apollo experience report window contamination

8/8/2019 Apollo Experience Report Window Contamination

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N A S A T E C H N I C A L N O T E - ASA 0-6721( 2 . 1

=-I

LOAN COPY:A F W L (

KIRTLAND

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AFB, N,DOUL)

APOLLOEXPERIENCE REPORT -W I N D O W CONTAMINATION

by L. J . Leger d n d R . W . Bricker

Mdnned Spucecrczft CenterH o ~ s t o n ,Te x m 77058

N AT I O N A LA E R O N A U T I C SA N DS PA C EA D M I N I S T R AT I O N WA S H I N G TO N ,D. C.-.. . . , .MkRCH'1972


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TECH LIBRARYKAFB, NM

1. Report No. 2. Government Accession No.

NASA TN D-67213. Recipient's Catalog No.

- _ ~ _ ~~~ ~~

4. Title and Subtitle I 5. Report Date

APOLL.0 EXPERIENCE REPORTWINDOW CONTAMINATION F -. Performing Organization Code

7. Author(s)~~~ . ~~

8. Performing Organization Report No.

L. J. Leger and R. W. Bricker, MSC10. Work Unit No..

~

MSC 5-284

9. Performing Organization Name and Address 914-13-20-21-72Manned Spacecraft CenterHouston, Texas 77058

11. Contract or Grant No.

~~

". . 13. Type o f Report and Period Covared

2. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D. C. 20546

I Technical NoteI14. Sponsoring Agency CodeI~~~~ ~_ _ _ _ _

5. Suoolementarv Notes

The M SC Director waived the use of the International System of Units (SI) forthis Apollo Experience Report, because, in hi s judgment, use of SI Units would impair the usefulnessof the report or result in excessive cost.

6 . Abstract

The prob lem of window contamination in the Apollo command module is reviewed. All fivecommand module windows were con taminat ed while in ea rt h or bi t on the first three mannedApollo flights. The conta minatio n, sources were identif ied and elimina ted. Prefli ght testingof lu nar module windows showed that no se ri ou s contamination should occ ur, and this conclu-sion was verified in subsequent manned flights. In this report , the command module windowdesigns and materials are described, the window-contamination so ur ce s ar e identified, andthe inflight and chemic al analyse s of the contamination are outlined. The corrective actionsthat were taken are reviewed. For the lunar module, the window desi gn and mate rial s andthe preflight and flight evaluations are desc rib ed. Window-design recommendations a r e made.

7. Key Words Suggested by Author(s1

' Spacecraft Windows' Window Contamination' Contamination

18. Distribution Statement

I

19. Security Classif. (of this report)~~ ~

20. Security Classif. (of this page) 22. Rice'1. NO. of Pages

None $3 007one"~.

For sale by the National Technical Information ervice, Springfield, Virginia 22151


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APOLLO EXPERIENCE REPORT

W I N D O W C O N TA M I N AT IO N

By L. J. Leger and R. W . Br ickerM a n n e d S p a c e c r a ft C e n t e r

S U M M A RY

All five command module windows on each of the first three manned Apollo flightswere contaminated while the spacecraft was in earth orbit. This contamination accu-mulated most of ten on the inne r surf ace f the outer pane and on the oute r s urfa ce ofthe middle pane of the three-pane window. By mea ns of astronaut debr iefing and post-flight examination of the recovered command module, three main sources of contamina-tion were identified: escape-tower-engine plumes, expulsion of waste water, andoutgassing of elastomeric sealants.

Because contamination on the middle pane w a s a serious problem, a concentratedeffort was made to eliminate the source of thi s contamination. The contamination w a sidentified, by mean s of chemica l ana lys is, as a silicone oil that resulted from outgas-sing of ela stomeri c window mat eri als . By mean s of weight -loss experiments on thesematerials? it was demonstrated that a solution to the problem would be a 48-hour vacu-um bake of all appropriate sealants at 200" C before installation. Use of this techniquereduced window contamination to an acceptable level.

As a re su lt of the problems encountered with the Apollo command module win-dows, tests were conducted on a lunar modu le forward window before manned flight ofa lunar module. The re su lt s of the se tests indicated that no serious contaminationwould be encountered on the lunar module window. This conclusion w a s verified insubsequent manned flights.

INTRODUCTION

The first window used in the U. S. manned space flight program was on the Mer-cury spacecraft. In thi s case , only one window (consisting of four panes) w a s provided.Because of its si ze, th e window provided only limited visibility, and it received onlylimited exposure to the space environment because of th e r elat ivel y s hor t flight dura -tions. Even for these conditions? window Contamination was noticed on some of theflights (refs. 1 and 2).

The Mercury window design influenced the Gemini and Apollo window designs.Major changes such as window size and number of a ss emb lie s were made; however,


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the basic concept of a multipane assembly, wherein elastomeric sealants were used fora p r e s s u r e seal, was retained . At the same time, other portion s of the spacecraftunderwent major redesign, but the resultant effects of these changes on the windowass embl ies we re not considered. It is now known that these changes caused majorproblems in the maintenance of satisfactory windows in the Gemini and Apollo P ro gra ms .Window requirement s became more string ent as space flights increased in complexity.Rather than simply providing a po rt fo r viewing and g ener al photography, the windowsbec ame cri tic al i tem s in docking maneuvers, in scientific photography, and, eventually,in the lunar landings.

The first serious window-contamination problems were encountered in the earlyGemini flights. It was determined by means of analy sis that the majority of this con-tamination was a silicone oil that appeared on the out ermo st window su rfa ce as a highlyviscous fluid. This contamination was produced by outgassing of the Gemini nose-capheat shield. During launch, ascent heating of the spacecraft had caused the siliconeoils in the ablator to vapori ze and, subsequently, to redeposit on the cooler windowsurfaces. Also, very light deposits were located on the interior surfaces of the threepanel windows, and the so ur ce of these deposits was identified as outgassing of window-sealant materials.

Because the contamination on the Gemini spacecr aft windows was a cle ar oilyfluid (during flight), it did not in terf ere seve rely with the functions of the spacecraft;therefore, immediate changes to the spacecr aft were not considered necess ary. Sev-e r a l window modifications were made for the later flights of the Gemini Program,however. One modification wa s a window cov er for us e duri ng the ascent phase offlight; another modification was a vacuum bake for the window sealants in an effort toreduce the amount of outgassing (ref. 3). These changes reduced the contaminationconsiderably; however, during the last Gemini flight, contamination could be detectedon the windows.

Unfortunately, the analysis of the Gemini window-contamination re su lt s was notconcluded early enough to be used in the Apollo Program, and contamination was aproblem again on the command module (CM) windows. In an effo rt to provide cleanwindows for docking and other operation s, the cause of this contamination was examinedin detail. The physic al character istics of the contamination during flight were obtainedthrough crew debriefings, and, in most cases, chemical analyses of the contaminationcould be made from samples collected postflight. From these data and fro m the out-gassing character istics of the materials used in the window assemblies, it was possibleto identify the ma jo r so ur ce of contamination: condensation of high-molecular-weightsubsta nces outgassed by seala nt compounds. The re su lt s of the investigation that ledto the identification of the contamination so ur ce as well as the approach that was takento reduce the contamination a re presented in this repo rt. A similar program to evaluatethe lunar module (LM) windows is reported also.

C O M M A N D M O D U L E W I N D O W S

Window D es ign and Mate r i a ls

Five multipane window asse mbl ies are used on the Apollo CM. The standard win-dow designa tion used f or t he CM is shown in figu re 1. Each assembly has three panes;

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the two nner panes are par t of the pressure-vessel structure, and th e outer pane ispar t of the heat-shield structure. Cross-sectional views of these windows are shown infigures 2 to 4. The outer pane protects the inner panes during various phases of themission, ncluding entry of the spacecr aft nto the atmosphere. The rendezvous win-dows (numbered 2 and 4 in fig. l ) , as the name implies, were designed to provide aview of the LM during docking maneuvers. Most of the visual observations and photo-

graphs were made using the other three windows becaus e the surface area of thesewindows is larger.

Rendezvous wmdowSide window

Rendezvous windowHatch wlndow - 2

+ X

AfHatch window

Rendezvous wmdowRendezvous window

Side wmdowSide window

Figure 1 . - Apollo CM window designation.

q K b l a t o r o u t e r silica pane

,-Vacuum du ring orbit

\I nn er aluminosilicate panes

Figure 3 . - Detail of CM rendezvouswindows 2 and 4.

The panes of the window assembliesare constructed of two types of glass. Threedifferent coatings are used in each assemblyto improve the optical properties. Thepressure-vessel panes a re made of temper edaluminosilicate glass, and each surface(surfaces 3 to 6 in fig. 5) of the double-paneconstruction is coated with a proprietary,

Outer silica pane

Vacuum during orbit

ner aluminosilicate panes

Figure 2. - Detai l of CM sidewindows 1 and 5.

rAblator

aluminosilicateDanes L D r y Nz, 389 torr

duringorbit

Figure 4. - Detail of CM hatch window 3 .

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antireflectiongent.hisoatingncreases @ M ~ F ~oatingthe total light transmission through the as-sembly by reducing the surface reflectionscreated by each pane of the multipane con-struction. The heat-shield window panesare made of 100 percent fused silica (Cor-ning glass 7940) and a re coated on the ex- - 4terior surface (surface 1 in fig. 5) withmagnesium fluoride (MgF2) and on the in-terior surface (surface 2 in fig. 5) with ablue-red coating that is used as a spect ra lfilter.

Insulation

Because of the high temperaturesthat a re encountered during entry, insula- Figure 5. - Command module window-tion is neces sary between he heat shield surface notation used n contaminationand the pressure-vessel tructure. hisocations.

insulation, as shown in figure 5, extendsto the edge of the aluminum window f ra me and consi sts of four 1/2-inch l aye r s ofsilicone-resin-impregnated glass fibers bonded with sili cone rubber RTV-511 (GeneralElectr ic). The exter ior of this composite, silica fibers and silicone rubber, is coatedwith an add itiona l layer of RTV-511. Any area of the insulation that is exposed toastro naut view is sealed in a layer of aluminum foil reinforced with glass cloth and isovercoated with black silicone rubber 92-018 (Dow Corning) for an tiglar e pur pose s.

The gasket that surrounds the heat-shield window is made of a glass-cloth-reinforced, heat-molded silicone rubber.The center is filled with a silica-fiber in-

sulating material. This gasket is appliedto the glass in two sections, and the s e c -tion joints a re bonded with silicone rubberRTV-560 (General Electric). Also, sili-cone rubber RTV-560 is used as a seal*on the press ure-v essel window assemblyand is applied in the areas shown in fig-u r e 6 by the use of da ms of heat-moldedsil icone rub ber . Although slight geomet-ric differences exist for the materials usedaround the five windows, the conside ra-tions mentioned previously are applicablegenerally.

RN-560Window

. .

-Sealantdams

Figure 6. - Detail of pressure-vessel-panel seal.

After the pressure-vessel panes are mounted and sealed to the frame, the areabetween the panes is evacuated and repressurized to 389 torr with dry nitrogen (N2).

Then, the assem bly is sealed and rem ains se aled throughout its use. The space be-tween the press ure-v essel windows and the heat-shield windows is vented to ambient

pressure. During flight, he pr es su re in this space is approximately 1 X 10 t o r r.4

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I nflight Analysis of Window Contamination

Contamination of t he spa cec raf t windows was obser ved as a problem on the earlyunmanned Apollo flights. The window conditions of the se fli gh ts were evaluated by useof the results of inflight photography and postflight analyses. These evaluation tech-niques were limited because, by the use of photography alone, the window cannot be

evaluated from various angles, and postflight analyses must e made on surfaces thatwere contaminated during entry. A complete objective evaluation of the contaminationproblem was obtained when the: Apollo 7 and 8 astronaut reports became available. Ingeneral, these crew reports indicated that the hatch and ide windows were unusa blefor most pur pos es bec aus e of contamination. In comparison, the rendezvous windowshad much less contamination and met the mission requirements.

Se vera l st ages of contamination could be identified from the flight reports. Thischronological information, in conjunction with a descri ption of th e ap pearan ce of thecontamination, helped considerably in the identification of the contamination sources.A descr ipt ion of each identifiable stage of contamination is given in the followingparagraphs.

Becau se the spacec raft i n the prelaunch phase of the mission is exposed to atmos-pheric weather conditions, rain has been a prob lem on some of the flights. The watercan provide collection sites for other contamination o r can collect in certain areas and,by evaporization and recondensation, deposit on the windows during flight. This sourceof contamination was a major problem on the Apollo 12 mission.

During jettison of the launch escap e tower and boost protective cover, a lightgray film was deposited on sur face 1 of the Apollo 7 windows. Although th is conditionw a s not repo rte d on any of the other fligh ts, it may have contributed some windowcontamination on all flights.

Soon after earth -orbit al inser tion, afilm appeared on surfaces 2 and 3 of thehatch and side windows of the Apollo 7and 8 command modules. This contami-nation continued to accumulate duringflight and resulted in a heavy film in somelocations and droplet formation in otherlocations. A photograph, aken with acamera focused on the hatch window duringthe Apollo 7 mission, is shown in figure 7.During the night p as s of the spacec raft ineart h orb it and during other similar cold-

window conditions, the contaminationseemed to crystallize and become opaque,particularly on the hatch window.

Another so ur ce of contamination hasbeen the waste-water-system expulsions.Dropl ets were depos ited on the windowsduring al l Apollo flights when this systemwas activated. Figure 7. - Apollo 7 inflight hatch-window

contamination.

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In general, however, the most serious contamination was on su rf ace s 2 and 3 ofthe hatch and side windows on the Apollo 7 and 8 missions. Crewmembers from thesetwo missions felt that, i f the contaminants on these surfaces could be reduced or elim i-nated, the minor amount of contamination on the out er sur face s was not sufficientlyseri ous to inte rfe re with visual observations and photography.

C h e m i c a l A n a l y s i sof Contamina t ion

Identification of the sources of major contamination on surfaces 2 and 3 of thehatch and side windows and confirmation of escape-tower-engine exhaust products asa mino r s ourc e of contamination on surfa ce 1 of all windows were obtained by meansof chem ical anal ysi s of the contaminants. Surface 1 of the Apollo 7 and 8 windows wassampled for optical-emission spectrographic analysis. Samples were collected fromsurfaces 2 and 3 of the hatch and side windows for infrared analysis.

The results of optical-emission spectrographic analysis of the contamination onsurface 1 of the spa cecr aft windows

areshown in table

I. Theelements

that were

TABLE I. - OPTICAL-EMISSION SPECTROGRAPHIC RESULTS

Element

Aluminum

CalciumCopper

Iron

Magnesium

Sodium

Nickel

Lead

Silicon

Silver

Tin

1

aM3

M3bS

M2

M3

C

d M l

S

S

M2

S

S

2

M3

M3S

M3

M3

M1

S

S

M3

S

S

Window number

3

M 3

M 3

S

M3

S

M1

S

S

M 3

S

S

4

M3

M 3

M2

M 3

M1

S

S

S

M1

S

S

15

M 3

M 3S

M3

S

M1

S

S

S

S

a

b~ = 0.1 to 1 percent, minor constituent.

M = 5 to 10 percent, second major constituent.

dM = 10 to 15 percent, first major constituent.

M = 1 to 5 percent, third major constituent.3

2

1

6


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present in large concentrations (such as calcium, magnesium, and sodium) a re commonto seawate r, o which the window w a s exposed during recovery operations. Elements(such as aluminum, iron, and lead) that were classed as the second and third majorconstituents of the contamination are not common to any so ur ce of Contamination thatwould resul t fro m reco ver y ope rati ons and spe cial handling of the spacecr aft. How-ever, these elements are common to the escape-tower-engine fuel. Therefore, thisengine probably is a so ur ce of contamination.

Because of high en try tem per at ure s and recovery conditions, no residue fro m thewaste-water-dump system w a s expected to remain on the heat-shield windows. Noneof the elem ents found on the window by me ans of s pec trog raphic ana lys is c an be t raceddirectly to the waste-water-dump system.

The infrare d spec tra of contamination collected from surfaces 2 and 3 on thewindows of the Apollo 7 and 8 spacecraft were typical of organosilicon high-molecular-weight compounds. A represen tat ive spect rum of this contamination is shown in fig -u r e 8; the spectrum contains absorption bands common o the methyl-phenyl siliconest ru ctu re . The same type of contamination was identif ied on the window of an ApolloCM (2TV-1) that w a s used for thermal-vacuum testing at the NASA Manned SpacecraftCenter.

I I 1 I I 1 I I I I I I I I2 3 4 5 6 1 8 9 10 11 12 13 14 15 16

Wavelength. pL I I 1 1 - I I I L L I I I I J

Moo Mo o m 1400 1m 1oW 800 650

Figure 8. - Typical infrared spe ctru m of cont aminatio ns from surface s 2 and 3on the Apollo 7 and 8 CM's.

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To identi fy the s pecific sourc es of contamination on surfaces 2 and 3, all nonme-tallic materia ls (16 types) that were used in the window areas were tested for outgas-

sing at 66" C in a vacuum (1 X to r r ) ( ref . 4). Outgassing tests we re conducted inall-glass sample chambers that were connected to individual cold traps. The outgass ingcondensate was collected in the upper portions of the sample chambers (at 25" C) andin the cold traps (at 0" C). At. the end of the exposu re time (9 days), each outgassingcondensate was rated according to relative volume y use of an ar bi tr ar y but commonscale. The results of these screening tests and comparisons of the inf rar ed spe ctr a ofthe collected condensate and contamination fr om the Apollo 7 spacecraft indicated thatthe ma jor s our ces of contam inatio n were silicone rubber s 92-018, RTV-560, andRTV-511.

Isothermal-vacuum eightoss of 1the problem materials was determined byuse of a continuous-recording vacuum bal-ance. The vacuum weight lo s s of thesematerials at 9 3 " C is shown in fig ure 9 .This temperature w a s the maximum tem-perature xpected in flight or materials Ethat were used in the window area. Sam -ples of all three materials had arge ni- E ;tial weight-loss rates (approximately1.25 ercent/hr) hat ubsequently, fter P15 hours, decreased to a much smallerrateapproximately .01ercenthr). 1

-0 )a

--OI.-

Correc t ive Ac t ion

Because the weight-loss rates forthe problem materials were smal l after

0

Temwrature: 25 " C and 93" C

Temperature: 93' C

-Temperature 25" CI I I

Tim e . hra

16 24

__ I

32

-

exposure to vacuum at tempera- Figure 9 , - Vacuum weight lo ss of windowtur es, an apparent solution to the silicone-outgassing problem was to bake themanufactured Darts in a vacuum before

materials that caused contamination onsurfaces 2 and 3 .

installation nthe spacecraft. This proc-e s s w a s used with success in r educing theGemini window contamination (ref. 3). In applying a vacuum bake, it is desirable tous e as high a temperature as possible in the thermal-vacuum exposure. However, athigh temperatures (approximately 2 5 0 " C), polymer degradation is possible. There-

fore, a temperature of 2 0 5 " C was chosen as the maximum allowed in the vacuum bake.At this temperature, the weight loss levels off (fig. 10). This condit ion indicates nomajor thermally induced polymer degradation has occurred.

Furth er cons iderat ions of the corrective action indicated that the vacuum bakeshould be conducted at low pre ss ur es a nd f or long periods of time. However, because

of limited chamber availability, a maximum pressure of 1 x 10-1 torr was allowed, and

an exposure time of 1.72 X 10 seconds (48 hours), divided into two 8.64 X 10 -second(2 4 hours) steps, was used. The weight loss during and after exposure of the materials

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Temperature: 25C and 205.

i 2l- emperature 205" C

I I0 6 128 24

Time, hr

Figure 10. - Vacuum weight lo s s of RTV-560a t temper atures used in the manufac-

turing rocess at 1 x t o r r.

6r Temperature: 205' C

LPrebake pressure

1 IO" torr

r h s t b a k e pressure

1 x torr

I

0 15 30 45Time, hr

Figure 11. - Vacuum weight oss of RTV-560for manufacturing cycle simulation.

to a vacuum of 1 x 10-1 o r r at 205" C isshown in figure 11. Because the totalweight loss af ter the vacuum bake is smal l(approximately 0.08 percent) and becauseonly a small amount (approximately

0.008 percent) of this weight los s is con-densable material at 2 5 " Cy a vacuum bake

at the conditions selected (1 X 10-1 t o r r fo r

two 8.64 x 10 -second periods) was con-sidered adequate to prevent contaminationof the spa cec raf t windows.

4

To verify that vacuum baking was anadequate solution for full-scale windows,an Apollo hatch was assembled from bakedcomponents and was tested in a large vac-

uum chamber. Heating was provided by asolar lamp and supplemental infrared heat-ers . The hatch was tested under simulatedearth-orbital thermal conditions and underconditions of maximum heating that may beencountered in a maneuver such as an at-titude hold for docking purposes.

During the maximum heating condi-tions of the test, the maximum tempera-tu re that was measured for the nonmetallicmaterials was 87" C , whereas the tem-per atu re of surfa ce 2 of the hatch windowattained -20" C in the cold conditions ofthe test. A s mal l amount of contamina-tion wa s noticed during a sh o rt portion ofth e first cold-soak period of the hatchtest. The contamination disappearedear ly in the first heating cycle and didnot reappear. This phenomenon hadbeen noticed in ea rl ie r tests and wasattrib uted to wa ter va por t rappe d betweenthe heat shield and the pressure-vesselstructure. Post-test nfrared analysisof the hatch-window surfaces indicatedno detectable contamination.

The Apollo 9 mission was the first flight on which vacuum-baked parts wereused. The rendezvous windows had no contamination problems on earlier flights;therefore, it was decided to use baked parts only on the hatch and side windows.

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Crewmember reports and the postflight examination did not indicate ny contaminationproblem s on the windows made of baked pa rts . The rendezvous windows, however,were contaminated slightly. This problem did not cause major concern because sub-sequent flig hts would involve the use of baked parts for all Uiindows. A photograph(fig. 12) of the hatch window of 2TV-1 (a test vehicle that revealed window contamina-tion similar to that found on the Apollo 6, 7, and 8 flights) is compared with a gridphotograph (fig. 13) of the Apollo 9 hatch window after recovery. In figure 1 3 , the grid,which is behind the window ass emb ly, is shown clear ly. Conversely, large amountsof contamina tion over the entir e window are shown in figure 1 2 . In this photograph,condensation of the contamination had accumulated to such an extent that droplet streakswere detectable. Although the photographs were not taken under the same conditions(the 2TV-1 photograph had no grid), they both indic ate that a significant improvementwas obtained by the use of vacuum-baked parts.

The Apollo 1 0 mis sion was the fi rs t flight for which vacuum-baked par ts we reused for all window appl icat ions . The astr onauts repo rted no window contaminationexcept a narr ow stre ak of droplets on surface 1 of side window 5. This streak wasnoticed after tower jettison, and the streak remained throughout the flight. Postflightobservations indicated that this contamination probably was caused by ex traneous waterunder the boost protect ive cover. The Apollo 11 sp ace cr aft had vacuum-baked pa rt sfor al l window applications; no window contamination was encountered that was causedby elastomer outgassing.

When the major contamination prob-lem had been solved, other sources ofcontamination found on the Apollo 9 and 10missions were reevaluated on subsequentmissions. These additional sources

Figure 12. - Hatch window of 2TV-1.

1 0

Figure 1 3 . - Apollo 9 hatch window.


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included tower -jettison-engine plumes, waste-water expulsions, reaction-control-engine plumes, and rain. Contamination caused by the tower-jettison-engine plumesand the reaction-control -engine plumes was not noticed by either crew. Contaminationcaused by the expulsion of waste water w a s evident on all flights. Because this con-tamination was minor and was almost completely removed by sho rt-duration exposu reto the sun, corrective action was not considered necessary. Rainwater, wh ich was amajor problem on the Apollo 12 mission, was noticed before flight on the exterior sur-faces of the heat shield window. Apparently, this water collected contamination fromthe escape-tower engine-firing sequence. After the water vaporized, th e contaminationremained on the window surface throughout the flight. This residue affected windows 1to 4 . Alterations in the boost protective cover were made in an attempt to preventwater leaks on subsequent vehicles.

LUNAR MODULE WINDOWS

Window Des ign and Mate r i a l s

Each of th e th re e windows used onthe LM has tw o panes: one pane formspart of th e pressure-vessel structure, andthe other pane provides micrometeoroidprotection for the pressure -vessel panes.The locations of the two triangular win-dows that a re designated as the forward orlanding windows and a smaller dockingwindow a r e shown in figure 14. The for-ward windows a re in front of the two crewsta tions, and the docking window is directlyabove the left-hand crew station.

Dockingwindow

Forward-windows

Figure 14. - Locations of LM windows.

The materials used in both types of window as se mb li es ar e sim ila r. The outerpanes of all windows are made of 96 percent fused silica (Corning glass 7900) and theinner panes of chemically strengthened glass (Corning glass 0311). The outer surfacesof all the outer windows (surface 1 in figs. 15 and 16) are coated with a blue-red coat-ing that is used as a spectral filter. To prev en t fogging of the windows when the LM ispressurized, the manufacturer provided an electrically conductive coating on the outersur fac e of the inner pane (surface 3 in figs. 15 and 16). This conductive coating per -mi ts the ast ron aut s to apply resis tive heating to the windows. The inner surface of theinner and outer panes is coated with a high-efficiency antireflection agent to improvethe overall optical properties of the window assemb lies. Als o, various mark ings onthe window surfaces are used to aid in navigation.

All three outer panes are sealed into the vehicl e struc ture by the use of a heat-molded rubber (Dow Corning Silastic 675-24-480), as shown in fig ure s 15 and 16. Thethermal-insulation blanket that protec ts the window fr ame is held in place by an epoxy/fiber-glass retainer ring. The forward-window pressure seal is made by pressi ng,with a Teflon and Silastic-jacketed metal spring, the inner pane against a silicone-rubber seat. The docking-window pressure seal is made by adhesively bonding theinner pane directly to a Kovar metal frame. Because the LM does not have to survive

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f -

- utboard

Figure 15 . - Cro ss sec tio n of LM forwa rdwindows.

Surface 1

B- Pressure-vessel wallOutboard

Figure 16. - Cr oss sec tio n of LM dockingwindow.

an atmospheric entry, the amount of insulation material that is needed around the win-dows is much less for the LM than for the CM.

Preflight Evaluation

As soon as it was recognized that a serious contamination problem existed on theCM, the LM window assem blies were examin ed fo r a similar problem. Because themaj or sou rce of contamination on the CM was outgassing of nonmetalli c ma terials ,weight loss was determined for all the nonmetallic materials in the LM windows. Theseweight-loss data a re shown in table 11. In the past, it has been determined that a totalweight loss of more than a fe w tenths of 1 percent can cause contamination in certainconfigurations. Because some of the ma te ri al s tested for use in the LM windows havetotal weight los se s that fall in this class, it was decided that a full-scale configurationtest should be made to determine i f a contamination problem existed. It should benoted that, in general, the weight loss of the LM-window materials was much less thanfor the CM-window materials.

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TABLE II. - VACUUM WEIGHT LOSS OF LM WINDOW MATERIALS

~

Material

Silastic 675 -24-480

Black velvet paint

Metlbond 329

Epon 828

RTV 90

Teflon TFE

Aluminized H-film

-~Total vacuum weight loss, percent

0.08

1.4

.05

.65

.6

.04

- 1 4

The full-scale configuratio n window test con siste d of a 2.16 X 10 -second(60 hours) thermal-vacuum exposure of a complete forward-window assembly. Through-

out this exposure, the pres sure was 1 x l o m 6 o r r o r less. The thermal environmentconsisted of three continuous phases that were designed to simulate the most severeheating portions of flight. The first phase was used to simulate the unmanned trans-

lunar phase by me ans of cycling between 5" and 32" C for 7 . 2 x 10 seconds (20 hours);then, one cycle to 93" an d -29" C was executed. The second phase consisted of a

7.2 X 10 -second simulation of th e lunar-orbital phase and included cycling between93" and 149 C, ollowed by o ne cycle from 149" to -29" C. The final phase was used

to simulate a hot lunar stay, and the cycle included exposure for 1.8 X 10 seconds(5 hours) a t 154" C, a drop to -29" C, a short-term cycle back to 154" C, and a recycleto -29 C. All the temperatures for the exposures were measured at the pressu re -pane outer surface.

5

4

4

4

In addition to pressure and temperature measurements, photographs of a resolu-tion grid were taken through the window assembly at intervals during the es t to eval-uate optical-property changes. A more detailed optical-property evaluation was madeby light-transmission measurements before and after testing. Provisions were made

to collect and analyze samples of any surface residue that remained after testing.

The results of the thermal-vacuum exposures indicated that no contaminationproblem existed within the temperature and pressure limits of the space envi,ronment.Vi s u a l inspection after tests and photographic coverage during te s ts did not indicatecontamination; therefore, no samples were collected fo r analysis. Post-test trans-mission measurements were representative of a loss of 3 to 4 percent, which possiblycould have been caused by dust collection in the handling process. This loss is verysmall and would not affect visual sightings o r photography.

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I 111 I I I

Inflight Evaluation

I I I .I11 I I

An evaluation of the LM window performance was first obtained on the Apollo 9mission. On this an d al l subsequent LM flights, the astronauts detected no unexpectedcontamination on the windows. A s expected, upon pressurization of the LM, water

condenses and collects on the inne r surface s of the inner panes. This condensate isremoved by activation of the window hea ter s. In general , it can be concluded that theLM windows performed satisfactori ly.

The relatively contamination-free condition of these windows might seem sur-prising because it was indicated earlier that the weight loss of the materials that wereused in construction was rather large in some instances. However, if table It andfigures 15 and 16 ar e examined, it is c lear that the materials that sustain large weightlosses are used in very limited amounts and, in general, a re not exposed to the ar eabetween the panes. There is one exception: the f iber-glass retainer that involves anepoxy as the matrix material. The weight loss for this material was 0.65 percent.Because no contamination problem w a s encountered, the amount of condensab le out-

gassing products from the epoxy was assumed to be very low.

CONCLUDING R E M A R K S

All window assemblies in the Apollo spacecraft were examined for contaminationproblems. Primarily, reduction of the opti cal p rope rtie s of the command module win-dows was caused by the contamination from outgassing of silicone sealants that wereused in the vicinity of the windows. This contamination was reduced to an acceptableleve l, with re sp ec t to mission objectiv es, by mean s of vacuum baking the siliconesealants.

The lunar module windows als o were examined fo r the contamination problems.A test program on the individual materials that were used in the window-assemblyconstruction and a full-scale configuration test indicated that no significant contamina-tion problem s exis ted on the lunar module window assem blie s. The re su lt s of inflightevaluations substantiated this conclusion.

As far as contamination is concerned, th e best spacecraft-window design shouldinvolve the use of a direct glass-to-metal seal. Currently, such a design is not prac -tical fo r windows la rg er than approximately 4 inches in diameter; nonmetallic eals mustbe used. Even for an application such as in the Apollo Pro gr am, the use of nonmetallicmaterials around optical surfaces must be controlled very closely. For these applica-tions, the first choice for a sealant m aterial should be heat-molded elastomers thatcan be cured at elevated temperatures and low pres sur es. If a room-temperature-vulcanizing agent is requ ired , new pro ducts have become available recently. Thesenew products have superior outgassing properties, and weight-loss values (in a thermal-vacuum environment) for these products are quite low and a re comparable to the weightloss of the vacuum-baked pa r t s that are used on the Apollo spacecraft.

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Window seals and sealants are indeed important when contamination prob lems a reconsidered. However, as was evident on the Apollo command module, the overa ll win-dow design must be considered also. Materials that are used in the main or supportingstructure, as well a s in other applications near the window assembly, must be testedto ensure low outgassing rates. Such an approa ch w a s import ant during the ApolloProgram and will be even more important in future programs.

Manned Spacecraft CenterNational Aeronautics and Space Administration

Houston, Texas, September 17, 1971914-13-20-21-72

REFERENCES

1. Anon. : Mercur y Project Summar y Including Result s of the Fourth Manned Orb italFlight May 15 and 16, 1963. NASA SP-45, 1963.

2 . Anon. : Result s of the Thi rd United States Manned Orbital Space Flight , October 3,1962. NASA SP-12, 1962.

3. Blome, James C. ; nd Upton, Bruce E. : Gemini Window Contamination Due toOutgassing of Silicones. Proceed ings of the 11th National Symposium and Ex-hibit of the Society of Aerospace Ma terial and Pro cess Engin eers, Scien ce ofAdvanced Material s and Proces s Engineering Series, vol. 11, 1967, pp. 217-225.

4. Kimball, L. G. : Postfl ight Invest igation of Apollo Command Module 101 Windows.Rept. NA-68-922, Lo s Angeles Division, North American Rockwell, Dec. 1968.

NASA-Langley, 1912 - 1 S 2 84 15

apollo experience report window contamination

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