 - l -6 . 2 -5 . 2 +1 . 0
Transearth inject ion 3283 . 2 3262 . 5 -20 . 7 -17 . 9 +2 . 8
Second midcourse correct ion 4.7 4.8 +0 .1 +0 . 2 +0.1
Distance
Time ,
Group Set/Marks Star Horizon from earth , Jemarks
hr :min
miles
2 1/3 1 Alpherat z Earth near 24 : 2 0 126 800 Optics calibration was zero. Not
entere d . Comput ed automat ic maneu-
2/3 2 Diphda Earth near ver onboard which did not consider
the lunar module ; therefore , diffi-
3/4 45 Fomalhaut Earth far 25 : 2 0 culty i n locating first star was
encountered as optics pointed at
lunar module . Ground-computed ma-
neuver was used and sightings pro-
ceeded satisfactorily .
8-15
Ac celerometers
Gyroscopes
c
X - Null bias drift , mERU -1. 2 1.7 9 0.4 -1 . 8 +2 . 4 -1. 2
Acceleration dri ft , spin reference
axis, mERU/g . -5.4 3. 8 9 -3. 3 -6.0
d
y - Null bias drift , mERU -1. 5 1.1 9 -2.4 -o. 6 +0.7 -1. 4
Acceleration drift , spin reference
axis , mERU/g 1. 7 2. 0 8 1. 3 3. 0
Acceleration drift , input axi s ,
mERU/g 7.1 5.6 14 9.0 5.0
e
z - Null bias d.rif't , mERU -0. 9 1.6 9 -2. 3 -0.2 -o . 6 -0. 1
Acceleration drift , spin reference
axis , mERU/g 8.4 6.6 8 20.4 5.0
Acceleration drift , input axis ,
mERU/g 0.8 6. 4 9 -4. 7 1.0
72
..-
- y
I-"'"
64
�-
f-""'
56 """'
/
.....
/
48
v
v
_,.V
./
....v
..
vv
vI--
/ r-
1
I
/
V"
8
......
0
X
z
-a
-16
0 liD uo 160 2111 3211 360 520 560 721
Time, sec
figure 8.6-l. - Velocity to���r��� ison 1111w1:en instrument unil n spaacnft guidance during·ascent.
8-19
8 . 7 .1 Servi ce Module
At the time the c ommand and s ervi ce modules s eparated from the S-IVB ,
the crew reported that the propellant is olat i on valve indi cat ors for
quad B indi c ated the "b arb er-pole " pos ition . This indi c ation corresponds
to at least one primary and one s econdary valve being in the clos ed pos i
tion . Twenty to thirty seconds after closure , the crew reopened the
valves according to checkli s t procedures , and no further problems were
experienced ( s ee section 16 . 1 . 6 ) .
All me as ured system pre ssures and temperatures were normal through
out the mis s i on , and except for the prob lem with the yaw engine , b oth
sys tems operated as expected during entry . Ab out 1 mi nut e after command
module /service module s eparation , system 2 was di s ab led and system l was
us ed for ent ry control , as planned. Forty-one pounds of propellant were
us ed during ent ry .
8. 8 SERVI CE PROPULSION
Servi ce propuls ion system performance was s atis fact ory during e ach
of the five maneuvers , with a total firing time of 5 31 . 9 se conds . The
actual ignition times and firing durat i ons are listed in table 8 . 6-rr .
8-20
Th e longest engine firing was for 357 , 5 s e c onds during the lunar orb i t
ins erti on maneuve r . T h e fourth an d fifth s ervi ce propuls i on firings were
preceded by a plus-X reaction control trans l at i on to e ffect prope llant
s ettling , an d all firi ngs were c onduct e d under aut omat i c control .
The engine t r an s i ent pe rforman ce during all s t art s and shut downs
was s at i s fact ory . The ch amber pre s s ure ove rshoot during the s t art of
the space cr aft s eparat i on maneuver from the S-IVB was approximat ely
120 ps i a , whi ch corresponds t o the upper s p e c i fi cati on limit for s t arts
using only one b ank o f propellant valves . On s ub s e quent firings , th e
chamb e r pres sure ove rshoot s were all le s s than 120 ps i a . During the
separat i on firing , minor os ci llat i ons i n the me as ure d chamb e r pre s s ure
were obs e rved beginning approximat e ly 1.5 s e c onds aft e r the i n i t i al firing
s i gnal . �oweve r , the magn itude of the os ci llat i ons was l e s s than 30 psi
( pe ak-to-peak ) , and b y approximat ely 2.2 s e c onds after i gni t i on , the cham
ber pres s ure dat a were indi c at i ng normal s t e ady-state operat i on .
The helium pre s s uri z at i on system funct i oned normally throughout the
mi ss ion . All system t empe ratures were mai nt ai ned within th eir red-line
limits with out heat e r operat i on .
t ank oxidi zer gaging error , is known to cause both the initial decrease
readings and a s tep increas e in the unbal an ce at cro s s over . The crew
were briefed on thes e conditions prior to flight and , therefore , expected
both the ini tial decreas e readings and a step increase at cros sover of
150 to 200 pounds . When the unbalance start e d to increase ( approach zero )
prior to cros sove r , the crew , i n anti cipat ion of the increas e , properly
interpreted the unbalance meter movement as an indication of a low mixture
ratio and moved the propellant uti li zat ion valve to the "increas e" pos i
t ion . As shown in figure 8 . 8- 1 , the unbalance then s tarted to decrease
in response to the valve change , and at cros sover the expected s tep in
creas e did oc cur . At the end of the firing , the crew reporte d that the
unbalance was a 50-pound increas e , which agrees well with the telemetered
dat a shown i n figure 8 . 8- 1 . This early re cognition of a lower mixture
rat io and the movement of the propellant uti liti zat ion valve to the "in
creas e " pos ition during lunar orbit ins ert ion resulted in a higher-than
pre di cted average thrust for the firing an d a durat i on of 4 . 5 s econds l e s s
than predicted.
The durat ion of the firing as determined by Mis s i on Control , was de
creas ed to reflect the higher thrust level experience d on the lunar orbit
ins ert ion firing . However , during the t ran searth i nj e ct i on firing , the
propellant utili zation valve was cycled from the normal to the decrease
position two times . This result e d in les s than the exp ected thrust and
consequently res ulted in an overburn of 3 . 4 s econds . ab ove the recalculated
transearth inj ect ion firing prediction .
NASA-S-69-3740
1
60
I I I
.
0 XI,d IZer -----
" Fuel --
-- '
50
r-......
' " I
r--
"-..
\),. Sump tanks
40
"'
c: �' primary gages -,
"' -....._ ! I
� '
0-
30 I'- '
12 "---., ....... I
m " ......
c:
'
0
::> .....
"---., '
20
' Storage tanks,
i / primary gages
! I ---.f_
... .......
�
10
I i
I
I "-...,
I
"---.,
I I
L
' I
I i
0
F i ri n g : 2 3
600
, · 1 - Lunar orbit i nsertion
--C--1 2 - lunar orbit circularization
400 -�
'
3 - Transearth injection
- -- - -- ,-
:ll
m
--
� 200
<..>
'
c
\
nJu
.,-
£! -.. ,--,
r .)..
I
0
<..> ' l �r""" '
·y f ' ,A .•
c
m "'
- � lV I �
v -J Y
Yll I
c � 200
.l'l :ll
::> :,:
:.; o
Oxidizer crossover
I Fuel crossover
.0!
"0
i ;
·;;; 400 I
0
Normal f-
�se1-
600
i
Decr
Propellant utilization valve position: No mal h
II
800
�rease
I
I1I
1000
Normal I I ncrease
The command module oxygen systems were us ed for parti culat e lunar
surface back-contaminat i on control from final command module docking
unt i l e arth landing .
At about 128 hours , the oxygen flow rate was adj us ted to an indi
cated reading of approximately 0 . 6 lb /hr to estab lish a pos itive differ
ent i al pre s sure between the two vehi cles , caus ing the cabin pre s s ure t o
increas e t o ab out 5 . 4 p s i a . The oxygen purge was terminated at 130 hours
9 minutes following the command module tunnel hat ch leak check .
The primary coolant system provi ded adequat e thermal control for
crew comfort and spacecraft equipment throughout the mi s s i on . The s e c
ondary c oolant system was activat e d only during redundant component checks
and the earth entry chilldown . The evaporators were not acti vat e d dur
ing lunar orbit coas t , since the radi ators provi de d adequat e tempe rature
cont rol .
Gas in the spacecraft potable water has been a problem on all manned
Apollo flights . On this mi s s ion , a two-membrane water/gas s eparator was
installed on both the wat er gun and the outlet at the food preparation
unit . The s eparat ors allow only gas to pas s through one membrane i nto
the cabin atmo sphere , while the s econd memb rane pas s es only gas-free
water to the outlet port for crew consumption . The crew indi c ated that
performan ce of the s eparators was s at is factory. Water in the food bags
and from the water pis tol was nearly free of gas . Two interface problems
were experienced while using the s eparators . There is no pos itive lock
between the wat er pi stol and the inlet port of the s eparator ; thus , oc
casionally the s eparat or did not remain i n place when use d to fill a food
bag from the water pistol . Also , the crew commented that s ome provis ion
for positively retaining the food bag to the s eparator outlet port would
be highly desirab le . For future spacecraft , a redes ign of the s eparator
will provide pos itive locking b etween the water pis tol and the i nlet port
of the separator . Als o , a change has been made in the s eparator outlet
prob e to provi de an improve d interface with the food bag .
8 . 10 CREW STATION
The di s plays and cont rols were adequat e except the miss ion clock in
the lower equipment bay ran s low , by les s than 10 s econds over a 24-hour
period , as reported by the c rew . The mis si on clo cks have a his tory of
slow operat ion , whi ch has been attributed to electromagnetic interference .
In addit ion , the glas s face was found to be cracked. This has also been
experienced in the past and is caus ed by stress introduced in the glas s
during the as sembly proc e s s .
The lunar module mis s ion clock is i dentical to the command module
clock . Bec ause of the lunar module clock problem dis cus s e d in section
16 . 2 . 1 , an improved-design t imer is b eing procured and will be incorpo
rated in future command modules .
8 . 11 CONSUMABLES
The service propuls ion propellant us age was within 5 percent of the
preflight estimat e for the mis s ion . The deviations which were experienced
have been attributed to the variations in firing times ( see s ect ion 8 . 8 ) .
In the following t able , the loadings were calculat e d from gaging system
readings and measured densities prior to lift-off .
Loaded
In tanks 15 6 33 2 4 967
In lines 79 124
Total 15 712 25 091 40 803 40 80 3
Loaded
Quad A 110 225
Quad B 110 2 25
Quad c 110 225
Quad D 110 2 25
Total 440 900 1340 1342
Loaded
System A 44 . 8 78 . 4
System B 44 . 4 78. 3
Tot al 89 . 2 156 . 7 245 . 9 245 . 0
Consume d
System A 15 . 0 26 . 8
System B 0.0 0.0
Total 15 . 0 26 . 8 40 . 8 39 . 3
8 . 11 . 3 Cryogeni cs
The oxygen and hydrogen usages were within 5 percent of those pre
dicted. This deviat i on was caus e d by the los s of an oxygen tank heater
element , plus a reduced reaction control system heater duty cycle . Us ages
listed in the following t ab le are based on the ele ct ri cal power produced
by the fuel cells .
Available at li ft-off
Tank 1 27 . 3 300 . 5
Tank 2 26 . 8 314 . 5
T ot al 54 . 1 56.4 615 . 0 634 . 7
Consumed
Tank 1 17 . 5 174 . 0
Tank 2 17 . 4 18o . o
Total 34.9 36 . 6 35 4 . 0 371 . 1
8.n.4 Water
Loade d
Pot ab le wat e r t ank 31 . 7
Waste wat e r t ank 28
Produce d i nflight
Fuel cells 315
Lithium hydroxi de , met ab ol i c NA
The crew inspected the descent stage thermal shielding after lunar
landing and observed no significant damage .
9-2
The knob on the as cent engine arm circuit breaker was broken , prob
ably by the aft edge of the oxygen purge system hitting the breaker dur
ing preparations for extravehi cular activity . In any event , this circuit
breaker was closed without diffi culty when required prior t o as cent ( sec
tion J6 . 2 . ll ) .
At staging , the des cent batteries had supplied 105 5 A-h of a nominal
total capacity of 1600 A-h . The di fference in load sharing at staging
was 2 A-h on b atteries 1 and 2 and 23 A-h on batteries 3 and 4 , and both
of these values are acceptable .
During the entire extravehi cular activity , the lunar module relay
provided good voice and extravehicular mobility unit dat a . Occas i onal
breakup of the Lunar Module Pilot ' s voice occurred in the extravehicular
communic ations system relay mode . The most prob able cause was that the
sensitivity of the voice-operated relay of the Commander ' s audio center
in the lunar module was inadvertently s et at less than maximum speci fie d .
This anomaly is dis cus s e d i n s e ction 16 . 2 . 8 .
After crew ingress into the lunar module , the voi ce link was lost
when the portable li fe support system antennas were stowe d ; however , the
dat a from the ext ravehi cular mobility unit remained good .
Lunar module voice and data c ommuni cat i ons were normal during the
lift-off from the lunar surface . The steerable antenna maintained lock
and tracked throughout the as cent . Uplink s ignal strength remained
stable at approximate ly minus 88 dBm.
9.5 INSTRUMENTATION
The gui dance and control system power-up sequence was nominal except
that the crew reported an initial di ffi culty in aligning the abort guid
ance system . The abort gui dance system is aligned in flight by trans fer
ring inert i al measurement unit gimbal angles from the primary guidance
system , and from these angles establishing a direction cos ine matrix .
Prior to the first alignment after activat i on , the primary system c ou
pling data unit s and the abort system gimbal angle registers must be
zeroed to insure that the angles accurately reflect the platform atti
tude . Failure to zero could c ause the sympt oms report e d . Another pos
s ible cause is an incorre ct s etting of the orbital rate drive electroni cs
( ORDEAL ) mode swit ch . If this switch is set in the orbital rate position ,
even though the orbital rate drive unit is powered down , the pitch atti
tude displayed on the fli ght director attitude indi cator will be offset
by an amount corresponding to the orbital rate drive res olver . No data
9- 4
After the des cent orbit insertion maneuver , an alignment check was
performed by making three teles cope sightings on the sun . A compari son
was made between the actual pitch angle required for the s un marks and
the angle calculated by the onboard computer . The res ults were well
within the allowable tolerance and again indicated a properly function
ing platform .
The inertial measurement unit was aligned five times while on the
lunar surface . All three alignment options were success fully utili zed ,
including an alignment using a gravity vector calculated by the onboard
accelerometers and a prestored azimuth , one utiliz ing the two vectors
obtained from two different star sightings , and one using the calculated
gravity vector and a s ingle star sighting to determine an azimuth .
The Lunar Module Pilot reported that the optical sightings as s oci
ated with these alignments were bas e d on a te chnique in whi ch the average
of five success ive sightings was calculated by hand and then inserted
into the computer . An analy s i s of thes e s uccess ive sightings indicated
that the random sighting error was very small and that the only signif
icant trend observed in the s uccess ive sightings was lunar rat e .
The plat form remained inerti al during the 17 . 5-hour period between
the third and fourth alignments . Because both of thes e alignments were
to the s ame orientat ion , it is possible to make an est imate of gyro drift
while on the lunar surface . Drift was calculat ed from three s ources :
the gyro t orquing angles , or mis alignment , indicated at the se cond align
ment ; the gimbal angle change hist ory in comparis on to that predicted
9-5
from lunar rate ; and the comparis on of the actual gravity tracking his
tory of the onboard accelerometers with that predicted from lunar rate .
The results ( table 9 . 6-II ) indicat e excellent agreement for the granu
larity of the data utili zed.
The abort guidance system was aligned to the primary system at least
nine times during the mis s i on ( t able 9 . 6-III ) . The alignment accuracy ,
as determined by the Euler angle differences between the primary and
abort systems for the eight alignments available on telemetry , was within
specification tolerances . In addition , the abort guidance system was in
dependently aligned three times on the lunar surface us ing gravity as
determined by the abort system accelerometers and an azimuth derived from
an external s ource . The resulting Euler angles are shown in table 9 . 6-IV .
A valid comparis on following the first alignment cannot be made because
the abort guidance system azimuth was not updated. Primary gui dance align
ments following the s econd alignment were incompatible with the abort . guid
ance system because the inertial measurement unit was not aligned to the
local verti cal . A comparison o f the Euler angles for the third alignment
indicated an azimuth error of 0 . 08 degree . This error resulted from an
incorrect azimuth value received from the ground and loaded in the abort
guidance system manually . The resulting 0 . 08-degree error in azimuth
caus ed an out-of-plane velocity difference between the primary and abort
systems at insertion ( see section 5 . 6 )
•
The ascent maneuver was nominal with the crew again reporting the
wallowing tendency inherent in the control technique use d . As shown in
table 9 . 6-V , the velocity at insertion was 2 ft/sec higher than planned .
This h as been attributed to a difference i n the predicted an d actual tail··
off characterist ics of the engine .
9-6
The abort guidance system , as stat e d , was used to monitor all pri
mary guidance system maneuvers . Performance was excellent except for
s ome isolated procedural problems . The azimuth mis alignment whi ch was
inserted into the abort guidance system prior to lift-off and which con
tributed to the out-of-plane error at insertion is dis cussed in the pre
vious section . During the as c ent firing , the abort guidance system
velocity-to-be-gained was used to compare with and to monitor the primary
system velocity to be gained. The crew reported that near the end o f the
insertion maneuver , the primary and abort system dis plays differed by 50
to 100 ft /sec . A s imilar comparison of the reported parameter differences
has been made postflight and is shown in figure 9 . 6-l . As indi cated , the
velocity di fference was as large as 39 ft /sec and was caused by the time
synchronization between the two sets of data not being precise . The cal
culations are made and displayed independently by the two computers , whi ch
have outputs that are not synchronized. Therefore , the time at which a
given velocity is valid could vary as much as 4 seconds between the two
systems . Both systems appear to have operated properly .
The inertial measurement unit was replaced 12 days before launch and
exhibited excellent performance throughout the mis sion . Table 9 . 6-VI
contains the preflight history of the inertial components for the inertial
measurement unit . The accelerometer bias history is shown in table 9 . 6-VII .
An accelerometer bias update was performed prior to undocking , with results
as shown .
The lunar module guidance computer performed as des i gned , except for
a number of unexpected alarms . The first of these oc curred during the
power-up s equence when the display keyboard circuit breaker was closed
and a 520 alarm (RADAR RliPT ) , which was not expected at this t ime , was
generated. This alarm has been reproduced on the ground and was caus ed
by a random setting of logic gates during the turn-.on se�uence . Although
this alarm has a low probability of occurrence , it is neither abnormal
·nor indicative of a malfunction .
The shift between the pre-installation calibrat ion dat a and the flight
measurements were as follows . ( The capability estimate limits are b ased
on current 3-s igma capabili ty estimates with exp ected measurement errors
included. )
Accelerometer Pre-installation
Free fall 48-day Capability
calibrat ion
( July 20 , 1969 ) shift e stimate
( June 6 , 196 9 )
X l -6 5 -66 185
y -17 -41 -2 4 185
z -66 -84 -18 185
When telemetered dat a were regained after the inflight calibration and
after powered ascent , excellent accelerometer stab ility was indicated as
follows . ( The capability estimate limits are b as ed upon current 3-sigma
capability estimat e s with expected measurement errors . included . )
Accelerometer
Capability
Before de s cent After ascent Shift
e stimate
Inflight calibration data on the gyros were reported and two lunar sur
face gyro calibrat ions were performed with the following result s . The
degree of stability of the ins truments was well within the expected
values .
9-9
The only hardware dis crepancy reported in the abort guidance system
was the failure of an electrolumines cent segment in one digit of the data
entry and display assembly . This is discussed in detail in section 16 . 2 . 7 .
'P
TABLE 9 . 6-I . - LUNAR MODULE PLATFORM ALIGNMENT SUMMARY
1-'
0
Alignment mode Telescope Star angle Gyro torquing angle , deg Gyro drift , mERU
Time , Type
detent e /star difference ,
hr :min aligr..ment a b
Option Technique used c.eg X y z X y z
a
3 - REFSMMAT ; 4 - Landing s ite .
b
l - REFSMMAT plus g ; 2 - Two bodies ; 3 - One body plus g .
c
l - Left front ; 2 - Front ; 4 - Right rear ; 6 - Left rear.
s
Star names :
25 Acrux
33 Antares
l2 Rigal
3 Navi
l3 Capella
10 Mirfak
9 -11
Lunar Surface
Inflight
Maneuver
Descent orbit Powered descent Coelliptic se- ConstWJt differ- Terminal phase
Condition Ascent
insertion initiation quence initiation ential height initiation
Time
hr :min :sec
a a
Ignition, 101:36 :14 102:33:05.01 124:22:00.79 125:19:35 126 : 1 7 : 4 9 . 6 127 : 0 3 : 51 . 8
Cutoff, hr :min : sec 101 : 3 6 : 44 102 :45 : 41 . 4 0 124 : 2 9 : 1 5 . 67 125:20:22
47 .o 1"1 .8
126:18:29.2 127:04 :14 . 5
Duration, sec 30.0 756.39 434.88 22.7
Gimbal drive actuator, in. (b) Not applicable Not applicable Not applicable Not applicable
Initial
Pitch +0.43
Roll -0.02
�laximum excursion
Fi"tch +0.03
Roll - 0 . 28
Steady-state
Pitch +0.59
Roll -0.28
+3.2 -o.l.
C-laximun attitude excurs ion, deg
Pitci1 (b) +1.2 (b) -l. 6
Roll -1.6 -2.0
-2 . 0 !0, 4
+0.8 -0.4
Ya�; -2.4 + :J . 8
a Reported by crew.
l::JTE : PG!fCS - Primary guidance , navigation, an d control system; DPS - Descent propulsion system; APS - Ascent propuls ion syste:r:,
RCS - Reaction control system.
!lendezvous maneuvers after terminal phase initiation are reported in section 5 and are based en crev reports.
Accelerometers
Gyros copes
2
Bias , em/sec
Condition
X y z
Updated value + 0 . 66 +0 . 04 + 0 . 03
122 : 31 : 02 -137 . 6 0 . 05
124 : 09 : 12 -177 . 6 -0 .15
126 : 10 : 14 -30 1 . 3 -2 . 01
9- 1'7
X -53 42 15 1 0
y -22 9 15 -11 -23 . 7
z -79 22 15 -66 -71 . 2
X 14 9 -430 -463 . 5
y 28 9 324 299 - 5
z 12 9 1483 1453 . 4
X 0 . 33 0 . 05 15 0 . 27 0 . 27
y 0 . 04 0 . 05 15 0 . 03 0 . 03
z 0 . 51 0 . 07 15 0 . 41 0 . 41
X -0 . 67 0 . 12 15 -0 . 6 5 -0 . 6 5
9-l8
NASA-S-69 -3741
1400
""'
1'. 1- t- A bort g � idance _
1-
1200 system data loss
'
"
1000 �
�
"'
(.)
1"'-
v """
Vl 800
2 Primary guidance system -
"
F--.'
l:'
� Abort gu idance system
"'
(.)
0
600
>
"
�' I
�'
'
400
�'
200
�'
1'\
0
�I'
1 24:28:00 :10 :20 :30 :40 :50 :29:00 :10 :20
Time , hr:min:sec
9 .7 REACTION CONTROL
The crew reporte d thrust chamber ass emb ly warni ng flags for three
engine pairs . The A2 and A4 flags occurre d s imult aneous ly during lun ar
module stat i on-keeping prior to des cent orbit insert i on . The B4 flag
appeared shortly thereafter and als o twice j us t b efore powered des cent
initiation . The crew beli eved these flags were accompanied by mas ter
alarms . The flags were re s et by cycling of the cauti on and warning elec
troni cs circuit breaker . See s e ct i on 16 . 2 . 14 for further di s cus s i on .
The chamber pre ssure switch in react i on control engine BlD failed
close d approximately 8 . 5 minutes aft er powered des cent initi at i on . The
swit ch remained close d for 2 minute s 5 3 s econds , then opened and func
tioned properly for the remainder of the mi s s ion . The failure mode i s
beli eved t o be the s ame as that o f pressure switch failures on Apollo 9
and 10 ; th at i s , parti culate contaminat ion or propellant residue holding
the switch clos ed . The only potenti al consequence of the fai lure would
have been the inability t o detect an engine fai led ·"off . "
Thermal characte ristics were s at i sfactory and all tempe ratures were
within pre di ct e d values . The maximum quad t emperature was 232° F on
quad l subsequent to t ouchdown . The fuel t ank temperatures ranged from
6 8° to 71° F .
The reaction control system was used in the as cent interconnect mode
during powere d ascent . The system us e d approximately 69 pounds of pro
pellant from the as cent propuls ion tanks .
\0
I
1\)
0
NASA-S-69-3806
120
I
System B
100
::::t- -
'
' ,,
::! System A 1-' r System A
�
80
-- --
.,;
\� :-::----
...
-o
"
"
l..
c.
f-1
"
X
'
.. ...
" System B
" 60 ....- ""
�
� ,_ l. -
c.
"
0 ' ,· ,
a:
40
20
0
98 100 102 104 120 122 124 126 128 130
Time , hr
360
"'
320
280
I
l /
t- I-I
24 0 I
!
p
t-\
� I
"U
Actual
I
v
"
"U 11
200
"
I
"
"-
X
I
" J
�
" ,J-1 r�
..'.:!
"
160
I 1,-7 " f- P lanned
"- 1\
e
a. I
I
1--i
120
- _./
I
P lanned�
80
....
·'
r
r· ,.J
L-Pv
' f- Ac\ual
40
0
w- -
1_-:;-:
Time , hr
\0
Pi
I
F i g ure 9 . 7-2 . - Total prope llant consumption .
9-22
The des cent propuls i on system operation was s atis factory for the
des cent orbit insertion and des cent maneuvers . The engine transients
and throttle response were normal .
9 . 8 .1 Inflight Performance
The des cent orbit insertion maneuver laste d 30 seconds ; the result
ing velocity change was 76 . 4 ft /se c . The engine was started at the mini
mum throttle setting of 1 3 . 0 percent of full thrust an d , after approxi
mately 15 s econds , was throttled to 40 percent thrust for the remainder
of the firing .
The duration of the powered des cent firing was 756 . 3 seconds , corre
sponding to a velocity change of approximately 6775 ft /sec . The engine
was at the minimum throttle setting ( 13 percent ) at the beginning of the
firing and, after approximately 26 seconds , was advanced to full throttle .
There was about a 45-second data dropout during this period but from crew
reports , the throttle-up condit ions were apparently normal . Figure 9 . 8-1
presents des cent propulsion system pressures and throttle settings as a
function of time . The data have been smoothed and do not reflect the
data dropout , and the throttle fluctuations just before touchdown .
During the powered des c ent maneuver , the oxi di zer interface pres
sure appeared to be oscillating as much as 67 psi peak-to-peak . The s e
oscillations were evident throughout the firing , although of a lower mag
nitude ( fi g . 9 . 8-2 ) , but were most prominent at about 50-percent throttle .
The fact that os cillations o f this magnitude were not observed in the
chamber pressure or the fuel interface pres sure measurements indicates
that they were not real . Engine performance was not affecte d . Oscilla
tions of this type have been observed at the White Sands Test Facility
on numerous . engines , on s imilar pressure measurement ins t allations . The
high magnitude pressure oscillations observed during the White Sands Test
Facility tests were amplifications of much lower pressure os cillati ons
in the system. The phenomenon has been demonstrated in ground tests
where small actual oscillations were amplifi e d by cavity res onance of a
pressure transducer as s embly , which contains a tee capped on one end with
the trans ducer on another leg of the tee . This is similar to the inter
face pressure transducer installation . The resonance conditions will
vary with the amount of helium trapped in the tee and the throttle set
ting .
9-23
The me asure d pres sure profile in the super criti c al helium tank was
normal . The preflight and inflight pres sure ri s e rates were 8 . 3 and
6 . 4 ps i /hr , respe ctively .
During propellant vent ing after landing , the fuel interface pressure
increased rapi dly to an off- s c ale reading . The fuel line had frozen dur
ing venting o f the supercrit i c al helium , trapping fuel between the pre
valve and the helium heat exchange r , and this fue l , when heat e d from en
gine soakb ack , c aus ed the pres sure ri se . See s ecti on 16 . 2 . 2 for further
di s cussion .
During the des cent orbit insert i on maneuver and the early portion
of powered des cent , the two oxi di zer propellant gages were i ndi cating
off- s c ale ( greater than the maximum 9 5 -percent indi cat i on ) , as expect e d .
·The fuel prob e s on the other hand were indi c at ing approximat ely 9 4 . 5 per
cent instead of re ading off-s cale . The propellant loade d was equi valent
to approximat ely 9/ . 3 and 96 . 4 percent for oxidizer and fuel , re spe ctively .
An initi al low fuel reading also had occurre d on Apollo 10 . As the firing
cont inued , the propellant gage s began to indi c ate consumption correctly .
The t ank 1 and t ank 2 fuel probe measurements agreed throughout the fir
ing . The t ank 1 and tank 2 oxi di zer prob e measurements agreed initi ally ,
but they began to diverge unt il the difference was approximately 3 per
cent midway through the firing . For the remainder of the firing , the
difference remaine d constant . The divergence was probably c aus ed by oxi
dizer flowing from t ank 2 to t ank 1 through the propellant cros s over line
as a result of an offset in vehicle center of gravity .
Landing
Propellant go/no-go Calculat e d
low level Engine de cision propellant
li ght on cutoff point deplet i on
116 5 20 0
Firing time remaining , s ec
NASA-S-69-3742
120
--
t- - t- - t- - t- - - - - "7 t-- r--r-- - -
II'
·v;
Throttle position __/
80 � 80
"E .,- r C hamber pressure
::!
"'
�
�
t'
Q.
"' �
I
60
'�
60
� ,_
r- - r- -
� -
c"
.>!
.0
�
�
"'
0 E
,'
�
'
.c
40 "' 40
�
Q.
u
"' ',
.c
--
>-
�-
"""'
20 20
1-
0
246
v
j..../
242
Regu lator outlet pressure
I -"'>
238
v
...=
--
�
�
234 '
�
.,-
::!
�
�
f'" ' -- P"
� 230
0.. " Fuel interface pressure
226
Time, hr:min
\.0
I
1\)
0\.
250
"'
c.
"'
�
"
"'
"' 200
c.
�
u
"
-£
<=
1l
N
Q;
-c
150
X
0
-c
<=
"'
"
;;
"'
"'
c.
� 100
Q;
-"
E
"'
-"'
u
50
Firing t i m e , sec
F i gure 9 . 8-2 . - Oxidi zer interface pressure and chamber pressure oscillations.
9-27
9 .9 ASCENT PROPULSION
The ascent propulsion system was fired for 435 seconds from lunar
lift-off to orbital insertion . All aspects of system performance were
nominal .
The regulator outlet pres sure was 184 psia during the firing and
returned to the nominal lock-up value of 188 . 5 psia after engine cutoff.
Table 9 . 9-I presents a comparison of the actual and predi cted perform
ance . Bas ed on engine flow rate dat a , the engine mixture ratio was esti
mated to be 1 . 595 . The estimated usable propellant remaining at engine
shutdown was 174 pounds oxidizer and 121 pounds fuel ; these quantities
are equivalent to 25 seconds additional firing time to oxidizer depleti on .
After ascent propulsion system cutoff and during lunar orbit , the
fuel and interface pressures increased from their respective flow pres
sures to lock-up , and then continued to increase approximately 3 . 6 psi
for fuel and 11 to 12 ps i for oxidizer . Loss of signal occurred approx
imately 39 minutes after engine shutdown as the vehi cle went behind the
moon . Pressure rises in the system were obs erved during both the Apollo 9
and 10 missions . This in'itial pressure rise after shutdown was c aused by
a number of contibuting factors , such as , regulator lockup , heating of
the ullage gas , and vaporization from the remaining. propellants .
�reflight predi ction b as ed on acceptance test dat a and assuming nominal system p erform�
ance .
b
Actual flight data with known bias es removed .
9-29
During the sleep period on the lunar surface , the crew reported that
they were too cold to sleep . Analysis of the conditions experienced in
dicated that once the crew were in a cold condition , there was not enough
heat available in the environmental control system to return them to a
comfortable condition . Ground tests have indicated that in addition to
the required procedural changes which are designed to maintain heat in
the suit circuit , blankets will be provided and the crew will sleep in
hammocks.
Shortly after lunar module ascent , the crew reported that the carbon
dioxide indicator was erratic , so they switched to the secondary car
tridge, Also , the secondary water separator had been selected since one
crewman reported water in his suit.
the optical section of the s ens or . Further dis cus s ion of both the errat
ic carbon dioxi de readings and water in the crewman ' s suit i s contained
in section 16 . 2 . 3 and 16 . 2 . 1 3 , respectively .
9 . 11 RADAR
Performance of the rende zvous and landing radars was satis factory ,
and antenna temperatures were always within normal limits . Range and
velocity were acquired by the landing radar at s lant ranges of approxi
mately 44 000 and 28 000 feet , respective ly . The tracker was lost brief
ly at alt itudes of 240 and 75 feet ; these los ses were expected and are
attributed to zero-Doppler e ffects as sociated with manual maneuvering .
9 .1 2 CREW STATION
The Commander and Lunar Module Pilot were provide d with communi ca
tions carrier adapter eartubes , having molded earpieces , for use in the
lunar module cabin. The purpose of these earphone adapters is to increase
the audio level to the ear . The Lunar Module Pilot use d adapters through
out the lunar module des cent and landing phase , but after landing , he
found the molded earpieces uncomfortab le and remove d them. The Commander
did not us e adapters s ince hi s preflight experience indicat e d audio volume
levels were adequate ; the us e of the adapters i s based on crew preference .
The Apollo 10 Lunar Module Pilot had used the adapters during his entire
lunar module operat ional period and reported no dis comfort . The Apollo 12
crew will als o be provided adapters for optional us e .
The crew commented that the inflight coverall garments would be more
utilitarian if they were patterned after the standard one-piece summe r
flying sui t . More pockets with a better method of closure , preferab ly
zippers , were re commended an d will be provide d for evaluat ion by future
crews .
9-31
The crew reported repeat ed fogging of the lunar module windows while
the sunshades were installed. They had transferred two of the command
module tis sue di spensers to the lunar module and made use of them in
cleaning the windows rather than using the window heaters for defogging .
Tissue di spensers are being adde d to the lunar module s towage list .
9 .13 CONSUMABLES
On the Apollo 11 mis s ion , the actual usage of only three consumable
quantities for the lunar module deviated by as much as 10 percent from
the preflight predicted amounts . These consumab les were the des cent
s tage oxygen , ascent stage oxygen , and react ion control system propellant .
The actual oxygen requirements were les s than predi cted be cause the leak
age rat e was lower than expected. The actual reaction control propellant
requirement was greater than predi cted becaus e of the increased hover t ime
during the descent phas e .
Consumed
Nominal 17 010
Redesignat ion 103
Margin for manual hover ll4
Tot al 6724 10 690 17 414 17 227
The actual ascent propuls ion system propellant us age was within .
5 percent of the preflight predictions . The loadings in the following
table were determined from me asured dens ities prior to lift-off and from
weights of off-loaded propellants . A porti on of the propellants was used
by the react ion control system during ascent stage operat i ons .
Consumed
By ascent propulsion sys- 1833 2934
tern prior to as cent stage
jett i s on
By react ion control system 23 46
Totai 1856 2980 4836 4966
The increased hover t ime for lunar landing resulted in a devi at ion
of over 10 percent in the re acti on control system propellant usage , as
compared with the prefl ight predi ct ions . Propellant consumpti on was cal
culated from telemetered helium t ank pre s s ure histories using the rela
t ionships between pres sure , volume , an d temperature . The mixture rat i o
was as sume d t o b e 1 . 9 4 for the calculat i ons .
Loaded
System A 108 209
System B 108 209
Total 216 418 634 633
Consume d
System A 46 90
System B 62 121
Total 108 211 319 253
9 .13. 4 Oxygen
The actual oxygen us age was lower than the preflight predictions
because oxygen le ak rat e from the cabin was less than the specifi cat ion
value . The actual rate was 0 . 0 5 lb /hr , as compared with the speci fi cat i on
rate of 0 . 2 lb /hr . In the following t ab le , the actual quantities loaded
and consumed are bas ed on telemetered dat a .
9-34
Loaded ( at lift-off )
Des cent st age 48 . 2 48 . 2
As cent stage
Tank l 2.5 2.4
Tank 2 2.5 2.4
Total 5 .0 4.8
Conswned
Des cent st age 17 . 2 21 . 7
As cent s t age
Tank l 1.0 1.5
Tank 2 0 .1 0.0
Tot al l.l 1. 5
9 . 13 . 5 Wat e r
The actual water us age was within 10 percent of the preflight pre
di ct i ons . In the following tab le , the actual �uantities loade d and con
s wne d are b as e d on telemet ered dat a.
9- 3 5
Loade d ( at li ft-off )
Des cent stage 217 . 5 217 . 5
As cent s t age
Tank 1 42 . 4 42 . 4
Tank 2 42 . 4 42 . 4
Tot al 84 . 8 84 . 8
Consumed
Des cent s t age 147 . 0 15 8 . 6
As cent st age
Tank 1 19 . 2 17 . 3
Tank 2 18 . 1 17 . 3
Total 37 . 3 34 . 6
9 .13 . 6 Helium
The consumed quantities of helium for the main propuls ion systems
were in close agreement with the predi cted amounts . Helium was stored
ambiently in the as cent s t age and supercriti cally in the des cent stage .
Helium loading was nominal , and the us age quanti ties in the following
t ab le were calculated from telemetered dat a . An addit i onal 1 pound was
stored ambiently in the des cent stage for valve actuat i on and is not re
flecte d in the values report e d .
Loaded 48 . 1 48 . 0 13 . 2 13 . 0
Consumed 39 . 5 38. 4 8.8 9.4
a8 . 6 b
Remaining 9.6 4.4 3 .6
a
At lunar landing.
b
At as cent stage j ett i s on .
10-1
The Command Module Pilot had a prob lem with the fit of the lower
ab domen and crot ch of his pre ssure garment as s emb ly , caus ed by the urine
collect ion and transfe r as sembly flange . Pres sure points resulted from
insuffi c i ent s i ze i n the pres sure garment as s embly . On future flight s ,
fit checks will be performed with the crewman wearing the urine collec
tion and transfe r assembly , fe cal containment system , and liquid cooling
garment , as applicab le . In addition , the fit check . will include a pos i
t ion s imulat ing that which the crewman experiences during the countdown .
All three pres sure garment as s emblies and the li quid cooling garments
for the Commander and Lunar Module Pilot were donned at approximat ely
97 hours in preparat i on for the lunar landing and surface operat i ons .
Donning was ac compli shed normally with help from another crewmen , as
required. The suit integrity check prior to undocking was completed
succes s fully with suit pressures decaying approximat ely 0 . 1 ps i .
of whether correct ive acti on is require d will be made after as ses sment of
Apollo 12 .
Extravehi cular act ivity preparat i ons proce ede d smoothly . However ,
more ime was required than planned for completing the unstowage of equip
t
ment and p erforming other minor t asks not normally emphas ized i n training
exerci ses .
The oxygen purge system checkout was p erforme d succes s fully . The
crew encountered two problems during pre-egre s s act ivities : ( 1 ) diffi
culty in mat ing the remote control unit connector and ( 2 ) bumping items
in the cab i n because of the bulk of the portab le life support system and
oxygen purge system ; as a result , one circuit breaker was broken and the
pos itions of two circuit bre akers were changed .
About 10 minutes was required to make each remote control unit con
nect or . Each t ime the crewman thought the conne ctor was aligned , the
lock lever rotation caus ed the connector to cock off to one s ide . The
problem i s dis cus s e d further in s ect ion 16 . 3 . 2 .
While wai ting for the .cab i n t o depressurize , the crew were comfort
ab le even though the inlet t emperature of the liquid-cooling garment
reached about 90° F prior to sublimator s tartup . No thermal changes were
not ed at egres s . The portable life support system and oxygen purge system
were worn quite comfortab ly , and the b ack-support e d mas s was not objec
ti'Onable in 1 /6-g .
a
Approximately 0 . 06 pound required for sui t i ntegrity check .
b
Approximat e ly 0 . 6 pound require d for start-up and t rappe d wat e r .
c
Minimum pre launch charge .
Crewman mob ility and b alance i n the ext ravehi cular mob i lity unit
were suffi ci ent to allow stab le movement while p erforming lunar surface
tasks . The Lunar Module Pilot demonstrat e d the capab i li ty to walk , t o
run , to change direct i on while running , and t o stop movement without dif
fi culty . He reporte d a t e ndency to t ip b ackwards in the soft s and and
noted that he had to be careful to compens at e for the different lo c at i on
of the center of mas s . The crewme n were ob served to kneel down and con
tact the lunar surface while retri eving obj e ct s . The crew stated that
getting down on one or both knee s to retri eve s amples and to allow closer
inspection of the lunar surface should b e a normal operat i ng mode . Addi
t i onal wai s t mobility would improve the ab i li ty to get clos er to the
lunar surface and , in addi ti on , would i ncre ase downward v i s ib i lity .
10-4
Throughout the ext ravehi cular activities , the crewmen made detailed
obs ervat i ons and ph ot ographs to document the activities and lunar surface
charact eri s ti cs . A televi s i on came ra provi ded real-time coverage of crew
ext ravehi cular activities .
The planned time line of maj or surface activiti es compared with the
actual time requi red is shown in t ab le 11-I . The table lists the events
sequenti ally , as presented in the Lunar Surface Operations Plan , and als o
includes s everal maj or unplanned activities . Crew rest periods , system
checks , spontaneous observations , and uns cheduled evaluations not neces
s arily related to the task being accomplished are not listed as separate
activities but are included in the times shown .
ll . 1 . 1 Summary
In the vicinity o f the lunar module , the mare s urface has numerous
small crat ers ranging in diameter from a few centimeters to s everal tens
of meters . Just s outhwest of the lunar module i s a double crat er 1 2 me
ters long , 6 meters wide , and 1 meter dee p , with a sub due d rai s e d rim .
About 50 meters east of the lunar module i s a s teep-walle d , but shallow ,
crater 3 3 meters in diameter and 4 meters deep , which was visited by the
Commander near the end of the extravehicular peri o d .
ll-4
All of the craters in the imme diate vi cinity of the lunar module
h ave rims , walls , and floors of rel at i ve ly fine grai ne d mat eri al , with
s c attered coars er fragments that occur in ab out the s ame abundance as on
the intercrater areas . Thes e crat ers are up to a meter deep and suggest
h aving b een excavat e d ent i re ly in the regolith b e c ause of the lack of
b locky e j e ct a.
The evi den ce suggests that proces s es of e ros i on are t aking place on
the lunar surface whi ch lead to the gradual rounding of the expos e d sur
faces of rocks . Several proces s es may b e i nvolve d . On s ome rounde d
rock surfaces , the indi vi dual c lasts ( fr agmented mat eri al ) and grains
11-5
that compos e the rocks and the glassy linings of pits on the surfaces hav'e
been le ft in raised reli e f by general wearing away or ab lat i on of the s ur
face . This different i al eros ion i s most prominent i n mi crobre c c i a ( rocks
consis ting of small sharp fragme nts emb edded in a fine-grained mat ri x ) .
The ab lat i on may b e caused primarily by small particles b omb arding the
surface .
Location of the landing site . - The landing s ite was tentat ively i dent
ified during the lunar surface stay on the b as is of ob servat i ons transmit
ted by the crew . The Commander report e d avoi ding a blo cky crater the
s i ze of a football field during landing , and ob s erve d a hill that he es
timated to be from 1/2 to 1 mile wes t of the lunar module . The lunar
module was tilt e d 4 . 5 degrees east ( backward ) on the lunar s urface .
During the firs t command and s ervi ce module pas s after lunar module
landing ( about 1 t o 1-1/2 hours aft er landing ) , the first of s everal dif
ferent landing s ite locations , comput ed from the onboard computer and from
tracking dat a , was transmitt ed t o the Command Module Pilot for visual
s e ar ch ( see section 5 . 5 ) . The first such es timate of the landing s ite
was northwest of the planned landing ellips e . The only site near this
compute d lo c at ion that could have mat ched the report e d des cription was
near North crater at the northwes t boundary of the landing ellips e . How
eve r , thi s region did not mat ch the des cription very closely . Later ,
computed e stimates indicated the landing s ite was cons iderably south of
the earlier determinat i on , and the areas near the West crat e r mos t closely
fit the des cription . Thes e , data were t ransmitt e d t o the Command Module
Pilot on the last pas s be fore lunar module lift-off , but the Command Mod
ule Pilot ' s activities at this t ime did not permit vis ual s earch . The
location j us t west of West crate r was confirmed by rende zvous radar track
ing of the command module by the lunar module near the end of the lunar
stay period and by the descent photography .
The crater that was avoi ded during landing was reporte d by the crew
to be surrounded by ej ect a containing blocks up to 5 meters i n diameter
and whi ch extended 100 to 200 meters from the crater rim , indi c ating a
relatively fresh , sharp-rimme d ray crat e r . The only crater in the 100-
to 200-meter size range that meets the des cription and is i n the vi cinity
indi c ated by the radar is West crat e r , near the s outhwest edge of the
planned landing ellipse . A des cription by the Commander of a double
crater ab out 6 to 12 meters in s ize and s outh of the lunar module shadow
plus the identi fi c at i on of West crater , the hill to the wes t , and the 21-
to 24-meter crat er reporte d behind the lunar module , formed a unique pat
tern from which the landing site was determined to within ab out 8 meters .
The 21 to 2 4 meter crater has been s ince identi fi e d by photomet ry as being
33 meters in diamet er . The returned s equence-camera des cent photography
confirmed the landing point lo c at ion . The pos i tion corresponds to coor
dinates 0 degree 41 minutes 15 s econds north lat itude and 23 degre es
26 minutes 0 second east longitude on figure 5-10 .
Geo logy . - The surface of the mare near the landing site i s unusually
rough and of great er geologic interest than expected before flight . Tele
vis i on pi ctures indicat ed a greater abundance of coarse fragmental debris
than at any of the four Surveyor landing s ites on the maria except that
of Surveyor I (ref . 8 ) . It is likely that the obs erve d fragments and the
11-7
s amples returned to e arth had b een derived from varying depths b eneath
the original mare surface and have had wi dely di fferent histori es of ex
posure on the lunar surface .
The maj or t opographic features in the landing are a are large crat e rs
a few hundred me ters acros s , of which four are b road sub dued features and
the fi fth i s Wes t crat e r , located 40 0 meters e as t of the landing point .
Near the lunar module , the surface i s pocked by numerous small crat e rs and
strewn with fragment al debris , part of which may have been generat e d dur
ing the impact formati on of Wes t crater .
Among the smaller crat e rs , both sharp , rai s ed-rim craters and rela
t ively sub dued crat ers are common . They range in s i ze from a few centi
meters to 20 meters . A s li ght ly subdue d , rai s ed-rim crater ( the reported
21- t o 24-meter crat e r ) 33 meters i n diamet e r and 4 meters deep occurs
ab out 5 0 meters e as t of the lunar module , and a doub le crater ( the re
ported doublet crat e r ) ab out 12 meters long and 6 meters wide lies
10 meters wes t of the lunar module at 260 degrees azimuth ( s ee fi g . 5 - 8 ) .
The walls and floors of most of the crat ers are smooth and uninter
rupte d by eithe r out crops or conspi cuous strat i fi cat i on . Rocks pres ent
in the 33-met e r crater are large r than any of thos e s een on the surface
in the vi cinity of the lunar module .
The bulk of the surface layer consists of' fine-grai ned part i cles
.which tended to adhere to the crewmen ' s b oot s and sui t s , as well as equip
ment , and was mol ded int o smooth forms in the footprint s .
The crewmen ' s b oot t re ads were sharply pres erve d and angles as large
as 70 degrees were maintained in the print walls ( s ee fig . ll-4 ) . The
surface disturbed by walking tended t o b re ak into slab s , cracking outward
ab out 12 to 15 cent imet ers from the e dge of footprint s .
The fine s t parti cles of the surface had s ome adhe s i on t o boot s ,
gloves , suits , hand t ools , an d rocks on the lunar s urface . On repeated
11-8
contact , the coat ing on the boots thickened to the point that their color
was completely ob s cure d . When the fine parti cles were b rushed off the
suits , a stain remai ne d .
During the televis ion panorama , the Commander pointed out seve ral
rocks west of the televi s i on camera , one of whi ch was t abular and stand
ing on edge , protruding 30 centimeters ab ove the surface . Strewn fields
of angular b locks , many more than 1/2 meter long , occur north and west
of the lunar module . In general , the rocks tended to be rounded on top
and flat or angular on the b ottom.
Televi s i on and phot ographi c coverage of the lunar surface act ivities
constitute most of the fundamental dat a for the lunar geology experiment
and complement informat i on reported by the crew . ( Refer to s ection 11 . 6
for a discussion of lunar surface phot ography . )
Phot ographi c document ation of the lunar surface was acqui red with
a 16-mm s equence camera, a close-up stereo camera , and two 70-mm sti ll
cameras ( one with an 80 -mm lens and the other with a 60-mm lens ) . The
camera with the 60-mm lens was intended primarily for gathering geologi c al
dat a , and a transparent plate cont aining a 5 by 5 mat rix of crosses was
mounted in front of the fi lm plane to define the coordinate system for
the opt i cal geomet ry .
The s ingle s ample survey was des i gned to record structure s that were
part i cularly s i gni fi c ant to the crew . The are a was photographed from a
distance of 1 . 6 meters . As with the s ample are a survey , the first pi cture
was t aken approximate ly down sun , and the next two were t aken cros s sun .
The third larges t rock i n the contingency s ample was collect e d with
in 2 meters of the lunar module . The rock has an ovoi d shape , t apere d at
one end , with broadly rounde d top and ne arly flat bottom ( s ee fig . 11 -6 ) .
It i s ab out 5 . 5 centimeters long , 2 to 3 centimeters wide , and 1-l/2 to
2 centimet ers thi ck . Part of the top and s i de s are cove red with fine dust
but the bottom and lower s i des indi c at e a very fine-grained clast i c rock
with s c attere d subrounded rock fragment s up to 5 millimeters in di ameter .
The rounde d ovoid shape of the top and s i de s of this specimen i s i rregular
in det ail . In the cent ral part , there i s a broad depre s s i on forme d by
many coales cing shallow i rre gular c avit ies and round pits . Adj acent to
thi s , t oward the t apered front end , round deep pits are abundant and s o
closely spaced th at some inters e ct others an d indi c at e more than one gene
rat i on of pittin g . The b ott om is marke d by two parallel flat surfaces ,
separat e d by an i rregular longitudinal s c arp ab out l/2 to l mi llimeter
high . A few small c avit i es are pres ent , but no round pits of the type
found on the top . An irregular fracture pattern occurs on the bottom of
the rock . The fractures are short , di s c ontinuous , and largely filled with
dust . On the top of the rock ne ar the t apere d end , a set of short frac
tures , 3 to 9 millimeters long , i's largely dust-filled and does not appear
11-10
to penetrate far into the rock . On a few s ides and corners , there are
short , curved fractures which may be exfoliation fe atures . This rock is
a brec ci a of small subangular lithic fragments i n a very fine grained
mat rix . It resemb les the materi al of the surface layer as photographed
by the stereo clos eup camera , except that this specimen i s indurat e d.
The two core-tube s amples were taken in the vi cinity of the solar
wind compos ition expe riment . The. first core locat i on was documented by
the televis ion camera and by two indivi dual Has s elblad photographs . The
second core-tube locat ion , as report e d by the crew , was in the vi cinity
of the solar wind compos ition experiment .
The sites of three of the cont ingency s ample rocks have - been loc at e d
and thos e of two tent at i vely i dent i fi e d b y comparing thei r shapes and
s i ze s from the lunar module window and surface photographs with photo
graphs t aken of the specimens at the Lunar Re ceiving Lab oratory . Evi dence
for the ident i fi c at i on and ori ent ation of rock A ( fi g . 11-9 ) was obtained
from the pre s en ce of a s addle-shaped notch on its expose d s i de . Rock C
( fi g . 11-10 ) was charact erized by the pitlike depress ion vi sible on the
photogr aphs . Rock B ( fi g . ll-9 ) is only ab out 2 centimeters acros s and
at this time has not b een correlat e d with the spe cimens in the Lunar Re
ceiving Lab orat ory ,
During bulk s amp ling , rock fragments were colle cte d primari ly on the
northeast rim of the large double crat e r s outhwest of the lunar module .
The large s coop , att ached to the extens i on handle , was us e d primar
ily during bulk s ampling to collect rocks and fine-grained mat erial . The
large s coop was us ed ab out 22 times in colle cting the bulk sample . As
expected from l/6-g s imulat i ons , s ome lunar materi al tende d to fall out
of the s coop at the end of s cooping motion .
The hamme r was us ed t o drive the core tubes att ached to the extens ion
handle . Hard enough b lows could be struck t o dent the top of the exten
sion h andle . The extens ion handle was att ached to the large s coop for
bulk s ampling and to the core t ubes for t aking core s ample s .
Two core tubes were driven and e ach colle cted a s at i s factory s ample .
Each tube had an internally t apere d bit that compre s s e d the s ample 2 . 2 : 1
within the inside of the tub e . One tub e colle cted 10 centimeters of
11-12
s ample and the other 13 centimet ers . The t ubes were diffi cult to dri ve
deeper th an ab out 20 centimete rs . Thi s di ffi culty may h ave b e en part i
ally c aus e d by the increas ing den s i ty o f the fine grained mat e ri al with
depth or other me ch an i c al charact e ri s ti c s of the lun ar regolith . The
di ffi culty of penet rat i on was also a functi on of the t apered b i t , whi ch
cause d great e r re s i st an ce with incre ased penet rat i on . One tube was di f
fi cult t o att ach t o the extens i on h andle . When thi s t ub e was detached
from the ext ens i on handle , the b utt end o f the tube un s crewed and was
lost on the lun ar surface . The tubes were opened aft e r the fli ght and
the s plit liners i ns i de b oth were foun d to be offset at the b i t end . The
Te flon core follower in one t ub e was origi nally insert e d ups i de down , and
the follower in the other tube was insert e d without the exp an s i on spring
which h olds it s nugly agai nst the ins ide of the spli t t ube .
The wei ghing s c ale was used only as a h ook to suspend th e bulk s am
ple b ag from the lunar module during the collect i on of bulk s ample s .
The lunar surface at the Apollo 11 landing s i te was s imi lar in ap
pe arance , b ehavi or , and me ch ani c al propert i e s to the surface encountered
at the Surveyor mari a landing s i tes . Alth ough the lun ar surface mat erial
di ffers c on s i derably in compos iti on and i n range of p art i cle shapes from
a t e rrestrial soil o f the s ame part i cle s i ze distribut i on , it does not
appear t o differ s i gni fi c antly in i t s engineering behavi or .
A variety of dat a was ob tained through detai led crew ob s ervat i ons ,
ph ot ography , telemetere d dynami c dat a , and examinat i on of the returned
lunar surface material and rock s ample s . Thi s i nformat i on permitte d a
prelimi nary as s es sment of the phys i c al and me chani cal prope rt i es of the
lunar surface mat e ri als . S imulat i ons b ased on current dat a are planned
to gain further i ns i ght i nt o the physi cal characteri st i cs and me ch an i c al
behavi or of lunar surface materi als .
11 . 2 . 1 Ob served Charact e ri s ti c s
Inspection of the area below the descent stage after landing re
vealed no evidence of an eros ion crater and little change in the apparent
topography . The surface imme diately underneath th e engine skirt had a
singe d appearance and was slightly etched ( fig . 11-14 ) , indicating a
s culpturing effect extending outward from the engine . Visible streaks
of eroded mat erial extended only to a maximum di stance of about 1 meter
beyond the engine skirt .
The landing gear foot pads had penetrated the surface 2 to 5 centi
meters and there was no di s cernible throwout from the foot pads . Fig
ures 11-15 through 11-18 show the foot pads of the plus Y and minus Z
and Y struts . The same photographs show the postlanding condi tion of
the lunar contact probes , which had dug into and were dragged through
the lunar surface , as well as s ome surface bulldoz ing by the minus Z
foot pad in the direction of the left lateral motion during landi ng .
The bearing pres sure on each foot pad i s 1 or 2 psi .
Natural clods of fine-grained material crumb led under the crewme n ' s
boot s . This behavior , while not fully understoo d , indicates cementation
and/or natural cohes ion between the grains . Returne d lunar surface s am
ples in nitrogen were als o found to cohere again to some extent after
being separat e d , although to a le s s er degree than ob s erved on the lunar
surface in the vacuum .
The surface mat erial was loos e , powde ry , an d fine-grained and ex
hibited adhes ive charact eristic s . As a result , the surface material
t ended to stick to any obj e ct with which it came i n contact , including
11-14
the crewmen ' s boots and suits , the televi s i on cab le , and the lunar equip
ment conveyor . During operat i on of the lunar equipment conveyor , the
powder adhering to it was carri ed into the spacecraft cabin . Als o , s uf
fici ent fine-graine d mat erial colle cted on the equipment conveyor to
caus e binding .
The thin lalfer of materi al adhering to the crewmen ' s b oot s oles
cause d s ome tenden cy to s lip on the ladde r during ingre s s . Similarly ,
the powdery coating of the rocks on the lunar surface was also s omewhat
s lippery ( s ee s e ction 4 . 0 ) . A fine dust confined betwe en two relatively
h ard surface s , such as a boot s ole and a ladder rung or a rock surface ,
would b e expected t o produce s ome tendency to s lip .
The lunar surface provide d adequate bearing stren gth for standing ,
walking , loping , or jumping , and suffi ci ent tract i on for s t arting , turn
ing , or stopping.
The material on the rim and walls of larger-si ze craters , with wall
s lopes ranging up to 35 degrees appeare d to be more compact and stable
than that on the smaller craters whi ch were trave rs ed.
The returned lunar material may be divided into the following four
group s :
g. All the rocks display glass-lined surface pits which may have
been caused by the impact of small particles .
ll-16
h . The fine mat e r i al and the bre c c i a contain large amount s of all
noble gases with e lemental and i s ot op i c abundances that almost cert ainly
were derive d from the s olar wind . The fact that interior s amples of the
b re c ci as c ont ain the s e gases implies that the b re c c i as were formed at
the lunar surface from mat eri al previ ous ly exposed t o the solar win d .
i . The 4 0 K/4 0 Ar me asurement s on i gne ous rock indi c ate that those
rocks crys t alli z e d 3 to 4 b i lli on ye ars ago . Cosmi c-ray-produce d nuclides
indi c at e the rocks h ave been within 1 mete r of the s urface for periods of
2 0 t o 160 million ye ars .
m . Elements that are enri che d i n i ron met eorites ( that is , n i ckel ,
cob alt , an d the plat inum group ) were e ither not observed or were low in
abundance .
Except for th e occas i onal o c currence of t rans i ent s i gnals , the b ack
ground sei smi c s ignal level on the long period vert i cal component seis
mometer is b elow system nois e ; th at i s , b elow 0 . 3 millimi cron over the
period range from l to 10 s econds ( s ee figs . ll-21 and ll-22 ) . Thi s is
between one hundred and ten thousand times les s than the ave rage b ack
ground levels ob s erve d on earth in the normal period range for mi cro
sei sms ( 6 to 8 se conds ) .
11-18
Sign als produced b y crew act i vities were promi nent on the short
period s eismomet er from initial turn-on unti l lunar module as cent . S uch
s ignals were part i cularly large when the crewmen were i n phys i cal contact
with the lun ar module . The s i gn al produce d when the Commander as cended
the ladde r to reent e r the lunar module is shown in fi gure ll-2 3 .
The pre domi nant frequency o f all o f thes e sign als i s 7 . 2 to 7 . 3 hert z .
The spe ct rum of the s i gnal produced by the Commander on the lunar module
ladde r , shown i n fi gure ll-23 , contains this prominent peak . This fre
quency is approximate ly equal to the fundamental res on ant mode of vibra
tion of the lunar module structure . The spectrum of the s i gnal generated
when one of the port ab le li fe s upport systems , weighing 75 p ounds , struck
the ground after b eing e j e ct e d from the lunar module is shown in figure
ll-24 for compari s on . The spectrum again shows the 7 . 2 hert z pe ak ; how
eve r , it is import ant to note that the two peaks at 11 . 3 and 12 . 3 hert z
would b e dominant i f the spectrum were corre cte d for instrume nt respons e .
The s i gn al at 7 . 2 hertz was presumab ly generat e d becaus e the portab le life
support system struck the lunar module porch and the ladder as it fell
to the surface .
S ome of the ob served high frequency s i gnals might pos sib ly have been
from ne arby meteoroi d impact s . An analysis i s being made of s everal high
frequen cy s i gnals whi ch may correspond to meteoroid impacts at ranges of
a few kilometers , or le ss , from the pas sive s ei smi c experiment package .
Substantive remarks on these event s cannot be made until spect ra of the
s i gnals are comput e d .
long period wave train ob s erve d on the re c ord is simply the s ummat ion of
t ransi ents corresponding t o these pulses and , hence , is of instrumental
orlgln . A di sper s i on of this type i s commonly ob served on earth in var
ious types of surface wave s and is well unde rs to o d . The dispers ion , or
gradual t rans format ion of an ini t i al impuls ive s ource to an extende d
oscillat ory t rai n o f waves , is produced by propagat i on through a wave
guide of s ome type . The events ob s e rve d appear only on the hori zontal
component s ei smometers . Such hori zont ally polari zed waves , when ob serve d
on e arth , would b e called Love wave s . On e arth , surface waves which have
a ve rt i c al component of mot ion ( Rayleigh wave s ) are usually the mos t prom
inent waves on the rec ord from a di stant event . Several pos sib i lities
are pre s ently under study to explain these waves .
11 . 4 . 4 Engineering Evaluat i on
From acquis ition of initial dat a to turn-off , the pass ive s ei smi c
experiment package operated a t ot al of 319 hours 18 minutes . The power
and dat a sub systems pe rforme d ext remely well , parti cularly in view of
the abnormally high operati ng t emperatures . The output of the s olar cell
array was within l to 2 wat t s of the expect e d value and was always higher
than the 27-watt minimum des ign spe c i fi cat i on .
About 9 9 . 8 percent of the dat a from the pas s ive s e i smi c experiment
package are pres e rved on t ape . Several oc currence s of dat a dropout were
determined to be c aus ed by other than the s e i smi c experiment system .
The pas sive s ei smi c experiment showed good respons e , detecting the
crewmen ' s foot steps , port ab le life s upport sys tem e j e ction from the lunar
module , and movement s by the crew in the lunar module prior to li ft -off .
The downlink s ignal s trength re cei ve d from the pas sive s ei smi c ex
periment package agree with the pre di ct i ons and for the 30-foot antennas
ranged from minus 135 to minus 139 dBm and for the 85-foot antennas
ranged from minus 12 5 to minus 127 dBm .
The i nitial impact of the los s of command capab i lity was the in
abi lity to re-level the long peri od s eismi c s ens ors . As a result , all
three axes became s o unbal anced that the dat a were me aningles s ; howeve r ,
me aningful dat a cont i nue d t o b e received from the short period s ens or.
The las er ranging retro-re fle ctor was deployed approximat ely 14 meters
s outh-s outhwe s t of the lun ar module in a relat i ve ly smooth are a ( s ee fi g .
11-26 ) . The bubble was not pre ci s ely in the center of the leveling devi ce
but was between the cent e r and the innermost divi s ion i n the s outhwe st
dire c t i on , indi c at i n g an off-level condit i on o f le s s than 30 mi nutes of
ar c . The shadow lines and sun compas s marki ngs were clearly vi s ible , and
the crew report e d that these devi c e s showed that the alignment was pre ci s e .
The s e ob servat i ons , made a few days b efore lunar sunset and a few
days aft e r lunar sunri s e , show that the thermal de s i gn of the refle ctor
permit s operat i on during s un i llumi nat e d peri ods and that the refle ctor
survived the lunar night s at i s fact ori ly . They als o i ndi cate no seri ous
degradat i on of opti cal per forman ce from fl aked insulat i on , debri s , dust ,
or rocket e xh aust product s which s c at t e re d duri ng lun ar module li ft -off .
11- 2 3
The experiment was deployed approximat ely 6 me ters from the lunar
module . The s t aff of the experiment penetrated 1 3 . 5 centimeters into the
surface .
The foi l was retreived after 77 mi nutes exposure t o the lunar en
vironment . The return unit was place d into a special Te flon b ag and re
turned to e arth in the lunar s ample return containe r . A port i on of the
foil was cut out , placed int o a metal gasket vacuum container , and heat
steri li zed at 125 ° C for 39 hours . The s ection of foi l has been released
for analysi s , and results will be reported in s ci ence reports .
11 . 7 PHOTOGRAPHY
SO- 3 6 8 , color 16 5 64 80 35
70 2
35 1
S 0-16 8 , color 16 8 * 63 32
70 2
3 40 0 , black 70 5 40 no 70
and white
When the lunar surface was vi ewed from the command module window ,
the color was report e d to vary with the viewing angle . A high sun angle
caus ed the surface to appear brown , and a low sun angle caused the sur
face to appear slat e gra;y . At this distance from the moon , distinct
- color variat ions were s ee n i n the maria and are very pronounced on the
proces s ed film . Acc ording to the crew , the 16-mm photographs are more
repres entat ive of the true s urface color than are the 70-mm photographs .
However , print s from both film types have shown t ints of green and other
shades which are not realist i c . Underexposure contributes to the green
t int , and the printing process can increas e this effe ct . Each generation
awa;y from the original copy will cause a further increase in thi s t int
ing . On the original film , the greenish t int in the dark , or underex
pos e d , areas is a funct i on of space craft windmv transmi s s i on character
i st i cs and low sun angles . For Apollo 12 , the master film copies will
be color corre cte d , whi ch should greatly minimize unreal i st i c t inting .
A 16-mm film s equence from the lunar module window shows crew activ
ities in both gray and light brown areas . As the crewmen move d , the gray
area , which i s apparently s ofter , deeper material , t urned almost b lack .
The crewmen ' s feet vis ibly sank i n this gray materi al as they ki cked mod
erat e quant it ies . The light brown are a did not appre c i ably change color
with crewmen ' s movement .
The color pictures i n which the fine grained parts of the lunar
surface appear gray are properly exposed , while those pictures i n which
the lunar surface i s light brown to light tan are generally overexp os e d .
ll-26
The rocks appear light gray to brownish gray in pi ctures that are pro
perly exposed for the rocks and vary from light t an to an off-white where
overexpos ed . The crew reporte d that fine grained lunar mat e ri al and rocks
appeared to be gray to dark gray . Thes e materi als appeared s lightly
brownish gray when ob s erve d near zero phas e angle . Small brownish , t an ,
and golden re fle ctions were ob s erve d from rock surfaces .
The targets and as sociat e d exposure values for e ach frame of the
lunar surface film magazines were carefully planned before flight . Nearly
all of the photographs were t aken at the re commende d expos ure s ettings .
A l franganus
0 10 20 30 40
Long i tude , deg E ast
NASA - S -6 9 - 3 7 45
; ......_
:
.
Approxi mate surface
'-.... contac t
......
}
.
.......
�--
.
·
·
.
. .
.
.
..
. ..
. . .
0 1 . .
..
I I I I I . . .. . .
Approximately
NASA-S-69-3746
എ Ꮶ።ᦹភቊᱎ
എᚓᚓᣗឃᏧᘅᣗᱎ Ꮸᘆᱎ ᦹមጕၢᆋቋᱎ ᖲၢᣗቋម Ꮹၢᓗ ՝ᱎ
ƀƁï@͘
ϯnwnŭwŮb̖"˨ ̈ ` ̏ ז
͓͚͕ז
'
'
$"'
!" ' % &# '
்ټʠ
éͤ ýז
$"'
' ;ķ O;æ ķ IA iķ
11-33
NASA- S -6 9 -.3 7 49
Front
end
White "crystals" and cleavage fragments in � Vertical irregularity hachures indicate down side
the matrix
�
Circular pits, many with raised rims; some with
thin wllite halos and locally on top
NASA-S-6 9-3 7 5 0
Front
end
0 1
NASA-S -69-3751
I
Television
\-
1
e;_........- -�.
'
) I
\. _
,
.
.
_ / / A rea from which core 2
·
I
• - sample was taken
.1 -
I
.
�- I -
,
_
•••
/
---;-
j -
B u l k sample
a rea
52 meters to
• 33-meter
dia meter crater
""'- � ./ '
...____ . - . - - ------
• 400 meters to/
/
+?. -� ...
Contingency
.
L-
est crater.-
/""!
_ _ _ sample a rea 1 :- --.....
�/ �-
·
· "'-.../
.- j ">,.., :.
.
k"
i..
-� -�: � �
r. j "-· � \
'- Documented
(" >,..
: sample area
I
J
j 7-�(. ... . .
:_
� ) \
�/"'\-
_j .
_J1.-- _ x:_ _--D/' (
.
\
10 '
_ _ ) F-"
of view
�
.... ........
•
''-JJ._ '•, .'
'
_
Gl
•
.
"'-· OD •
/ t"f",
.... . ..
Laser ranging �
retro reflector _/ ' ''?K
•
0 5
NASA-S-69-3 7 5 2
NA SA -5 -69-3754
--
F igure 1 1-13 . - Photograph of area shown in figures 11-11 and 1 1 - 1 2 , taken during extravehicu lar activity.
u-4o
NASA- S-69-3757
NASA-S-69-37 5 8
NASA- S -6 9 - 3 7 6 0
NASA-S-69 - 3 7 6 1
,.., I
Short per iod
.w Long per i od
I ..
I
c:
<II I ',
E I
,
"
<II
() ' '\
ru , , ---- --- '
0.. ' A
C/'1 I I \
'
',
I
\\
I
\'
-c '
'
-c I
I
I
c:
I ' \
\
:::s
0 I \
I I
\
...
C'l
' ' \
E I I \
() \
. ' ' \
:;::;. I I \ \
:::s
' ' \ '
0
' ' \
2
I
. ' \\
\
'
c: I \
:::s
' ' \ \
�C'l
I I \ \
' ' \ '
I \ \
-c
I
' ' \ \\
' '
\
�
<II . '
C/'1 ' \
I \
I
c:
&. \
C/'1
'
I I
: \
\
\\
� I I \
... . ' \
2 ' '
' \
Q)
I \ '
E \
'
'
'
\ '
'
0
E
.
C/'1
<II ' '
:
.
(/) ' \
' '
' ' '
. ' '
'
'
.
I
\ \
'
10 ����
2
°
���
1 0 - 2 10 - 1 10
Peri od , sec
NASA- 5 -69-3763
I1 (X a nd Y III (X, y Zl I
S i m ultaneously)
Seismic
Tide
No cha nge
20 m i n utes
NASA-S -69-.3 7 6 4
20 Type I, Y
oL-��L_ __
20
lO Type TI , X and Y
<lJ 0
•
Jill1
•we�
(f)
ra
..c Type m, z
...
:::!
0
� 10
Type m, x
Type ffi, Y
22 2.3 24 25 26 27 28 29 .3 0 .3 1 1 2 .3
J u ly -t- August
T ime , days
0 . 225
0 . 200
Ul
.....
0 . 175
s::
=s
...
>-
0 . 15 0
...
C1:l
.....
..0
...
C1:l 0 . 125
'
<II
-c
.....
=s
0 . 100
0..
E
C1:l
(/') 0 . 075
::2
�
0 . 05 0
0 . 025
0 2 4 6 8 10 12 14 16 18 20
Frequency , H z
NASA-S-69-.3 766
0 . 20
0 . 16
"'
......
1::
:::::;
>
...
!<:l
!:l 0 . 12
..0
...
!<:l
'
<1l
"'C
·;;_ 0 . 0 8
......
:::::;
�
(/")
:::2:
0:::
0 . 04
0 8 12 16 20
Frequency , H z
NASA-S-69-3767
700
i.
QJ
"0
100 ::s 400
::s :!::!
�
!ij
a.
....
QJ
a.
0 "0 300
QJ
E
� E
41 E
(J ::s
.f!
.... - 1 00
(f)
200
::s
(f)
-200 100
-300
21 22 23 24 25 26 27 28 29 30 31 1 2 3
July August
T i me , days
NASA-S-69-3768
NASA-S-69 -3 769
NASA-S-69-3 7 7 0
_-
12-l
12 . 0 BIOMEDICAL EVALUATION
The crew ' s health and performance were excellent throughout the
flight and the 18-day post flight quarantine period. There were no s ig
ni fi c ant physiological changes observed after this miss ion as has been
the cas e on - all previous miss ions , and no effects attributable to lunar
surface exposure have been obs erve d .
The biomedical data were of very good quality . Only two minor prob
lems occurred , both late in the flight . Data from the Command Module
Pilot ' s impedance pneumogram became unreadable and the Lunar Module Pilot ' B
electrocardiogram s ignal degraded because of drying of the electrode paste
under the sens ors . The Lunar Module Pilot replaced the electrocardiogram
leads in his bioinstrumentation harnes s with the spare set from the medi
cal kit , and proper readings were restore d . No attempt was made t o cor
rect the Command Module Pilot ' s respiration signal because of entry prep
arations .
The average heart rates during the entire mis s ion were 71 , 60 , and
67 beats /min for the Commander , Command Module Pilot , and Lunar Module
Pilot , respectively . During the powered des cent and ascent phases , the
only data planned to be available were the Commander ' s heart rates , which
ranged from 100 to 150 beats /min during descent and from 68 to 120 during
as cent , as shown in figures 12-l and 12-2 , respectively .
Commander ' s heart rat e during the last phas e s of this activity is indi ca
tive of an increas ed work load and body heat s torage . The me taboli c pro
duct ion of each crewman during the extravehi cular activity is reported
in s ection 12 . 3 .
12 . 2 MEDICAL OBSERVATIONS
During the first 2 days of the flight , the Command Module Pilot re
port ed that half a meal was more than enough to sat is fy his hunge r , but
his appetite subsequently returne d .
12 . 2 . 2 Medicat ions
The Commander and the Lunar Module Pilot each took one Lomotil tablet
prior to the sleep period to retard bowel movements before the lunar mod
ule activity . They each carried extra Lomotil tablets into the lunar mod
ule but did not take them. At 4 hours b e fore entry and again after splash
down , the three crewmen each took anti-naus eant tablets containing 0 . 3 mg
Hyos cine and 5 . 0 mg Dexedrine . Aspirin t ab lets were also taken by the
crewmen , but the number of t ab lets per individual was not recorded. The
Lunar Module Pilot recalled that he had t ak en two aspirin t ab lets almost
every night to aid his s leep .
12 . 2 . 3 Sleep
It is int eresting to not e that the crewme n ' s sub j e ctive e stimat es
of amount of sleep were les s than those b as e d upon t elemetered biome di
cal dat a , as shown in table 12-I . By either count , the crewmen s lept
well in the command module . The s imult aneous s leep periods during the
trans lunar coast were carefully monitore d , and the crew arrived on the
lunar surface well rested. Therefore , it was not nece s s ary to wait until
after the first planned 4-hour s leep period before conduct ing the extra
vehicular act ivity . The crewmen slept very little in the lunar module
12-3
12 . 2 . 4 Radi ation
The personal radi ation dos imeters were read at approximately 12-hour
intervals , as planned. The total integrated , but uncorrect e d , dos es were
0 . 25 , 0 . 26 , and 0 . 2 8 rad for the Comm�der , Command Module Pilot , and
Lunar Module Pilot , respectively . The Van Allen Belt dosimeter indicated
tot al integrated doses of 0 . 11 rad for the skin and of 0 . 08 rad for the
depth reading during the ent ire mis s ion . Thus , the total dose for each
crewman is estimated to have been less than 0 . 2 rad , whi ch is well below
the medically s ignificant level . Results of the radio-chemi cal as s ays of
feces and urine and an analysis of the onboard nuclear emulsi on dosimeters
will be presented in a separate medical report .
12 . 2 . 5 Inflight Exercise
12 . 2 . 6 Drug Packaging
the kit . Venting of each of the plastic or foil containers will be accom
plished for future flights and should prevent this problem from recurring .
The Afrin nasal spray bubbled out when the c ap was removed and was there
fore unus able . The use of cotton in the spray bottle is expected to re
s olve this problem on fut ure flight s .
12 . 2 . 7 Water
A new gas /water s eparator was used with s atis factory results . The
palatability of the drinking water was greatly improved over that of pre
vious flights because of the abs ence of gas bubbles , which can cause
gastro-intestinal dis c omfort .
12 . 2 . 8 Food
The food supply for the command module included rehydratable foods
and beverages , wet-packed foods , foods contained in spoon-bowl packages ,
dried fruit , and bread . The new food items for this mis s i on were c andy
sti cks and j ellied fruit c andy ; spreads of ham , chicken , and tuna s alad
packaged in lightweight aluminum , easy-open cans ; and cheddar cheese
spread and frankfurters packaged in flexible foil as wet-packed foods .
A new pantry-type food system allowed real-time s e lection of food items
b as ed upon indivi dual preference and appetite .
Four meal periods on the lunar surface were s cheduled , and extra
optional items were included with the normal meal packages .
Prior to flight , each crewman evaluated the available food items and
s elected his flight menus . The menus provided approximately 2300 kilo
calories per man per day and included 1 gram of calci um , 0 . 5 gram of
phosphorus , and 80 grams of protein . The crewmen were well s at i s fied
with the quality and variety of the flight foods . They reported that
their food intake met their appetite and energy requirements .
12-5
12 . 3 EXTRAVEHICULAR ACTIVITY
The integrated rates of Btu production and the accumulated Btu pro
duction during the intervals of planned activities are listed in table
12-II . The actual average metaboli c production per hour was estimated
to be 900 Btu for the Commander and 1200 Btu for the Lunar Module Pilot .
These values are les s than the preflight estimates of 1350 and 1275 Btu
for the respective crewmen .
12 . 4 PHYSICAL EXAMINATIONS
The post flight medi cal evaluation included the following : mi crobi
ology studies , blood studies , physical examinations , orthostatic toler
ance tests , exercise response test s , and chest X-rays .
The recovery day examination revealed that all three crewmen were
in good health and appeared well rested. They showed no fever and had
lost no more than the expected amount of body weight . Each crewman had
taken anti-motion si ckne s s medication 4 hours prior to entry and again
after landing , and no seas i cknes s or adverse symptoms were experi enced.
The Pub lic Laws and Federal Regulat ions concerning contaminat ion
control for lunar s ample return mi s s i ons are des cribed in reference 9 .
An interagency agreement between the N ational Aeronaut i cs and Space Ad
mi ni s trat i on ; the Department of Agri culture ; the Department of Health ,
Educat i on and Welfare ; the Department of the Interi o r ; and the Nat i onal
Academy of S ci ences ( ref. 10 ) confirmed the exi sting arrangements for the
protection of the earth ' s biosphere and defined the Interagency Committee
on Back Contami nat i on . The quaranti ne s chemes for manned lunar mis s ions
were estab li shed by the Interagency Committee on Back Contaminat i on
( re f . 11 ) .
12 . 5 . 1 Lunar Exposure
Although each crewman att empt e d to clean himself and the equipment
be fore ingres s , a fairly large amount of dust and grains of lunar s ur face
mat e ri al was brought into the cabin . When the crewmen remove d their hel
mets , they not i ce d a di st inct , pungent odor emanat ing from the lunar mate
rial . The texture of the dust was like powdered graphit e , and both crew
men were very dirty after they removed the i r helmet s , overshoes , and
glove s . The crewmen cleaned their hands and fac es with t i s sues and with
t owels that had been s oaked in hot water . The Commander removed his
liquid-cooling garment in order t o clean his body . One grain of material
got into the Commande r ' s eye , but was e as i ly removed and caused no prob
lem. The dust -like mat e rial c ould not be removed c ompletely from bene ath
their fingernai ls .
The concern that part i cles remaining in the lunar module would float
in the cab in atmosphere at zero-g aft e r as cent caus e d the crew to remain
helmet e d to prevent eye and bre athing contaminat ion . However , float ing
part i cles were not a problem . The cabin and equipment were further
cleaned with the vacuum brush . The equipment from the surface and the
pres s ure garment as s emblies were placed in bags for transfer to the com
mand module . Before transfer to the command module , the spacecraft sys
tems were configured t o cause a posit ive gas flow from the c ommand mod
ule through the hatch dump / rel i e f valve in the lunar module .
The c ommand module was cleaned during the return to earth at 2 4 -hour
intervals using the vacuum brush and towels . In addit ion , the c irculat i on
of the cabin atmosphere through the lithium hydroxide filters continued
to remove traces of part i culate material .
swimmer ret ired wi th the s econd life raft to the origi nal upwind pos i
t i on . The hatch was opene d , the crew ' s biologi cal i solat i on garments
were ins e rt ed into th e command module , and the hat ch was clos e d .
The crewmen were brought up i nto the heli copter without i nc i dent
and remained in the aft compartme nt . As exp e ct e d , a moderate amount of
wat e r was pres ent on the floor aft er retri eval , and the wat e r was wiped
up with towels . The helicopter crewmen were als o prot e ct e d from pos s ible
contaminat i on .
The heli c opter was move d to the Mobile Quarantine Faci lity on the
lower deck of the re covery ves s el . The crewmen walked acro s s the de ck ,
entered the Mobile Quarantine Facility , and remove d their biological
i s ol at i on garments . The de s ce nt s teps and the de ck area between the
heli copter and the Mobile Quarantine Facil i ty were sprayed with glut aral
dehyde solution , which was moppe d up afte r a 30-minut e contact t ime . .
After the crewmen were p i cked up , the prot e cted swimme r s crubbed the
upper deck around the postlanding vents , the hat ch are a , and the flotat i on
collar near the hat ch with Betadine . The remaining Betadine was emptied
into the bottom of the recovery raft . The swimmer removed h i s biologi cal
i solat i on garment and pl ace d i t i n the Betadine i n the life raft . The
di sinfectant sprayers were di smantled and sunk . Afte r a 30-minute contact
time , the l i fe raft and remai ning equipment were sunk .
The crew became uncomfort ab ly warm while they were enclos e d in the
bi ological i s olat ion garme nts in the environment ( 90° F ) of the heli
copter cabin . On two of the garme nts the vi s or fogged up b e c aus e o f im
proper fit of the nos e and mouth cup . To alleviate this di s comfort on
future mi s s ions , con s i derat i on i s b e i ng given to : ( l ) replacing the
pres ent b iologi cal i solat i on garment with a lightweight coverall , s imilar
to whi teroom clothi ng , wi th respirator mask , cap , gloves , and b oot i es ;
and ( 2 ) wearing a liquid cooling garment under the biologi cal i s olation
garment .
12-9
The command module was taken ab oard the USS Hornet ab out 3 hours
aft e r landing an d att ached to the Mob i le Quarantine Fac i li ty through a
flexible tunnel . The removal of lunar surface s ample s , film, dat a t ap e ,
and medical s amples went well , with one exception . Two of the medical
s ample cont ainers leaked within the inner b i ologi cal i s olation container .
Corre ct ive me asures were promptly exe cut e d , and the quarantine pro c e dure
was not vi olat e d .
Tran s fer o f t h e Mob ile Quarantine Facility from the re covery ship t o
a C-14 1 air craft and from the aircraft t o the Lunar Receiving Lab oratory
at the Manne d Spacecraft Center was ac c omplished without any ques tion of
a quarantine violat i on . The transfer o f the lunar surface s amples and
the command module into the Lunar Re c e iving Laboratory was al s o accom
pli shed as planne d .
12 . 5 . 3 Quarantine
A tot al of 20 pers ons on the medi cal support teams were expos ed ,
directly or indirectly , to lunar material for periods ranging from 5 t o
1 8 days . Daily medical ob s ervat i ons and periodic lab oratory examinat i ons
showed no s igns or symptoms of infecti ous di s ease related to lunar ex
posure .
No microbial growth was obs erve d from the prime lunar samples aft er
15 6 hours of i ncubat i on on all types of differe nti al me di a . No micro
organisms whi ch could be attribut e d to an ext rat e rrestrial s ource were
recovered from the crewmen or the space craft .
No s igni f i c ant trends were noted in any bioch emi cal , immunologi c al ,
or hematologi c al parameters in either the flight crew or the medi cal sup
port pers onnel .
12-10
The personnel in quarant ine and in the crew re c ept i on area of the
Lunar Receiving Lab oratory were approved for release from quarant ine on
August 10 , 1969 .
The s amples of lunar material and other it ems st ored in the bi olog
i c al i s olat i on containers in the Lunar Receiving Laboratory are s cheduled
for release to principal s cient i fi c invest i gators in S ept ember 1969 .
' ,
Time of
Telemetry Crew report
crew report ,
hr :min
Command Module Lunar Module Comman d Module Lunar Module
Commander Comman der
Pilot Pilot Pilot Pilot
23 : 0 0 10 : 2 5 10 : 10 8 : 30 7 :00 7 : 00 5 : 30
71 : 2 4 9 : 35 (a) 9 : 20 7 : 30 7 : 30 6 : 30
95 : 25 6 : 30 6 : 30 5 : 30 6 : 30 6 : 30 5 : 30
Totals 36 : 10 -- 32 : 3 5 29 :00 30 : 0 0 2 5 : 30
a
No dat a available .
12-12
Commander
Environment al familiarizat ion ; deploy televis ion 109 : 4 4 14 1200 280 963
cable
Terminat e extraveh icular activity , ingre s s , and 111 : 2 3 14 1650 385 2945
transfer sample return containers
HOTE : Value s are from the integration of' three independent determinations of metabolic rate based on
heart rat e , decay of oxygen supply pressure , and thermodynami cs of the liquid cooling garment.
,' ' '
I
NASA-S-69-377 1
180
Land ing
160
1 0 0 0 fee d
altitude i
! "Go" for stay
v ��'\\_
.
<::
2 0 0 0 feed
E
'- 140
<II
altitude i
�
"' Powered descent
ClJ
1\
..c in itiation
a;
1U
�
�
lil
ClJ
12 0 V I '-.
:I:
100
1\
80
1 0 2:33 1 0 2:35 1 02:37 1 0 2:39 102:41 1 0 2 : 43 1 0 2:45 1 02:47 1 02:49 102:51
Time, hr:min
140
�
120
" G o " for Ascent eng ine Ascent engine!
t:
E
l i ft-off ignition cutoff
i a
! i
......
C/)
....
l1l
�
Q)
i I
100
Q)
1U
�
....
�
l1l
Q)
:::c
80
60
124:14 1 24: 1 6 1 2 4: 1 8 1 2 4: 2 0 124:22 1 2 4: 2 4 124:26 124:28 1 2 4: 3 0
T ime , hr:m i n
NASA-S-69-3 773
••··�=B�u�lk�s�a:m
- F lag and President ' s message
lp le col lection
Lunar module inspection
Experiment package deployment -
Documented sample col lection ••••
Transfer sample return containers -
Terminate extravehicu lar activity I
� 160 r------.------,------,------,------r------r------r-���
E
� 140 r-----
---r--+--�--�� --�---++-+4�L¥�
-;;;
QJ
� 1 2 0 r-----frt-�T-+----�--���-�f--4
�
�
1 0 0 r----t-t�r-�-r\r�f-����Hr--��rn��-t---r----1
ffi
� 8 0 �-----L-------L------�------�----�-------L--� _L __ ____ __J
109:00 1 0 9 : 2 0 109:40 1 1 0: 0 0 1 1 0:20 1 1 0:40 1 1 1:00 1 1 1: 2 0 1 1 1:40
Time, hr:min
(a) Commander (CD R l .
] 1 0 0 �-----
--+----��--1---.-�r-- ---�--,H�nT�
2
� 80
�
ffi
QJ 60
::t:
1 09 : 0 0 109:20 109:40 1 1 0:00 1 10:20 1 1 0:40 1 1 1:00 1 1 1:20 1 1 1:40
T ime, hr:min
(b) Lunar Module Pi lot (LMPl .
13 . 1 FLIGHT CONTROL
The flight control respons e to thos e prob lems identified during the
mis s i on was bas e d on real-time dat a . Sect i ons 8 , 9 , and 1 6 should be
consulted for the post flight analyses of th es e problems . Three of the
more pertinent real-time de cisions are dis cus s ed in the following para
graphs .
At acqui sition of signal after hmar orbit ins ert i on , dat a showe d
that the indi cated t ank-B nitrogen pres sure was about 300 ps i lower than
expect e d and that the pre s s ure had started to decrease at 80 seconds into
· the maneuve r ( s ee section 16 . 1 . 1 ) . To cons erve nitrogen and to maximi ze
system reliability for t rans earth inj e ct i on , it was re commended that the
circularization maneuver be pe rforme d using bank A only . No further leak
was apparent , and b oth banks were us ed normallY for t ransearth inj e ction .
During the crew rest period on the lunar s urface , two checkli st
changes were re commende d , b as ed on the events of the previ ous 20 hours :
( l ) the rende zvous radar would remain off duri ng the ascent · fi ring , and
( 2 ) the mode-s ele ct switch would not b e place d in the primary gui dance
pos ition , thus preventing th e computer from generating alt itude and al
titude rate for th e telemetry display . The reas on for these changes was
to prevent computer ove rload during ascent , as had occurred during des cent .
13-2
13 . 2 NETWORK PERFORMANCE
The Mi s s ion Control Center and the Manned Space Fli ght Network were
placed on mis s i on s t atus on July 7 , 19 6 9 , and s at i s factorily s upport e d
the lunar landing mi s s i on .
The s upport provi ded by the real-time c omputer complex was generally
excellent , and only one maj or problem was experi enc e d . During trans lunar
coast , a problem in updat ing di gital-to-televi s ion display s by the primary
computer res ulted in the loss of all real-time televi s i on displays for ap
proximately an hour . The problem was is olated t o the interface between
the compl:ter and the display equipment .
Operations by the communi c at i ons proc e s s ors were excellent , and the
few prob lems caus e d only minor los s e s o f mis s i on dat a .
Both C - and S-b and tracking support was very good. Los s o f tracking
coverage was experienced during trans lunar injection when the Mercury ship
was unable t o provide high-speed traj e c t ory dat a be c aus e of a temporary
13-3
problem in the central dat a proce s s or . Some stat i ons als o experienced
t emporary S-band power ampli fier failures during the mi ssion .
Televi sion support provi ded by Network and Jet Propuls ion Lab oratory
facilities was good through out the mi s s ion , part i cularly the s upport by
the 2 10-foot stat i ons at Parkes and Goldstone .
13 . 3 RECOVERY OPERATIONS
Support for tqe primary landing are a in the Paci fic Oce an was pro
·vided by the USS Hornet . Air s upport cons isted of four SH-3D heli copters
from the Hornet , three E-lB aircraft , three Apollo range ins trument at i on
aircraft , and two HC-130 res cue aircraft staged from Hickam Air Force
Base , Hawaii . Two of the E-lB aircraft were de s ignat e d as "Air Bos s " and
the third as a communi c at i ons relay air craft . Two of the SH-3D heli cop
ters carri ed the swimmers and requi red rec overy equipment . The th ird
heli copte r was us ed as a phot ograph i c plat form , and the fourth c arri ed
the decont aminat i on swimmer and the flight s urgeon and was us e d for crew
retrieval .
The command module imme diately went t o the stab le II ( apex down )
flot ation attitude after landing . The upri ghting system returne d the
spacecraft to the s t ab le I attitude 7 minutes 40 seconds later . One or
13-4
two quart s o f wat e r entere d the space craft while i n s t ab le I I . The swim
me rs were then deployed to ins t all the flot at i on collar , and the de c on
t aminat i on swimme r pas s ed the b i ologi c al i s ol at i on garments t o the flight
crew , ai ded the crew i nt o the li fe raft , and decontaminat e d th e exterior
surface of the command module ( s ee s e ct i on 12 . 5 . 2 ) . Aft e r the command
module h at ch was clos ed and de contaminat e d , the flight crew an d de cont am
inat i on swimmer washed e ach other with the decontami nat e s olut i on prior
to being t aken ab oard the re covery heli copt e r . The crew arrived onb oard
the Hornet at 1753 G . m . t . and ent e re d the Mob i le Quarant ine Faci lity
5 minutes l at e r . The first lunar s amples t o b e returned were flown t o
Johnston I s land , placed ab oard a C-141 ai rcraft , an d flown to Houston .
The s ec ond s ample shipment was flown from the Hornet directly t o Hickam
Air Force Bas e , H awai i , approximate ly 6-l/2 hours lat e r and placed ab oard
a range ins t rument at i on aircraft for t rans fer to Hous t on .
The command module was t aken to Ford Is land for deact ivat i on . Upon
comple t i on of de act ivat i on , the command module was shipped t o H i ckam Air
Force B as e , Hawaii an d flown on a C-133 ai rcraft to Hous ton .
Event Time , G . m . t .
July 24
NASA-S-69-3774
13°4 5 '
� Swim 2
� Relay 1
-"' 13°30 ' "'
;;:
0
2
Photo 1
""llliJP ..¥ Air Boss 2
Recovery 1 � ,...., I
....
U S S Hornet
Target point
Onboard computer landing point
I
......
e
Swim 1 I � .,.
" Air Boss 1
Landing point (recovery forces)
13°15' e
The 11 secondary obj e ctives are li sted in table 14-I and are described
in detail in reference 13 .
The s ingle primary obj e ctive was met . All secondary obj e ctives and
experiment s were fully s atis fied except for the following :
These two items were not completely s at i s fied in the manner planned pre
fli ght and a discussion of the deficiencies appear in the following para
graphs . A full as ses sment of the Apollo 11 det ailed obj e ctives and ex
periments will be presented in separate report s .
informat i on for locat ing the landing point using onb oard maps . In addi
tion , thi s informat i on was t o b e t ransmi tt e d t o the Command Module Pi lot ,
who was t o us e the s ext ant in an attempt to locate the landed lunar mod
ule . Further , i f it were not pos sible for the Command Module Pilot to
resolve the lunar module in the s extant , then h e was t o t rack a ne arby
landmark that had a known location relative to the landed lunar module
( as determined by the lunar module crew or the ground team ) .
Toward the end o f the lunar surface stay , the locat i on o f the lande d
lunar module was determine d from the lunar module rende zvous radar track
ing dat a ( confirme d post flight using de s cent photographi c dat a ) . However ,
the Command Module Pilot ' s activiti es di d not permit his attempting another
tracking pas s after the lunar module loc at i on had b een determined accu
rat ely .
The traj ectory parameters of the AS-50 6 launch vehicle from launch
to translunar inj e ction were all close to expected values . The vehi cle
was launched on an azimuth 90 degrees east of north . A roll maneuver was
initiated at 13 . 2 seconds to place the vehi cle on the planned flight azi
muth of 72 . 0 5 8 degrees east of north .
16 . 0 ANOMALY SUMMARY
During the lunar orbit insertion firing , the gas eous nitrogen in
the re dundant s ervi ce propuls ion engine actuat i on system decayed from
2307 t o 1883 psia ( s ee fig . 16-1 ) , indi cating a leak downstream of the
inj e ct or pre -valve . The normal pres sure de c ay as experienced by the
primary system i s approximately 50 ps i a for e ach firing . Only the one
system was affe ct e d , and no performance degradat i on result e d . This actu
at i on system was used during the t ransearth inj ection firing , and no leak
age was dete ct e d .
The fue l an d oxi di zer valves are cont rolled by actuators driven by
nitrogen pres sure . Fi gure 16-2 i s repre s entat i ve of b oth nitrogen con
trol systems . When power is appli ed to the service propulsion system in
preparation for a maneuve r , the inj ector pre-valve is opened ; howeve r ,
·p res sure i s not applied to the actuat ors becaus e the s olenoid control
valve s are clos ed. When the engine is commanded on , the s olenoi d control
valves are opene d , pre s s ure is applie d to the actuat o r , and the rack on
the actuator shaft drives a pini on gear t o open the fuel and oxidi zer
valves . When the engine is commanded off , the solenoi d control valve
vents the actuator and clos es the fuel and oxi di zer valves .
The mos t likely c ause of the problem was . cont aminat i on in one of the
components downstre am of the inj e ctor pre -valve , whi ch i s olates the nitro
gen supply during nonfi ring periods . The injector pre-valve was not con
s idered a problem s ource becaus e it was opened 2 minutes before i gnition
and no leakage oc curred during that peri od . The pos sibility that the
regulator and relief valve were leaking was als o eliminat e d s ince pres
sure was appli ed to these components when the pre-valve was opened .
Both of the s olenoi d control valves in the leaking system had been
found to be contaminated before flight and were removed from the system ,
rebuilt , and suc ce s s fully retested during the acceptance test cycle .
Space craft for Apollo 12 and sub s equent mi ssions have integral fi l
ters installed , and the facility manifolds are more clos ely controlle d ;
therefore , n o further corrective action will b e t aken .
The performan ce of the automat i c pres sure control system indi cated
that one of the two heater elements in oxygen t ank 2 was inoperative .
Dat a showing heater currents for prelaun ch checkout veri fied that b oth
he.at er elements were operati onal through the countdown demonstration
test . However , the current readings recorded during the tank pressur i za
tion in the launch count down showed that one heater in oxygen t ank 2 had
failed. This informat i on was not made known to proper channels for di s
pos iti on prior to the flight , s ince n o specifi c at ion limits were called
out in the t e st procedure .
The laun ch-site test requirements have b een ch anged to spe c i fy the
amperage level to verify that b oth t ank heaters are operati onal . Addi
ti onally , all launch-s ite procedures are being reviewed to determine
whether spe ci fi c at i on limits are required in other are as .
Post fli ght tests showed th at two pins in the terminal b oard ( f'i g .
16-3 ) were loose and c aused intermittent continuity t o the automat i c coils
of' the engine valve . Thi s type f'ai lure has previously been noted on ter
minal b oards manufacture d prior to November 1967 . This b oard was ma.riuf'ac
tured in 1966 .
The intermittent cont act was caus ed by improper clip position rela
tive to the bus b ar counterb ore . The improper pos itioning results in los s
o f' s ome s i de f'orce an d pre cludes proper cont act pre s sure against the bus
b ar . A des i gn change t o the base gasket was made t o insure pos it ively
that the bus b ar is correctly pos itione d.
The loc at i on of' pre-November 1967 terminal b oards has been deter
mined f'rom inst allat i on re cords , and it has been determined that none are
in circuits whi ch would jeopardize crew s af'ety . No act ion will b e t aken
f'or Apollo 12 .
This anomaly i s clos ed.
16-4
An ele ct rolumi nes cent s egment on the nume ri c display of the entry
monitor system velocity c ount er would not i lluminate . The s e gment i s in
dependently swit ch ed through a logi c network whi ch activat e s a s ili con
controlled re cti fi er to bypas s the li ght when not i lluminat e d . The
power s ource i s 115 volt s , 400 hert z .
Thi s anomaly is c l os ed .
During the initial lunar module pre s sur i z at i on , two mas ter alarms
were act i vat e d when the oxygen flow rate was decre asing from full-s cale .
The s ame condi ti on h ad b een ob served s everal times during alt it ude
chamber tests and during s ubse q_uent troub leshoot i ng . The cause of the
prob lem could not be i dent i fi e d b e fore launch , but the only c ons eq_uence
of the alarms was the nui s ance factor . Fi gure 16-4 shows the b as i c ele
ments of the oxygen flow s ens ing circui t .
Note i n figure 16-4 that i n order for a mas ter alarm to o ccur , relay
Kl mus t hold in for 16 s econds , after which t ime relays K2 and K3 wi ll
clos e , activat ing a mas ter alarm .
16-5
The filter c apacitor was open during postflight tests , an d the mas ter
alarms were dupl i c at e d with s low , de cre as ing flow rates .
There has been no previous fai lure h i story of these metali zed MYlar
capacitors as sociated with th e flow s ens ors . No corre ctive act i on i s
requi re d.
Two valves having the nominal lat ching force of 7 pounds were sele cte<i
for shock testing . It was found that shocks of 80g for 10 milli s econds
to shocks of lOOg for 1 mi lli second would close the valve s . The latching
forces for the valve s were reduced to 5 pounds , and the valves were
shock tested again . The shock require d to close the valve s at this re
duced lat ching force was 54g for 10 milli s econds and 75g for 1 mi lli s ec
ond. After completi on of the shock testing , the valves were exami ned and
teste d , and no degr adat i on was not e d . Higher shock levels malf h ave been
experienced in fli ght , and further tests will be c onducted .
A review o f the checkout procedures indi cat e s that the lat ching
force can be degrade d only if the procedures are not prope rly implemented ,
such as the appl i c at i on of reverse current or ac to the circui t . On
Apollo 12 a special test has indi c at e d that the valve lat ching force has
not b een degr ade d .
16-6
An odor s imi lar to burned wire insul at i on was detect e d in the tunnel
when the h at ch was first opene d . There was no evi den ce of di s colorat i on
nor indi c at i ons of overheating of the ele ctrical ci rcuits when examined
by the crew during the flight . Several other s ources o f the odor were
invest igated , including burned part icles from t ower j et t i s on , out gas s ing
of a s i l i c one lub ri c ant us ed on the hatch s e al , and outgas s ing of other
components use d in the tunnel are a . Odors from these s ources were re
produced for the crew to compare with the odors dete cted during flight .
The c rew st ated that the odor from a s ample of the docking h at ch abl at or
was very s imi lar t o that dete cted in flight . Apparent ly , removal of the
outer insulat i on ( TG-15000 ) from the h at ch of Apollo 11 ( an d s ubsequent )
resulted in h i gher ablator t emperatures and , there fore , a larger amount
of outgas s ing odor than on previ ous fli ght s .
An alysis of as s oc i ated dat a indi cat e d that the oxygen flow was norm
al , but that the indi c at e d flow rate was negat i vely b i as e d by approximately
l . 5 lb /hr . Postflight tests of the t rans duce r confi rme d this b i as , and
the cause was as s oc i at e d with a change in the heater winding resi stance .-
within the flow s ens or bri dge ( fi g . 16-5 ) . The re s i stan ce of the heater
had incre as e d from 1000 ohms to 1600 ohms , changing the temperature of the
hot wire e lement whi ch s upplies the reference volt age for the b alan ce of
the bri dge . Further testing to determine the c ause of the res is t ance
change is not pract i cal because of the minute s i ze of the potted res istive
element . Depotting of the element would destroy avai lab le evidence of
the cause of failure . Normally , heater re s i st an ce changes h ave occurred
early in the 100 -hour burn-in peri od when heater stability is achieve d .
16-7
Spacecraft 110 and lll were exami ned , and it was found that a clove
hitch was erroneous ly us ed on thos e vehicles als o .
An apparent anomaly exi sts with the glycol temperature control v alve
or the relat e d temperature control system. Temp erature of the water/
glycol enteri ng the evaporator i s normally maintai ned above 42° F by the
glycol temperature control valve , which mixes hot water/glycol with water/
glyc ol returning from the radiators ( s ee fig . 16-7 ) . As the radi ator out
let t emperature de creas es , the temperature control valve opens to allow
16-8
more hot glycol to mix wit.h the cold fluid returning from the radi ator
to maintain the evaporator inlet temperature at 42° to 48° F . The con
trol valve starts to close as the radiator outlet temperature increas e s
and closes completely at evaporator inlet temperatures above 4 8 ° F . I f
the automati c temperature control system i s lost , manual operation o f
the temperature control valve i s available by deactivating the automati c
mode . This is accompli shed by pos itioning the glycol evaporator tempera
ture inlet switch from AUTO to MANUAL , whi ch removes power from the con
trol circuit .
The control valve was remove d from the spacecraft , dis as s embled , and
inspecte d . A bearing within the gear train was found to have its retainer
dis engaged from the rac e . The ret ainer was interfering with the worm gear
trave l . The caus e of the failure of the retainer is under investigation .
Photographic dat a were obt ained of the service module entering the
earth ' s atmosphere and dis integrating near the command module . Pre flight
predictions indicated the s ervice module should have skipped out of the
earth ' s atmosphere and entered a highly ellipti cal orbit . The crew ob
s erved the s ervi ce module about 5 minutes after separation and indi cated
the reaction control thrusters were firing and the module was rotating
about the X plane . ·
Bas e d on the film , crew observation of the s ervi ce module , and data
from previous mis s i ons , it appears that the s ervice module did not per
form as a stable vehicle following command module /service module separa
tion . Calculations using Apollo 10 data show that it is poss ible for the
remaining propellant s to move axially at frequencies approximately equal
to the pre cess ional rate of the service module spin axis about the X body
16-9
axis . This effect causes the movement to res onate , and the energy trans
fer between the rotating vehicle and the propellants may be sufficient to
cause the s ervi ce module to go into a flat spin about the Y or Z axis and
become unst able .
16 . 2 LUNAR MODULE
The crew reported shortly after lunar landing that the mis s ion timer
had stopped . They could not restart the clock at that time , an d the power
'to the timer was turned off to allow it to cool . Eleven hours later ,
the timer was restarted and functioned normally for the remainder of the
mis s ion .
New mis sion timers and event timers which will be mechanically and
electrically interchangeable with present timers are being developed.
These new timers will use integrated circuits welded on printed circuit
boards instead of the cordwood construction and include design changes
as s ociated with the other timer problems , s uch as cracked glass and elec
tromagneti c interference sus ceptibility . The new t imers will be incorpo
rated into the spacecraft when qualifi c ation testing is complete .
On future mis s i ons , the s olenoid valve ( fig . 16-10 ) will be closed
prior to fuel venting and opened s ome time prior to lift-off . This will
prevent freezing of fuel in the heat exchanger and will allow the s uper
critical helium tank to vent later. The helium pres sure rise rate after
landing is approximately 3 to 4 psi /hr and constitutes no constraint to
presently planned miss ions . Appropriate changes to operational procedures
will be made .
Shortly after the lunar module as cent , the crew reported that the
measurement of c arbon dioxide partial pressure was high and erratic . The
secondary lithium hydroxide canister was selected, with no effect on the
indication . The primary canister was then reselected , and a c aution and
warning alarm was activated .
1 6 ll
-
To preclude water being introduced into the sensor from the drain
tank , the vent line will be relocated to an exi sting boss upstream of the
fans , effective on Apollo 13 ( see fig . 16-11 ) .
For future missions , the correct vehicle blockage and multipath con
ditions will be determined for the predicted flight traj ectory . Opera
tional measures can be employed to reduce the probability of this problem
recurring by selecting vehicle attitudes to orient the antenna away from
vehicle blockages and by selecting vehicle attitude hold with the antenna
track mode switch in the SLEW or manual position through the time periods
when this problem may occur .
Five computer program alarms occurred during des cent prior to the
low-gate phase of the traj ectory . The performance of guidance and con-
trol functions was not affected.
The alarms were of the Executive overflow type , which s ignify that
the guidance computer cannot accomplish all of the dat a processing re
quested in a computation cycle . The al arms indi c ated that more than
10 percent of the computational c apacity of the computer was preempted
by unexpected counter interrupts of the type generated by the coupling
data units that interface with the rendezvous radar shaft and trunnion
res olvers ( s ee fig . 16-14 ) .
The computer is organi zed such that input /output interfaces are
servi ced by a central proces s or on a time-shared basis with other pro
cessing functions . High-frequency dat a , such as accelerometer and cou
pling data unit inputs , are processed as counter interrupts , which are
as signed the highest priority in the time-sharing sequence . Whenever
one of these pulse inputs is receive d , any lower priority computati onal
task being performed by the computer is temporarily suspended or inter
rupted for 11 . 72 mi croseconds while the pulse is processed, then control
is returned to the Executive program for resumption of routine operations .
res erve another memory storage area for its use . When the Executive
program is requeBte d to s chedule a j ob and all loc ations are as signed ,
a program alarm i s displayed and a software rest art i s initiated. A
review of the j obs that can run during des cent leads to the conclus ion
that multiple s cheduling of the s ame j ob produced the program alarms .
The cause for the multiple s cheduling of j obs has been i dent i fied by
analyses and s imulations to b e primarily counter interrupts from the
rendezvous radar coupling dat a unit .
The interrupts during the powered des cent resulted from the con
figuration of the rendezvous radar I coupling data unit I computer inter
face . A s chematic of the interface is shown in figure 16-14 . When the
rendezvous radar mode switch is in the AUTO or SLEW position , the excit
ation for the radar shaft and trunnion res olvers is supplied by a 28-volt ,
800-hertz signal from the att itude and trans lation control assembly .
When the switch is in the LGC pos ition , the positi oning of the radar
antenna is controlled by the guidance computer , and the res olver exci ta-
t ion is suppli e d by a 28-volt , 800-hertz s ource in the primary gui dance
and navigation system . The output s i gnals of the shaft and trunnion
res olvers interface with the coupling dat a unit s regardles s of the excit
ation s ource . The attitude and translation control as s embly voltage is
locked in frequency with the primary guidance and navigation syst em
voltage through the system ' s control of the PCM and timing electroni cs
frequency , but it is not locked in phas e . When the mode swi tell is not
in LGC , the attitude and translation control ass embly voltage is the
s ource for the res olver output signals to the coupling data units while
the primary gui d!mce and navigation system 800-hertz voltage is us ed as
· a reference voltage in the analog-to-digital convers ion portion of the
coupling data unit . Any di fference in phase or amplitude between the
two 800-hertz voltages will cause the coupling data unit to recognize a
change in shaft or trunni on position , and the coupling dat a unit will
"slew" ( digit ally ) . The "slewing" of the data unit results in the un
des irable and cont inuous transmi s s i on of pulses representing incremental
angular changes to the computer . The maximum rate for the pulses is
6 . 4 kpps , and they are processed as counter interrupts . Each pulse re
ceived by the computer requires one memory cycle time ( 11 . 7 micros econds )
to proces s . I f a maximum of 12 . 8 kpps are received ( two radar coupling
data units ) , 15 percent of the available computer time will be spent in
--
process ing the radar interrupts . ( The computer normally operates at
approximately 90 percent of capacity during peak activity of powered
des cent . ) When the capacity of the computer is exceeded , s ome repeti
tively scheduled routines will not be completed prior to the st art of
the next computation cycle . The computer then generates a s oftware re
start and displ�s an Executive overflow alarm .
16-14
The me aningle s s counter interrupts from the rende zvous radar coupl
ing data unit will not b e proces s e d by the Luminary lB program us ed on
future mi s sions . When the radar i s not powered up or the mode swi t ch i s
not i n t h e LGC pos ition , the dat a units will b e zeroe d , preventing counter
interrupts from being generat e d by the radar coupling dat a units . An
additional change will permit the crew to monitor the des cent without
requiring as much computer time as was require d in Luminary lA .
On Apollo 12 and sub sequent vehicles , the b acteria filter will not
be use d , thus reducing the t ime for decompres sion from about 5 minutes to
less than 2 minut e s . In addition , the alt itude chamber test for Apollo 1 3
included a parti al cab in vent pro cedure whi ch verifi ed sat i s factory valve
as s embly operat ion without the bacteria filt er ins talled.
In order to ens ure proper operation under all conditions , for future
mi s s i ons a prelaw�ch test will activate all segments , then the intensity
will be varied through the full range while the display is obs e rved for
faults .
All analysi s and laboratory testing to date indi c ates that the voice
breakup experienced during the extravehi cular activity was not an inherent
system design problem . Testing has shown that any voice which will key
the extravehicular communic ation system will also key the lunar module
relay i f the s ensitivity control is s et at 9 .
16-16
When the lunar module voi ce-operat e d keying circuit i s properly ad
juste d , any s ignal that keys the extravehi cular communi cations system
will al s o key the lunar module relay . There are indications that the
lunar module voi ce keying s ensitivity was s et below maximum , as evidenced
by the relayed voi ce breakup experi enced by the Lunar Module Pilot ( s ee
s ection 16 . 2 . 8 ) . Therefore , it would have been pos sible for the extra
vehicular communi cat ions system to be keyed by breathing or by suit air
flow without this b ackground noi s e b eing relaye d by the lunar module .
Howeve r , the uplink turnaround voi ce could provide the additional lunar
module received audio s ignal level to operate the voice -operated keying
circuits , permitting the s ignal to be returned to the earth . The crew
indi c at e d that the voice-operated keying circuits in the extravehi cular
16-17
c ommuni cations system were activated by suit air flow for s ome p os it i ons
of the head in the helmet . Both voi ce-operated keying circuits were als o
keyed by bumping or rubbing of the communi cati ons carrier against the
helmet . The random echo problem is inherent in the communi cation system
design , and there does not appear to be any practical w� to eliminate
random voice keying or significantly reduce acoustical coupling in the
c ommuni cations carrie r .
During post flight tests , the recorder functioned properly for the
first 2 hours of operation . Then , the voi c e channel failed and recorded
no voice or background nois e , although timing and reference tones were
recorded properly . This failure does not duplicate the flight results ,
indicating that it di d not exist in flight .
Preflight data from the launch s it e che ckout procedure show that
b oth the timing inputs and the internally generated reference frequency
were not within speci fi c at ion tolerances and may be indicative of a pre
flight problem with the system . The procedure did not spe c i fy acceptable
limits but has now been c orre cte d .
The most probable cause of the damage w as impact of the oxygen purge
system ( aft edge ) during preparat i on for extravehi cular activities ; s uch
impact was demonstrated in s imulati ons in a lunar module .
The switch used to monitor the quad 2 aft-firing engine ( A2A) exhib
ited s low response to j et driver commands during most of the mis s i on .
During an 18-minut e period just prior t o terminal phase init i ation , the
switch fai led to respond to seven cons ecutive minimum impulse commands .
This resulted in a master alarm and a thruster warning flag , which were
reset by the crew . The engine operated normally , and the switch failure
had no effect on the mis s i on . The crew did not attempt any inves t iga
tive procedures to determine whether the engine had actually faile d . A
section drawing of the switch i s shown in figure 16-18 .
The crews for future mis s i ons will be briefed to recognize and
handle similar situations .
After the l1mar module achieved orbit , water began to enter the
Commander ' s suit in spurts ( estimated to be 1 tablespoonful ) at about
1-minute intervru.s . The Commander immediately selected the secondary
water separator , and the spurts stopped after 15 to 20 minutes . The
spurts entered the suit through the suit half vent duct when the crewmen
were not wearing their helmets . The pressures in all liquid systems
which interface ;rith the suit loop were normal , indicating no leakage .
The possible sources of free water in the suit loop are the water
separator drain tank , an inoperative water separator , local condensation
in the suit loop ,. and leakage through the water separator selector valve .
( s ee fig . 16-11 ) .. An evaluation of each of these poss ible s ources indi
cated that leakage through the water separator selector valve was the
most probable s ource of the free water .
The crew reported thrust chamber as sembly warning flags for three
engine pairs . Quad 2 and quad 4 warning flags for system A occurred
s imultaneously during lunar module station-keeping prior to descent
orbit insertion . Quad 4 flag for system B appeared shortly thereafter
and als o twice just be fore powered des cent initiation . The crew believed
these flags were accompanied by master alarms . The flags were reset by
cycling of the cauti on and warning electroni cs circuit breake r . Suffi
cient data are not available to confirm any of the reported conditions .
The first two possible causes are highly unlikely because simultane
ous multiple failures would have to occur and s ubsequently be corrected.
The third poss ible cause is the most likely to have occurred where a
single point failure exists . Ten o f the s ixteen engine pressure switch
outputs are conditioned by the ten buffers in one module in the signal
conditioner electronics as sembly ( fig. 16-20 ) . This module is supplied
with +28 V de through one wire . In addition , the module contains an
os cillator which provides an ac voltage to each of the ten buffers . If
either the +28 V de is interrupted or the os cillator fails , none of the
ten buffers will respond to pres sure switch closures . If engines mon
itored by these buffers are then commanded on , the corresponding warning
flags will come up and a master alarm will occur .
The cable for the lunar s urface televis ion camera retained its coiled
shape after being deployed on the lunar surface . Loops result ing from
the coils repres ented a potent i al tripping hazard to the crew .
For future mi s sions , the male half of the connector has been replaced
with one whi ch has a coupling lock ring with a positive rot at ional posi
t ion with the connector shell and can b e grasped for firm alignment of
the two halves . The ring is then rotated 90 degrees to capture and lock .
In additi on , e as ier insertion has been att ained with coni cal t ipped con
t act pins in place of hemi spherical tipped pins .
The force required to clo s e the s ample return containers was much
higher than expecte d . This high clos ing force , coupled with the inst
ability of the des cent stage work t able and the lack of adequate reten
tion provisions , made clos ing the containers very diffi cult .
Because of the cont ainer s eal , the force required to close the cover
reduces with each closure . The crew had extens ive training with a s ample
return container which had been opened and closed many times , res ulting
in clos ing forces lower than the maximum limi t of 32 pounds .
16-22
The container used for the flight had not been exercised as had the
container used for training . In addition , the cleaning procedures us ed
by the contractor prior to delivery removed all lubri cant from the latch
linkage sliding surfaces . Tests with similar containers have shown that
the cleaning procedure caused an increas e i n the clos ing force by as much
as 2 4 pounds .
A technique for burnishing on the lubri c ant after cleaning has been
incorporat e d . As a result , containers now being delivered have closing
forces no greater than 25 pounds .
NASA-S-69-3 7 7 5
;
24 00 I
I
I
I
Primary
I
r---....
I
I I
:
2000 -loo...
I
I I Secondary
I I
I I
I
�
-
Q)
1 20 0
I
J :
:::s
(/)
(/)
Q)
�
a.
800
400
0
75:48 75:50 75:52 75:54 75:56 75:58 76:00 76:02
T ime , hr:m i n
I-'
0\
I
F igure 1 6 - 1 . - N i trogen pressure dur i ng i n itial lunar orbit i n sertion firing . 1\)
'-"
NASA-S-69-3776 I-'
0\
�
I
Fuel
(i�a�:� Injector
pre-valve
I n let
Outlet
�'"""'ti�\�\ [\ ,II��;%1�-- Seals
(most probable
leak source)
liJ Oxidizer
� Fuel
D Gaseous
nitrogen
NASA-S-69-3 7 7 7
Terminal
board 9 19
( 1 6 gage pins)
+ 2 8 V de - Oxid i zer
: :
- Fuel : :
r I : :
r
I D irect c o i l s
•
..-- I
- Oxid i zer
I
:. to
: Failed
.. Fuel :
: : operate
+ 2 8 v dc - � Em
� '/ En
-�
Automati c coi ls
I ntermittent p m s
Reset
To
master
28 V de alarm
1 6-second t---""'"'"'
de lay K2
Vo ltage
Oxy gen 1----f Out put 1---.....---t leve l Relay
flow amp l ifier detector driver
sensor
T
F i gure 16-4 . - Oxygen flow sens ing circu i t .
N ASA-S- 69 -37 79
Oxygen
flow
� Sensor probe
U..-.:o��
Heating e lement
I
I
_ J
To te lemetry and
on board d i s p lay
NASA-S-69-3 7 8 0
Proper
Improper
NASA- S -69-3 78 1
Rad i ator outlet Primary evaporator
temperature sensor temperature sensor
w Primary 0g{�
Warm
� Evaporator in let t
temperature
sensor
Primary Cold plates and
radiators heat exchangers
�
Hot .. -
Pump Hot �
F i gure 16-7 . - Pri mary water g l ycol coo lant loo p . 1-'
0\
I
1\)
\0
NASA -S -69-3782
-- I I
N o r m a l perfor mance
--1 -- --'1
60 1--+-----1-1--
n "'
Temperat ure control
valve unders hoot
r--
.?\.
A \�-'' l
'
'JI
j 'l\
I
I
I
I
I
\�
50 1--t--Hi-+-----1
' I J f/-f- Expected temperatu re
V:···+/ due to norma l temperat ure
I
1, \ y
�
j
I
'\ V
L.L.
0
40
control valve modu lation
!
\
\ I
I i
� I
'\_'-'"'"�"
:J
1\ I \
I outlet
�
'\
I
\
\
1---,�'-+-----'1
a.>
� 30
' I
,_
I
'\
\
I
I
I
I
\
I I \
I Radiator out l et � I I \
I I \
20
\._/, I
I
\
\
\
I
I
10 1----+----1 1'-/
0 L----'----_J
101:45 102: 15 102:45 104:30 105:00 105:30 106:00 ll3:30 114:00 1 14:30 115:00 115:30
Time, hr: m i n
NASA-S -69-.378.3
C i rcuit board
Potting
mater ial
NASA-S-69-3 7 84
To vent
To oxi d i zer tank
_. ��======�
Heat
exchanger
NASA-S-6 9 -3 785
To water
Lunar
Modu le--o.. �===�=::J
__. Commander's
system P i lot ' s ' s u it
Lith i u m
hydroxide
can i ster
--
NASA-S-69-3786
..
:
'{.; ':"" : .
.. .
: :}
250
230 : l : ':
""'\.,.·.;;:::::: :::;:;j:;:;:;:::=:r::;;;.
'-- S pacecraft operational
•
':>:'::Jf:
l-l--+-----l--4l�
C;i2
data book boundary
8$Mc__.J--_+--- --1 prior to flight
·
/
/ blocking diagram I . _
I
tr Antenna
1+--+-'l---A-- pointing +--+--f--+1
\/
an le
\
\ - ing �
.�
D irect path
M u lti path
(reflected from
lunar surface)
Moon
......
0'.
F i gure 16-13 . - E xample of multi path .
I
w
Vl
NASA-S-69-3788
1-'
0\
I
w
0\
Frequency Frequency
sync PCM and sync
Attitude and trans lation
1--;..._ --1 timing
contro l assemb ly ( 1 . 6 kpps) e lectron ics ( 1 . 0 2 4 M H z)
-
2 8-vo lt, 8 0 0 H z
L
J- AUTO
e--G-:---
Primary
g
2 8-vo lt, 8 0 0 H z reference
u i dance and
S EW L -------t-----------1 nav igati on
C
s ystem
Mode switch
F igure 16- 1 4 . - I nterfaces from rendezvous radar antenna to primary g u idance system .
16 - 3 7
NASA-S-69-3 7 8 9
r
•
·.·. . .
B
� - -�===================�- -
. '
'
NASA-S-69-37 9 0
.-
�-
NASA- S -69-3 7 9 1
Acoustical
tube -
. . .. . . . . . .
Earpiece with
microphone and .. I
earphone dri vers
mo lded in___ _,
!�
Commu n i cat ion system conne �
--
i
I�
I
I
!
Contacts Pressure
Sense por
�i 11
sw itch
Prope l lant
Passage b locked va l ves
by contam ination
I
--
- Hand le
- - ·
Reset
A2D firing
command
Chamber 8 uffer
module
A2D chamber
pressure switch
A2A firing
command
To malfunction
A2A chamber logic circuits
pressure switch 1A
18
28
t
2A isolation 3A
valve closed Reset 3B
4A
� -+==;---]++2i2Bi'VV"dc� �
:
4B
::-
at this
point switch
Buffer 1
Figure 16-20. - Reaction control system malfunction detection circuits (caution and warning systemI.
' (
' '
N ASA-S-69-3795
f:J S ystem A
i System B
+X
1 left
3 ri ght
4 down
I-'
0\
I
F i gure 16- 2 1 . - Reaction contro l system geometry .
.,...
w
NASA-S-69-37 96
F i gure 16- 22 . - Connector between remote control un it and portable l i fe su pport system .
' '
17-l
17 . 0 CONCLUSIONS
The Apollo ll mis sion , including a manned lunar landing and surface
exploration , was eonducted with skill , pre cision , and relat ive ease . The
excellent performance of the spacecraft in the preceding four flight s and
the thorough planning in all aspects of the program permitted the safe and
efficient execution of this mis s ion . The following conclus ions are drawn
from the information contained in this report .
9 . The Misuion Cont rol Center and the Manned Space Flight Network
proved to be adequate for controlling and monitoring all phases of the
flight , including the des cent , surface activities , and ascent phases o f
the mission .
A-1
Very few change s were made to the Apollo 11 space vehicle from the
Apollo 10 con figurat ion . The launch e s c ape system and the spacecraft /
launch vehi cle adapter were i dent i cal to thos e for Apollo 10 . The few
minor changes t o the command and service modules , the lunar module , and
the Saturn V lam1ch vehicle are di s cus s ed i n the following paragraphs .
A de s cripti on of the ext ravehi cular mobi li ty unit , the lunar surface ex
periment equipment , and a li sting of spacecraft mass properties are also
pre sent e d .
The insulation in the are a of the command module forward hat ch was
modi fied t o prevent the flaking which oc curre d during the Apollo 10 lunar
module pre s suriz at i on . The fee dback circuit i n the high gain antenna was
s lightly changed to re duce s ervo dither. In Apollo 10 , one of the three
ent ry b atteries was modi fi ed t o make us e of cellophane s eparators . The
flight results prove d thi s materi al superior to the Fermion-type previ
ously used an d for Apollo 11 all three entry b atteries had the cellophane
s eparat ors . The b at tery chargers were modified to produce a higher charg
ing capacity . The s econdary bypass valve s for the fuel cell coolant loop
were changed from an angle-cone s e at de sign ( b lo ck II ) to a s ingle-angle
s e at (block I ) to re duce the possibility of parti culate cont aminat i on .
As a replacement for the wat er/gas s eparat i on b ag which proved ineffect ive
during Apollo 10 , an in-line dual membrane s eparat i on devi ce was added t o
b oth the water gun and the food preparat i on unit .
A.2.1 Structures
The most s i gnifi cant structural change was the adde d provi s i ons for
the functi on e� e arly Apollo s ci ent i fi c experiment package and the modular
equipment s towage as semb ly , b oth of which hous ed the experiments and tools
used during the lunar surface activities . Another change was the addition
of the react i on control system plume de fle ctors .
A. 2 . 2 Thermal
A change from Kapton to Kel-F was made to the des cent stage b as e
heat shield t o preclude the possibility o f inte rference with the landing
radar . Als o , i nsulat i on was adde d to the landing gear and probes to ac
commodate the requirement for des cent engine firing unti l touchdown .
A.2 . 3 Communicat i on s
The maj or modi fi cat i on s t o the communi cat i ons syst ems included the
addition of an extravehi cular activity antenna for lunar communicat i ons
between the crew , and the lunar module , an d an S-band erect ab le antenna
to permit communicat i ons through the lunar module communi cat i ons system
( fig . 16-16 ) while the crew was on the surface .
A.2 4
• Gui dance and Cont rol
The maj or di fference in the guidance and control system was the re
design of the gimb al drive actuator to a constant damping system rather
than a brak e . This was re des igned as a result o f the brake failing i n
b oth the disengage d an d engage d positi on . This change also requi re d mod
i fi c at i on of the des cent engine control as s emb ly and the phas e corre cting
network t o e liminat e the possibility of inadve rtent caut i on and warning
alarms .
The exterior tracking light had improvements in the flash head and
in the puls e-forming network .
The pushbut t ons for the dat a entry and display as sembly were re
wired to pre clude the e rroneous cauti on and warning alarms that had
oc curred on the Apollo 10 flight .
The gui dance and n avigat i on optics system was modifi ed by the addi
tion of Teflon locking rings t o the s extant and the s canning teles cope
to prevent the rot at i on of eye guards under zero-g conditions .
The inj ector filter for the as cent propuls ion system was modified
because the fine mesh in the original filter was causing a change in the
mixture ratio . An additional change was the incorporation of a light
weight thrust chamber .
In the · environmental control system relay box in the oxygen and cabin
pressure control section , a pressure transducer was replaced by a suit
pressure switch to improve reliability .
A.2 .7 Radar
'""'-.
Circuit breakers were adde d for the ab ort electroni cs as sembly and
the uti lity light . A circuit breaker was adde d for the ab ort electronics
as sembly to prote ct the de bus , and another circuit breaker was adde d to
accommodate the t rans fe r of the uti lity light to the de bus to provi de
redundant light .
The circuit breaker for the environmental control system suit and
cabin repre s sur i z at i on functi on was delete d i n conjunct i on with the modi
fi cat i on of the suit cooling as sembly . In addition , a low-level caution
and warning indi c at i on on the s e condary water glycol accumulator has been
provi ded .
Mas ter alarm funct i ons which were eliminat e d include the des cent
helium regulator warning prior t o pre ssuri z at i on with the de s cent engine
control as semb ly ; the re acti on cont rol system thrust chamber as semb ly
warning with quad circuit breakers open ; the rendezvous radar caut i on when
placing the mode s ele ct swit ch in the auto-track posit i on ; and the deleti on
of the react i on control system quad temperature alarm.
Cauti on and warning funct ions which were deleted include the landing
radar velocity " dat a no-good" and the de s cent propellant low-level quantity
which was change d to a low-leve l quantity indi c at i on light only .
A modi fi cat i on was made to the engine stop swit ch lat ching me chanism
to insure posit ive lat ching of the swit ch .
Additional stowage i ncluded prov�s �ons for a s e cond Hass elb lad
camera , a t ot al o f two port ab le life support systems and remot e control
unit s , two pairs of lunar ove rshoe s , and a feedwater collect i on b ag . The
C omman der had an attitude controller as sembly lock me chani sm adde d.
A-5
The extravehicular mobility unit provi des life support in a press ur
ized or unpressurized cabin and up to 4 hours of extravehicular life s up
port .
The liquid cooling garment was worn by the crewmen whi le in the lunar
module and during all extravehicular activity . It provi ded cooling during
extravehicular and intravehi cular activity by abs orbing body heat and trans
ferring excessive heat to the sublimator in the portable life support sys
tem. The liquid cooling garment was a one piece , long sleeved, integrated
stocking undergarment of netting material . It consisted of an inner liner
of nylon chiffon , to facilitate donning , and an outer layer of nylon Span
dex into which a network of Tygon tubing was woven . Cooled water , supplied
from the portable life support system or from the environmental control
:;>ystem , was pumped through the tubing .
The pressure garment assembly was the basic pres sure vessel of the
extravehi cular mobility unit . It would have provided a mobile life sup
port chamber if cabin pressure had been lost due to leaks or puncture of
the vehicle . The pres s ure garment ass embly cons isted of a helmet , tors o
and limb suit , intravehi cular gloves , and various controls and instrumen
tation to provide the crewman with a controlled environment .
The torso and limb suit was a flexible pressure garment that encom
passed the entire body , except the head and hands . It had four gas con
nectors , a multiple water receptacle , an electrical connector , and a urine
transfer connector . The connectors had positive locking devices and could
b e connected and dis connected without assistance . The gas connectors com
prised an oxygen inlet and outlet connector , on each side of the suit front
tors o . Each oxygen inlet connector had an integral ventilation diverter
A-6
valve . The multiple water receptacle , mounte d on the suit tors o , s erved
as the i nterface b etween the liquid cooling garment multiple water conne c
t o r an d port ab le li fe s upport system multiple wat e r conne ctor and the en
vironmental control system water supply . The pres sure garment as semb ly
electri c al connect or , mat e d with the vehi cle or port ab le li fe s upport
system elect ri cal umbi li cal , provide d a commun i cat ions , instrument at i on ,
and power interface to the pre s sure garment as semb ly . The urine trans fer
connector was us ed to transfe r urine from the urine colle ct i on trans fe r
as sembly to the waste management system.
The urine tran s fer conne ctor on the suit right leg , permitted dumping
the urine colle ct ion bag without depre s suri zing the pres sure garment as
s embly . A pres sure garment as s emb ly pre ssure relie f valve on the suit
s leeve , ne ar the wrist ring , vent e d the suit in the event of overpressuri
z at i on . The valve opened at approximat ely 4 . 6 psig and re seat e d at 4 . 3
psig. I f the valve di d not open , it could have been manually overri dden .
A pres sure gage on the othe r s le eve indi cated suit pre s s ure .
A. 3 . 4 Helmet
The helmet was a Lexan ( polycarbonate ) shell with a bubble type visor ,
a vent pad as s emb ly , an d a helmet att aching ring . The vent pad assemb ly
permitted a constant flow of oxygen over the inner front surface of the
helmet . The crewman could turn hi s head within the helmet neck ring are a .
The helmet di d . not turn independently o f the torso and limb suit . The
helmet had provis ions on e ach s i de for mounting an extravehi cular vi s or
assemb ly .
The communi cat i on s carri er was a polyurethane foam headpi ece with
two independent e arphone s and mi crophones whi ch were connected to the
suit 21-pin communi cat i ons electri cal conne ct or . The communi cat i ons c ar
rie r could be worn with or without the helmet during intravehi cular opera
tions . It was worn with the helmet during ext ravehi cular operations .
The integrat e d thermal mi cromete oroid garment was worn over the pre s
sure garment assemb ly , an d prot e cted the crewman from harmful radi at i on ,
heat t rans fe r , and mi cromete oroi d activity. The i ntegr ated thermal mi c
rometeoroid garment was a one piece , form fitting multi layere d garment
that was lace d over the pres sure garment as semb ly and remained with it .
The extravehi cular vi s or as s embly , glove s , and boots were donne d s epar
ately . From the outer layer in , the integrat e d thermal micrometeoroid
A-7
garment consi sted of a prot e ctive cover , a micromet eoroi d-shielding laye r ,
a thermal-barri er blanket ( multiple layers of aluminized Mylar ) , and a
prot e ct ive liner .. A zipper on the integrated thermal micromet eoroi d gar
ment permitte d conne cting or dis conne cting umbi li c al hoses . For extra
vehi cular activity , the pressure garment as semb ly gloves were replaced
with the extravehicular glove s . The extravehi cular gloves were made of
the s ame material as the integrated thermal mi crometeoroid garment to per
mit handling intens ely hot or cold obj e ct s out si de the cabin and for pro
tect i on against lunar temperature s . The extravehicular boots were worn
over the pre s sure garment as semb ly boots for extravehicular act ivity .
They were made of the s ame materi al as the integrat e d thermal mi cromet eo
roid garment . �['he s oles had additional insulat i on for prote ction agai ns t
intens e t emperature s .
The portable life support system ( s ee figure A-2 ) contained the ex
pendable materi als and the communi c at i on and telemetry equipment requi red
for extravehi cular operat i on . The system s upplied oxygen to the pressure
garment as semb ly and cooling water to the liqui d cooling garment and re
moved s olid and gas cont aminant s from returning ox:rge n . The port ab le
life support system , att ached with a harnes s , was worn on the b ack of
the sui ted crewman . The tot al system cont ained an oxygen ventilat ing
circuit , water J;eed and liqui d transport loops , a primary oxygen supply ,
a main power supply , communi cation systems , di splays and related s ensors ,
switches , and controls . A cover encompas sed the as semb led unit and the
top portion supported the oxygen purge syst em .
The remote control unit was a display and control unit chest -mounted
for e asy ac ce s s . The controls and displays consi sted of a fan swit ch ,
pump switch , space-suit communi c at ion-mode switch , volume control , oxy
gen quantity indi cator , and oxygen purge system actuator .
A-8
The oxygen purge system provided oxygen and pres sure control for
certain extravehicular emergencies and was mounted on top of the portable
life support system . The system was self-contained , independently pow
ered , and non-rechargeable . It was capable of 30 minutes of regulated
( 3 . 7 ± 0 . 3 psid) oxygen flow at 8 lb /hr to prevent excessive carbon di
oxide buildup and to provide limited cooling . The system cons isted of
two interconnected spherical 2-pound oxygen bottles , an automatic temper
ature control module , a pressure regulator assembly , a battery , oxygen
conne ctors , and the necess ary checkout instrumentation . The oxygen purge
system provided the hard mount for the VHF ant enna.
Earth stations that can beam lasers to the experiment include the
McDonald Observatory at Fort Davis , Texas ; the Lick Observatory in Mount
Hamilton , Cali fornia ; and the Catalina Station of the Univers ity of Ari
zona. Scientists in other countries also plan to bounce laser beams off
the experiment •
A-9
The primary aim of the Apollo lunar field geology experiment was to
collect lunar s runples , and the tools described in the following para
graphs and shown in figure A-5 were provided for this purpose .
Launch vehicle AS-506 was the sixth in the Apollo Saturn V series
and was the fourth manned Apollo Saturn V vehi cle . The AS-506 launch
vehicle was configured the s ame as AS-505 , used for the Apollo 10 mis
sion , except as des cribed in the following paragraphs .
Spacecraft mass properties for the Apollo 11 miss ion are summari zed
in t able A-I . These data represent the conditions as determined from
post flight analyses of expendable loadings and usage during the flight .
Variations in spacecraft mass properties are determined for each signifi
cant mis s i on phase from lift-off through landing . Expendables usage is
based on reported real-time and postflight dat a as presented in other
s ections of this report . The weights and centers of gravity of the indi
vidual command and service modules and of the lunar module as cent and de
s cent stages were measured prior to flight , and the inertia values were
calculated , All changes incorporated after the actual weighing were
monitore d , and the spacecraft mass properties were updated.
A-ll
Product of inertia,
Center of gravity , i n . Moment ot inertia, slug-ft 2
Weigh t , slug-ft2
Event
1b
lyy 1z I 1x ry
z XY z z
X Y Z I
A A A xx
Lif't-off 109 666 . 6 847 . o 2.4 3.9 6 7 960 1 164 828 1 167 323 2586 8 956 333 5
Earth orbit insertion 100 7 5 6 . 4 8o7 . 2 2.6 4.1 67 108 713 136 715 672 4745 ll 34_1. 3318
Transposition and docking
Command & service modules 63 4 7 3 . 0 934 . 0 4. 0 6.5 3 4 445 76 781 79 530 -1789 -126 3148
Lunar module 33 2 9 4 . 5 1236 .2 0.2 0.1 2 2 299 24 826 24 966 -508 27 37
Total docked 96 767 . 5 1038.0 2.7 4.3 5 7 006 532 219 534 981 -7672 -9 240 3300
Separat ion maneuver 96 566 . 6 1038.1 2.7 4.3 5 6 902 531 918 534 766 -7670 -9 219 3270
First midcourse correction
Ignition 96 418 . 2 1038 . 3 2.7 4.2 5 6 770 531 482 534 354 -7Tll -9 170 3305
Cutoff 96 204.2 1038.4 2.7 4.2 5 6 667 531 148 534 ll3 -7709 -9 147 3274
LW1ar orbit insertion
Ignition 96 061 . 6 1038.6 2.7 4.2 5 6 564 530 636 533 613 -7785 -9 063 3310
Cutoff 72 037 . 6 1079 . 1 1.7 2.9 4 4 117 412 855 419 920 -5737 - 5 166 382
Circularization
Ignition 72 019 . 9 1079 . 2 1.8 2.9 4 4 102 412 733 419 798 -5745 - 5 160 366
Cutoff 70 905 . 9 1081 . 5 1.6 2.9 4 3 539 407 341 413 864 -5403 - 5 208 316
Separat ion 70 760 . 3 1082 . 4 1.8 2.8 4 4 762 407 599 414 172 -5040 -5 4o4 286
Docking
Command & service modules 36 847 . 4 943.6 2.8 5.5 2 0 747 57 181 63 687 -2094 833 321
As cent stage 5 7 38 . 0 1168 . 3 4.9 -2 . 4 3 369 2 34[ 2 8[3 -129 54 -354
Total after docking
Ascent stage manned 42 585 .4 973.9 3.1 4.5 2 4 189 ll3 70'i 120 67'i -1720 -1 018 -50
Ascent stage unmanned 42 563 . 0 972 . 6 2.9 4.5 2 4 081 110 884 117 8o4 -2163 -811 -28
After ascent stage jettison 37 100.5 943.9 2.9 5. 4 20 807 56 919 63 4l'i -2003 7 30 305
Transearth injection
Ignition 36 96 5 . 7 943.8 3.0 5.3 2 0 681 56 775 63 303 -1979 709 336
Cutoff 26 'i9 2 . 7 961 . 4 -0 . 1 6.8 1 5 495 49 843 51 454 -824 180 -232
Command & service module
separat ion
Before 26 65 6 . 5 961 . 6 o.o 6 .7 15 4o6 49 739 51 338 -854 228 -200
After
Service module 14 549 . l 896 . 1 0.1 7.2 9 143 14 540 16 616 -83'i 885 -153
Command module 12 107 . 4 1040 . 4 -0.2 6.0 6 260 5 470 4 995 55 -4o3 -47
Entry l2 095 . 5 1040 . 5 -0 . 2 5.9 6 253 5 463 4 994 55 -400 -4 7
Drogue deployment ll 603 . 7 1039 . 2 -0 . 2 5.9 6 066 5 133 4 690 56 - 37 5 -48
Main parachute ll 3 1 8 . 9 1039 . 1 -0.1 5.2 5 933 4 947 4 631 50 -312 -28
deployment
Lunar Module
Lunar module at launch 33 297 . 2 185.7 0.2 0.2 2 2 304 25 019 25 018 228 454 77
Separation 33 683 . 5 186 . 5 0.2 0.7 2 3 658 26 065 25 922 225 705 73
Descent orbit insertion
33 669.6
·-
Ignition 186 . 5 0.2 o.B 23 649 26 045 25 899 224 704 7l
Cutoff 33 401 . 6 186 . 5 0.2 0.8 2 3 480 25 978 25 871 224 704 7l
Lunar landing 16 1 5 3 . 2 213.5 0.4 1.6 l2 582 13 667 16 204 182 555 ·r 4
Lunar lift-off 10 776 .6 24 3 . 5 0.2 2.9 6 BoB 3 475 5 971 20 214 45
Orbit insertion 5 928.6 255 . 3 0.4 5.3 3 457 3 082 2 273 17 135 43
Coelliptic sequence initi - 5 881 . 5 255 . 0 0.4 5.3 3 437 3 069 2 246 l7 137 44
at ion
NASA-S - 6 9 -3 7 9 7
Remote control u n i t
S u n g lasses pocket
Oxygen purge
S u pport straps system actuator
Connector cover
I ntegral thermal
Uti l ity pocket
and meteoro i d
garmet
Pouch
U r i ne col lection and transfer
connector/ b i omed i cal injector/
dosimeter access flap and
donn ing lanyard pocket
NASA-S-6 9- 3 7 9 8
VHF antenna
Actuat i ng cable
·-(stowed)
-
�·"- (stowed position}
· · · ----
Heater status
Pressure gage
Regu lator
Oxygen
• purge system
•·.· 'actuator__
Stowage p late ·
NASA-S-69-3 8 0 0
West
N ickel thermometer
dust detector
East
Scoop
Tongs
Gnomon
H ammer
The history of the lunar module ( LM-5 ) at the manufacturer ' s facility ,
Bethpage , New York , i s shown in figure B-3 , and the operations at Kennedy
Space Center , Florida, in figure B-4 .
NASA-S-69 -3802 �
f\)
1968 1969
December January March
• Individual systems checkout, modification
and retest
• Integrated systems test
I I I Data review
I Demate
I Pressure vessel leak check
I Aft heatshield installation
Command modu le
I Weight and balance
• Preshipment inspection
I Prepare for shipment and ship
1 1 1 1 Service propu lsion system test
Service module
I Thermal coating
• Preshi pment inspection
I Prepare for shipment and ship
F i gure B-1 . - Factory checkout flow for command and service modules at contractor faci lity .
t ., I , '
I '
NASA-S-69-3803
1969
September
1968 1969
September II November J December
October January
J February I March
- -� F inal hardware instal lation and checkout
• Plugs-in test
I I I I Install and test radar
I I I P lugs-out test
• - · Final factory rework and test
• • Install thermal shielding
I I Weight and balance
I I Final inspection
- Install base heat shield
I I Prepare for shipment and sh i p
I I Landing gear functional test
F igure B-3 . - Factory checkout flow for lunar module at contractor facility .
. ... ' , '
NASA-S-69-3805
1968 1969
December January August
I Radar alignment
The command module arrive d at the Lunar Receiving Lab orat ory , Houston ,
Texas , on July 30 , 1969 , after reacti on control system deactivation and
pyrote chnic s afing i n Hawaii . After de contaminat i on and at the end of the
quarantine peri od , the command module was shipped to the contract or ' s fa
cility in Downey , Californ i a , on Augus t 14 . Postflight te sting and in
spection of the command module for evaluat i on of the inflight performance
and investigation of the flight irregularities were conducted at the con
tractor ' s and vendor ' s facilities and at the Manne d Space craft Center i n
accordance with approve d Apollo Space craft Hardware Utili zati on Requests
( ASHUR ' s ) . The tests performed as a result of in flight problems are de
s cribed in t ab le G-I and di s cus s ed in the appropriate systems performance
s ections of thi s report . Tests being conduct ed for other purpos es i n ac
cordance with other ASHUR ' s and the b asi c contract are not included.
TABLE C-I . - POSTFLIGHT TESTING SUMMARY
Environmental Control
107001 To determine the cause of the down- End-to-end resis tance and contin- A capacitor in the electromagnetic inter-
shift in oxygen flow reading and ui ty check of the flow rat e trans- ference filter was open and the res i s -
i t s remaining at the lower _limit ducer calibration; calibrat ion tance of the heater element on one o f
except for periods of high flow check and failure analysis the two air stream probes was 600 ohms
above the requirement.
107019 To determine the cause for the de- Leak t e s t o n the primary water/ System was found to be t ight and well
crease in the primary glycol ac- glycol system ; leak t e s t on the within specification . Indication was
cumulator quantity glycol reservoir valves that the glycol res ervoir inlet valve
was not fully closed during flight and
allowed leakage into the res ervo ir .
107503 To determine the caus e for high and Measure the glycol temperature con- All resi stances and deadband prope r .
low water/glycol temperatures troller deadband and determine re- Control valve bound closed.
sensed at ·the evaporat or outlet sponse to a s imulated glycol temper-
during mixing mode operat i on in at ure sensor
lunar orOit
107039 To determine the cause for high and Remove control valve from space- Broken bearing found interfering with
low water/glycol temperatures craft and perform electrical and gear t rain ass embly. Analysis incomplete.
sensed at the evaporator out let mechanical acceptanee tests. Dis-
during mixing mode operation in ass emble control valve .
lunar orbit
Reaction Control
107014 To determine the caus e of the mal- Circuit continuity verification Continuity test determined that an inter-
function of the command module mi ttent existed on a terminal board.
negative yaw thruster Wiring was found to be prope r .
107016 To verifY command module circuit Circuit continuity verification Control circuit for s ervice module reac-
associated with service module tion control quad B propellant isolation
propellant isolation valves for valves and indicators was proper through
g_uad B the command module to the circuit inter-
rupter interface .
,
' '
t '•If
. .
Crew Equipment
107028 To determine the cause of high clos- Examine the seal for comparison Vacuum seal satisfactory . Latching force
ing forces on the sample return con- with ground test. Re-roll seal and above maximum specification limits because
tainers measure latching forces . of lubrication removal . Appli c ation of
lubri cation on s imi lar latches � us i ng
Apollo 12 procedures , resulted in closing
forces below maximum specifi cation limits .
107030 To investigate the loose handle on Visual inspect ion . Determine The handle was not attached to right end;
the medical kit and overpressuri za whether pin holes will prevent on]y bare]y attached to left end. lin
tion of pill containers overpressuri zation vented pill packages expand about
300 percent at 5 psia from ambient.
Vented packages (two needle holes in
film) do not expand at 5 psia from am
bi ent .
107034 To investigate the voice turnaround Turnarmmd test with extravehicular No defect ive circuits or components in
problem durjng extravehicular ac co�Eunications system packs and either carrier . Up-voice turnaround was
tivity Commander and Lunar Module Pilot present in both headsets but always ac
headsets in all possible connectors . quired with the Lunar Module Pilot car
rier, regardless of position of connec
tion . Turnaround was cause d by audio/
mechanical coupling , and could be ac
quired or eliminated by control of mech
anical isolation of headset and earphone
output level .
107038 Investigate leak in riser of X-ray and visually inspect hose and During preflight adjustment o f the liquid
liquid cooling garment . manifold. Verify corrective action . cooling garment , the spring reinforced
riser hose was improperly drawn over the
manifold nipple , cutting the inner wall
of the hose between the spring and the
nippl e . Water/glycol leaked through the
inner wall hole and ruptured the outer
wall of the Lunar Module Pilot ' s garment
during postflight tests at the qualifica
tion level of 31 psig. No le akage was
found in the Commander ' s garment because
the inner wall was sealed against the
0
nipple by the spring behind the cut . I
Proper installation with the necessary \.)J
between the nipple and spring will pre
clude cuts in the inner wall .
..______.._______________.____________
D-1
Tables D-I and D-II are summaries of the dat a made availab le for
systems performance analyses and anomaly inve s tigat ion s . Tab le D-I li sts
the dat a from the command and s ervic e modul es , and table D-I I , the lunar
module . Although the t ables re flect only data processed from Network
magnet ic tapes , Network data tabulations and comput er words were avail
able during the mis sion with approximately a 4-hour delay . For additi onal
informat ion regarding dat a avai lability , the status listing of all mi s sion
data in the Central Metri c Dat a File , building 12 , MS C , should be consult
ed.
D- 2
5 6 : 50 58 : 10 CATa X X
5 7 : 15 5 7 : 30 GDS X
57 : 30 5 7 : 45 GDS X X
5 8 : 10 7 3 : 09 CATS X X
73:15 7 3 : 48 MAD X X X
7 3 : 48 75 : 48 MAD X
7 5 : 48 7 5 : 57 D/T X X X X
7 5 : 57 76:15 D/T X X
77 : 39 78 : 24 GDS X
78 : 24 79 : 09 GDS X X X
7 8 : 41 80 : 2 2 MSFII X X
79 : 07 79 : 47 GDS X X X
79 : 5 4 80 : 37 GDS X X
80 : 10 80 : 43 D/T X X X X X
80 : 22 85 : 41 MSFII X x
81 : 40 83 : 11 D/T X
83 : 43 84 : 30 D/T X
85 : 00 85 : 30 GDS X
85 : 41 86 : 32 D/T X
85 : 42 89 : 11 MSFII X
87 : 39 88 : 27 D/T X
88 : 32 89 : 41 HSK X
89 : 37 90 : 25 D/T X
90 : 2 5 93:07 MSFII X X
90 : 29 91 : 39 HSK X
91 : 36 92 : 29 D/T X X X
92 : 30 92: 40 HSK X
9 3 : 26 99:07 MSFN X X X
9 3 : 34 94 : 31 D/T X X
9 4 : 22 94 : 34 MAD X X
9 5 : 32 96:20 D/T X X
96 : 30 98 : 20 MSFN X
97 : 3 0 98 : 5 2 D/T X X
98 : 20 100 : 00 MSFII X
98 : 50 99 : 00 MAD X X
99 : 29 100 : 32 D/T X X
100 : 3 5 100 : 45 MAD X X X X X
100 : 44 101 : 19 MSFII X
100 : 5 5 102 : 45 MSFII X X
101 : 15 101 : 27 MAD X
101 : 27 102:14 D/T X X
102:15 102 : 48 MAD X
102 : 49 106 : 48 MSFI� X X X
103 : 25 104 : 19 D/T X
105 : 23 106:11 D/T X
106 : 28 110 : 21 MSFI� X X
107 : 21 108 : 10 D/T X
109 : 17 110 : 09 D/T X
110:31 113 : 16 MSFI� X X
111 : 1 8 112 : 3 8 D/T X
112 : 0 6 113 : 00 MSFI� X
113 : 11 117 : 02 MSFI� X X
113 : 18 114 : 04 D/T X
115 : 17 116 : 02 D/T X
117 : 13 118 : 01 D/T X
118 : 00 122 : 06 MSFN X X
D- 4
APPENDIX E - GLOSSARY
euhedral having crystals whose growth has not been interfered with
gnomon im1trument use d for size and color comparison with known
standards
induration hardening
E-2
lithic stone-like
ray any of the bright , whitish lines seen on the moon and
appearing to radiate from lunar craters
terra earth
.8 . Cali forni a Insti tute o f Te chnol ogy , Jet Propuls i on Laborat ory :
1968 Surveyor Proj e ct Final Report Pt II B ci ence Results , Sect i on I I I
Televis ion Obs ervat i ons from Surveyor .
10 . NASA Headquarters : Prot e ction of the Earth ' s Biosph ere from Lunar
Sources of Contaminat i on ; An Interagency Agreement Between the
National Aeronauti cs and Space Admi ni strat i on ; the Department of
Agriculture ; the Department of Health , Educat i on , and Welfare ; the
Department .of the Interior ; and the Nati onal Academ[ of S ci ences .
August 2 4 , 196 7 .
Apollo 7 CSM 101 First manned flight ; Oct. ll ' 1968 Cape Kennedy ,
/
earth-orbital Fla .
Apollo 8 CSM 103 First manned lunar Dec . 21 , 1968 Kennedy Space
orbital flight ; first
manned Saturn V launch
Apollo 9 CSM 104 First manned lunar Msr. 3 , 1969 Kennedy Space
I.M-3 module flight ; earth Center, Fla.
orb it rendezvous ; EVA
Apollo 10 CSM 106 First lunar orbit Me¥ 18 , 1969 Kennedy Space
LM-4 rendezvous ; low pass Center, Fla .
; ,'
k
over lunar surface
Apollo 11 CSM 107 First lunar landing Ju:cy 16 , 1969 Kennedy Space
LM-5 Center , Fla.
Apollo 12 CSM 108 Second lunar landing Nov. 14 , 1969 Kelllledy Space
I.M-6 Center, Fla.