Apollo Experience Report Spacecraft Structural Windows

8/8/2019 Apollo Experience Report Spacecraft Structural Windows

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N A S A T E C H N IC A L N O T E

APOLLO EXPERIENCE REPORT -SPACECRAFT STRUCTURAL WINDOWSby Oruis E . Pigg and Stanley P . WeissLyndon B , Johnson Spuce CenterHouston, Texas 77058

NAT ION AL AERONAUTICS AND SPACE ADMINISTRATION WASHING TON, D. C. 0 SEPTEMBER 1973


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1. Report No. 2. Government Accession No.NASA TN D-74394. Title and Subtit le

APOLLOEXPERIENCEREPORTSPACECRAFT STRUCTURAL WINDOWS

3. Recipient's Catalog No.

5. Report DateSeptember 1973

7. Author(s) 8. Performing Organization Report No.JSC 5-377Orvis E. Pigg and Stanley P. Weiss, JSC 10 . Work Unit No.

Lyndon B. Johnson Space CenterHouston, Texas 770582. Sponso ring Agency Name and Address

Washington, D. C . 20546National Aeronautics and Space Administration

I 914- 13-20-13-72. Performing Organization Name and Address11. Contract or Grant No.

13. Type of Report and Period CoveredTechnical Note

14 . Sponsoring Agency Code

17. Key Words (Suggested by Author(s))'Apollo Spacecraft* Chemically Tempered 'Fracture Mechanics'h e a l e d Glass * Coatings

'Thermally TemperedWindows GlassG l a s s 'Optical Properties

The window structural design and verification experience is presented for the Apollo commandand lunar modules. This repor t present s window design philosophy, design cri te ria, hardwaredescription, and qualification and acceptance test p rograms and discusses the problems en-countered and solutions developed in these a re as .

18. Distr ibution Statement

The struc tura l characteristic s of glass are not generally well understood by designers. Theoptics and instrument glass covers were not considered to be st ruc tur al components and thuswere not normally subjected to the design, qualification, and acceptance standards necessaryto preclude fai lures. These two factors contributed significantly to window problems on bothApollo spacecraft.

19 . Security Classif. (o f this report) 20 . Security Classif. (of this page)None None 21. NO. of Pages2422 . Rice*

Domestic, S2.75Foreign, 55.25


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Section PageFRACTURE MECHANICS ANALYSIS OF GLASS . . . . . . . . . . . . . . . . . . 17

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18ONCLUSIONSREFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

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.TABLES

TableI LUNAR MODULE WINDOW TEST RESULTS . . . . . . . . . . . . . .II LUNARMODULEGLASSMORTESTSPEC IMENMATR IX . . . . . .m RESULTS O F TEST PROGRAM FOR LM FORWARD WINDOWHEATER BUS BAR FAILURE . . . . . . . . . . . . . . . . . . . .

Figure

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FIGURES

Command module window designation . . . . . . . . . . . . . . . . .Command module side windows (1 and 5)Command module rendezvous windows (2 nd 4) . . . . . . . . . . . .Command module hatch window (3) . . . . . . . . . . . . . . . . . .Command module window-surface notation . . . . . . . . . . . . . .Lunar module window location . . . . . . . . . . . . . . . . . . . . .Cross sect ion of edge of LM forward window . . . . . . . . . . . . .

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Cro ss section of LM upper docking window . . . . . . . . . . . . . .Lunar module window coatings . . . . . . . . . . . . . . . . . . . . .Lunar module window flexure t es tsLunar module window heat er electr ica l connectionRevised forward window heater electrical connectionRevised docking window heate r ele ctr ica l connection

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APOLLO EXPERIENCE REPORTSPACECRAFT STRUCTURAL WINDOWSBy O r v i s E. P ig g a n d S t a n le y P. W e i s s

L y n d o n B. J o h n s o n S pa ce C e n t e rSUMMARY

The Apollo command module and lunar module window structural design and veri-fication experience is presented. The design philosophy, design cr it er ia , hardware,qualification and acceptance te st s, problems, and problem resolutions are discussed.

Instrument windows and guidance and navigation optics were not considered to bestruc tural it ems during their design and had to be structurally verified late in the pro-gram. In addition to the lack of a good window definition, which contributed to inade-quate struct ura l design and verification rigor , a general misunderstanding of thestruc tural characte ristic s of gla ss by design engineers w a s noted.A st ructural window definition is proposed to identify g las s s tru ctu res that shouldbe verif ied flightworthy. A spacecraft window is struc tural ly defined as any piece ofglass that is thermally or mechanically stressed and will endanger either the crew or

mission success if it breaks.Fractu re mechanics is the best means known today to understand, explain, andevaluate the struct ura l cha rac ter ist ics of g lass and to provide a struct ural rationalefor specifying proof-test requi remen ts. Use of fra cture mechanics in all future win-dow designs and evaluations is proposed to ensure acceptable window designs and toeliminate the misunderstanding of the structural characte ristic s of g lass.

I NTRO DUCT ON

The significant aspects of the structura l design and verification of the Apollo com-mand module (CM) and lunar module ( LM ) windows are presented in this repor t. Designphilosophy and cri te ri a, hardware configuration, qualification testing, acceptance test-ing, and prob lems encountered (with resolutions) ar e discussed. Also presented arebrief discussions of the struct ura l characteri stics of gla ss and of fra ctu re mechanicsanalysis of glas s. Information concerning Apollo window contamination is given inreference 1.


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A spacecraf t window is structurally defined as any piece of gl as s that is thermallyo r mechanically st re sse d and wi l l endanger the crew or mission succes s if it breaks.This definition was not used during the design of the CM and LM. The lack of definitioncontributed to inadequate st ructur al design and verification of some g la ss i te ms that arenot normally considered structural.Nine CM windows and four LM windows we re integral p ar ts of the Apollo space-cra ft primary pres sur e ves sels (habitable volumes). Also, the re were many gl ass in-strument covers in the CM and LM that were not par t of the prim ary pr es su re ve sse lsbut which sealed and protected the inst rumen ts fro m the spacecraft environment. Thest ruct ural integrity of these windows affected crew safety and mission succ ess to vary-ing degrees. Windows and glass s tr uct ures were not treated as a sep arate technologyduring the Apollo spacecraft design and development; theref ore , consistent design phi-losophy and design cr i te ri a were not used initially throughout the program. This situa-tion was recognized late in the Apollo Program during a NASA Lyndon B. Johnson space'Center (JSC), formerly the Manned Spacecraft Center (MSC), review of-glass structuralcharac teri stics and the analytical tools for design, analysis, and verification of glassstructure. It was al so determined that the str uctura l characteristics of glass were not

well understood by window designers. This review identified fr ac tu re mechanics as thebest analytical means available to evaluate the st ructur al integrity of glass and to pro-vide a structural rational fo r specifying proof- tes t requirements. Fra ctu re mechanictechniques were used to reevaluate the glas s s tr uc tu re in the CM and LM; this methodresulted in some modifications and some requalification and rever ification te st s.STRUCTURAL CHARACTER1 S T I C S OF G L A S S

Both tempered and annealed gl as s we re used for Apollo spacecraf t windows. Abrief discussion of the st ruct ural cha racter is ti cs of g la ss is included to introduce theCM and LM structural design philosophy and str uct ura l design cri te ri a.

Essentially, glass is a noncrystalline, supercooled liquid that is isotropic andelastic. Two conditions, tensile s tr es s and surface f l aw s , are required fo r a fracturein glass, which has an inherent strength (without flaws) in the range of 1 to 3 millionps i. All bulk gl ass , such as window mate ria l, has microflaws caused by manufactureand handling that reduce i t s strength. In addition, visual ly undetectable flaws throughthe thickness can be present i n annealed gla ss, which necessit ates a rigo rous flaw-screening program.Annealed glass exhibits a characteristic called static fatigue, which is a degrada-tion of allowable st res s as a function of t ime at load. Static fatigue is f l aw growth

caused by a combination of stress and environment and is sometimes ref err ed to asstress corrosion. Water o r moisture is a pr im e contributor to st re ss corrosion ofglass..

Tempering in gla ss is a pro cess that puts the surface in compression, thus elim-inating surface tensile st re ss es . Thermal tempering re su lt s in a parabolic r esidualstress distribution through the glass thickness with a compressive stress on the exter-nal surface approximately twice that of the tensile s t r e s s at the middle of the glass.

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Chemical tempering ha s a much higher compression-to-tension ratio , but the compres-sive layer is much thinner than with thermal tempering.The strength of both tempered and annealed glass is defined by the modulus ofrupture (MOR), which is the short-time breaking s t re s s of gl as s caused by bending ina moist environment. The MOR is generated f o r each type of gl as s by using abradedtes t samples to dec rea se the data sca tte r. The MOR values are used for design by thegla ss industry for both annealed and tempered glas s. The type of abrasion is assumedto represent the worst f l aw to be expected in the gla ss . The MOR of tempered gl as s

the glass if it we re annealed. The lower limit strength of polished samples is approxi-mately the sam e as for abraded samples.

.consists of the residual str es s in the glass at the flaw tip plus the inherent strength of .

C O M M A N D M O DU LE W I N D O W SThe CM has five double-pane windows: one hatch window, two side windows, andtwo rendezvous windows (fig. 1). Each of these windows cons ists of inne r-pr es sure -ve ss el and heat-shield windows as shown in figures 2 , 3 , and 4 . Four single-pane win-dows wer e located in the CM guidance and navigation (G&N) optics. In this rep ort , theCM windows are separated into four categories: inner str uct ure , heat shield, G&N,and inst rument windows. Because the G&N and instrument window problems were sim-ilar for the CM and LM, they a r e discussed i n a section that applies t o both.

SideSide windowRendezvous window

Hatch windowJ -> LRendezvousindowt x

Hatch windowRendezvous window

Side windowRendezvous window

- X

Outboard

L l n j e c t e d elastomerL Dr y nitrogen, 7.0psia

Figure 2 . - Command module sidewindows (1 and 5).

Figure 1.- Command module windowdesignation.

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Outboard, -Ablator Outboard

Vacuum dur ing orbi t

Dry ni trogen 7.0 psi

Outer si l ica pane t

L nn er aluminosi l icate panes aluminosi l icate Dry ni trogen,panes 7.OpsiaFigure 3 . - Command module rendez-vous windows (2and 4). Figure 4. Command module hatchwindow ( 3 ) .

Inn e r - S t r u c t u r e W in do wsDesign philosophy. - The st ru ct ur al design philosophy adopted for the CM innerstr uct ure , which is the CM prima ry pres sur e vessel, was that all windows would bedouble-pane windows fo r redundancy -and made of tempered g la ss. The cavity betweenthe two panes would be evacuated and backfilled to 7.0 ps ia with a dry inert gas. Thewindows were to be mounted to preclude installation stresses, thus restricting theglass loading to differential pr es su re . The fr am es were to be designed to r es tr ic t thewindow loads to those induced by pressure on the window.Design crite ria . - The stru ctu ral design c ri te ri a used fo r the Apollo CM inner-st ru ctur e windows wer e different fro m those used throughout the ae rosp ace indus try,including those used for the Mercury and Gemini spacecraft. In general, a fac tor ofsafe ty of 3 .0 w a s required, based on the MOR of the gla ss used. Also, the limit pres -

su re was usually the maximum pres su re that could exist during the mission, excludingany fa il ur es . The factor of safety required f o r the CM inner-structu re windows was1.5, based on the residual compressive s tr e ss in the glas s caused by tempering. TheCM windows were thus designed so that the gla ss surfaces had ze ro tension s tr es s atultimate load. Because the CM windows were double pane and the cavity between thepanes was at le ss than atmospheric pre ss ur e, the possibility existed that the windowscould have developed a prelaunch leak and be subjected to 14.7 psid l imit pres sur e inspace, assuming a check-valve leak. Because the p re ss ur e in the window cavity w a snot verified after installation, the cri te rion was adopted that the design limit pressurewould take into account a check-valve se al leakage failu re. There fore, al l CM double-pane windows were designed for a 14.7psid limit pres sur e ac ro ss each pane.Description. - All CM inn er- structure windows were made of aluminosilicate g la ss. and wer e thermally tempered to 25 000 psi MOR for the hatch and sid e windows and to23 200 psi MOR or the rendezvous window. The two panes in each of the windows hadthe s ame thickness: 0.23, 0.25, and 0.20 nch fo r the hatch, side, and rendezvouswindows, respectively . To accomplish the str es s- fr ee mounting and to provide a seal ,a silicone elastomer was injected around the edge of each pane and cured in place.This construction essentially potted the windows in thei r fr am es . After the sea l hadbeen cured, the volume between the double-pane inner-structure window was evacuated

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and backfilled with dry nitrogen to 7. 0 psia. Each inner-structu re window was coatedon both si de s with a high-efficiency, antireflection (HEA) oating (fig. 5).

Outboard -@ Magnesium fluoride coating

Figure 5. - Command module window-surface notation.

Qualification. - The qualification testsfor the inner- struc ture windows consistedprima rily of strength and deflection tes tsand mission-life pressure-leak te st s. Thestrength tests were conducted on windowsabraded to s imulate the maximum expectedsurface flaws. The design limit pr es su refor the inne r-structu re windows was14 .7 psid, which assumed a check-valveleak of one of the window seals. Fivestrength tests wer e performed on each con-figuration, the last of which was a failuretest. The pr es su re s at failure for the hatchwindow, side window, and rendezvous win-dow were 29.5, 29.7, and 32.5 psid, re-spectively. These strength te st s we reperformed on abraded, uncoated windows.It was assumed that the coatings did notaffect the window strength . Thi s assump-tion l at er was proved invalid, and additional qualification tes ts to de termine the effectsof the coatings were require d. These te st s ar e discussed in the section of this repo rtentitled "Problems. "

Heat- Sh eI W in ow sDesign philosophy. - The initial structural design philosophy for the CM heat-

shield windows w a s that al l windows subjected to the entry environment would be single-pane windows made of annealed, fused silica glass. However, an outer pane, al so ofannealed fused sili ca, was provided to protect each heat-shield window fro m micro-meteoroid impact while in space. These micrometeoroid windows were not requir ed towithstand the entry environment. Later in the prog ra m, the micrometeoroid windowswere omitted to reduce weight.Design cri ter ia. - The prima ry struc tural design criter ion for the CM heat-shieldwindows was a factor of safety of 1 . 5 , based on the sandblasted, annealed, fused sil icaMOR degraded fo r the time at load. The heat-shield windows are exposed to stress forapproximately 15 minutes during entry, and the primary loading is thermal. The timeused to determine the allowable st rength (considering flaw growth) was 1 hour. The

cri ter ion generally re quired by the aerospace industry is a design factor of safety of3 with a proof factor of 2 , based on the MOR of abraded (sandblas ted) gl as s. The stru c-tu ra l degradation of gl as s caused by s t r e s s and humidity can cause problems duringproof testi ng that can be misinterpreted . Also, the proof test can contribute to flawgrowth. The st ruct ural crite rion fo r annealed windows in the CM was a lower facto rof safety based on a more realistic structural allowable and a limit s tr e s s that includedworst-case conditions.

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Description. - Each of the five CM heat-shield windows consisted of a 0.7-inch-thick single pane made of fused amorphous si li ca . Each pane was insulated around thesilicone elastomer bonding agent). The insulation was flexible enough to allow the edgesedge by a the rm al insulator encapsulated by a 0.02-inch-thick Fibe rglas layer (using aof the glass to rot ate . The 0.080-inch-thick 17-4 P H steel fra me and retainer were de-signed so that the flat glazing could be instal led on the conical heat shield. A 0.05-inchgap was provided on all edges between the insulation and fra me to allow the shel l to con-tr act in the cold-soak environment without loading the glas s. Each heat-shield windowwas coated as shown in fig ure 5..

Qualification. - The qualification tests for the heat-shield windows consisted ofstrength t es ts fo r the appropriate abor t design conditions and a ther mal te st of the de-sig n entry condition from an initial cold-soak condition. Limit pr es su re tests to -5.5and +11.14 psid and ultimate pr es su re te st s to -8 .3 and +16.7 psid we re conducted oneach of the th re e window configurations. The analytical and tes t-t emp era ture differen-tials between the inner and outer surf aces of the heat-shield window we re 1600, 930",and 1130" F fo r the hatch, rendezvous, and side windows, respectively.

Quality ControlTo screen f l aws , each of the Apollo CM windows was subjected to an acceptanceprogram. The acceptance te st s for each window included thermal-shock te st s, pr es -su re tes ts, and visual inspections.The thermal-shock te st of the heat-shield windows consisted of heating the win-dows in an oven to 1200" F and quenching in 68" to 77" F water . The thermal-shocktest of the inner-st ructure windows cons isted of heating the windows in a sa lt bath to500" F and quenching in 68" to 77" F wat er. Although the st re ss es in a window causedby the thermal-shock test a r e difficult to predict, the test s cr ee ns f l aw s over the entiresurface of the window, which is important f or heat-shie ld windows because of the ther -ma l loading.Each inner-structure and heat-shield window pane was subjected to a proof-pr es su re acceptance test to scree n flaws. The proof-test pr es su re fo r all inner-st ru ctur e windows was 22 psi. For the heat-shield windows, the pr ess ur es we re 49 psifor the hatch window, 40 psi fo r the s ide window, and 73 psi fo r the rendezvous window.Windows were visually inspected to ens ure that they were not damaged subsequent to theproof test. It was assumed that the surfa ce would have to be visibly damaged fo r thewindow to be significantly flawed.The CM window proof-test requ irements wer e reevaluated subsequent t o the MSCstudy that determined a general lack of understanding of gla ss str uct ura l cha rac ter ist icsamong window designers. The proof -test requ irements were reevaluated using fr ac tu remechanics technology (ref. 2) and rational analysis techniques. Although proof te s t sfo r some conditions were not of the magnitude that would be used i f the windows werebeing designed (specifically , a proof test to 1 .5 of the operating st r e s s instead of avalue derived fr om fr ac tu re mechanics), the windows were determined to be acceptablefor crew safety and mission su cc es s.

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P ob1em sAs previously stated, an initial assumption w a s made that the optical coatingsapplied to the windows did not degrade their st ructural capability; therefore, the win-dow qualification te st s were conducted using uncoated, abraded windows. Thi s assump -

tion was proved invalid during coating crazing anomaly, and additional qualification test-ing of the coated windows was required to verify thei r flightworthiness.To improve the optical proper ties of the CM heat-shield hatch window for the lunarmul tispectra l photography experiment (S-158) on Apollo miss ions 13, 14, and 15, theavailable hatch windows were surveyed to find those with acceptable optical propert ies,Two heat-shield side windows with acceptable properties wer e found, and it was decidedto remo ve the coatings and cut hatch windows from these side windows. At that time,the hatch scient ific window could have no coatings because of the S-158 exper iment op-tical req uire ments. The "no coating?' requirement w a s removed later for th e Apollo 14mission because the S-158 exper iment on that flight w as rescheduled to the Apollo 15mission. The fi rs t side window was damaged during the polishing procedure to removethe coatings. The second window was cut to size, and both the magnesium fluoride andthe blue-red (BR) coatings were left on the window. In the thermal-shock test, the win-dow was heated in air to 1200" F fo r 45 minutes and immediately imm ers ed in room-temp era ture tapwater. The BR coating crazed and cracks propagated through thetension layer developed by the the rmal gradient of the glas s on the unders ide of the BRcoating. La te r, an enti re side window from spacecraft 014 was tested in a similarmanner to verify that the cutting of the window had not caused the crazing ; si mi la r re-su lts were obtained.

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In August 1970, three coupon te st s were conducted, one of which w a s thermallyshocked in room temperatu re water and two of which were slow cooled in air. The fol-lowing conclusions were reached.1. The BR coating crazing is temperature dependent.2 . The BR coating crazing is independent of window size or cooling rate.3. The BR coating crazing occurs during cooling.4. Window glass fra cture sometimes occurs when the BR coating cr az es , butdoes not occur without it.To determine the minimum temperature a t which the heat-shield window coatingcrazes, a se ri es of te sts was conducted on 2- by 7-inch samples cut fr om a space-cra ft 017 heat-shield side window. The re su lt s of these te st s indicated that the thresh-old temperatu re for BR coating crazin g is 450" to 500" F. A separate series of testsw a s conducted on 3- by 3-inch samples cut from a spacecraft 014 heat-shield rendez-vous window and al so determined that BR coating crazing oc cu rs at approximately500" F.The decision wa s made to continue using the coated windows and to conduct adelta qualification test program to verify the adequacy of these windows at newly pr e-dicted flight temp era tures. The maximum flight temperatu re on the inner sur face ofthe heat-shield window w a s predicted to be 362" F. Because the coating crazing w as

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determined to be caused by peak temperature and not thermal g radient, it was proposedthat the heat-shield windows be tested to a maximum temperature of 462" F to accountfo r variations i n oven temperature and the temp eratu re a t which the coating was depos-ited. Three heat-shield side windows were heated in an oven to 462" F to representworst-case flight tempe ratures, and the windows were then pr es su re tested on bothsides to verify adequate structu ral capability fo r the remaining flight loads. The pro -gr am was completed successfully and the coated heat-shield windows were considered. flight qualified.

A test pr og ra m was also conducted to verify that the coating proc es s did not sig-nificantly degrade the struct ura l capability of the tempered gl as s windows. Three hatchwindows were pr essur ized to 23 psig on each side separately to verify that the coatingpr oc es s did not degrade the windows and that the windows would pass the acceptancetest . The th ree hatch windows were pr es su re tested to failure, which occurred at 47,42, and 41 psig in a humid environment.LUNAR MODULE W I N D O W S

The LM has four windows in its pri mar y stru ctu re. The two forward windows arelocated in t h e cabin front-face bulkhead. The docking window is located in the uppersection of the cabin wa l l . Each window contains two panes wi th the cavity between thepanes vented to the external environment. In thi s configuration, only the inner pane ofeach window was subjected to the cabin pressure loading. The four th window in theprimary structure is the G&N window (optical telescope) and is discussed in anothersection.

D e s i gn P h i l o s o p h yThe st ructu ral design philosophy f or the LM windows w a s to provide a window ofminimum weight wi th maximum crew visibility, which led to the selection of the single-pane-window concept using chemically tempered glass. The design consisted of a singlest ructur al pane and an external pane fo r micrometeoroid and radiation protection.

D e s ig n C r i t e r i aThe basic st ru ctur al design requirement was that the windows had to sustain the

, environment imposed on them by the cabin with the pri mary loads attributable to thepr es su re and the rma l environment. The design limit pr es su re , the maximum cabinrelief valve pressure, was 5.8 psid. The design ultimate fa ctor of safety was 2.0.. However, the minimum acceptable MOR bending modulus (strength) of the glas s, sub-sequent to all manufacturing p roc ess es, w a s specified to be 50 000 psi. This strengthcorresponds to a factor of safety of approximately 7.2, based on the s t r e s s at design-limit pres su re . The ther mal environment imposed on the windows was +350" to -90" F.

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Descriptiontwo 25- by 28- by 24-inch triangular forward windows located in the front-face bulk-head of the forward cabin section we re installed in the vehicle in a plane oblique to thevehicle axes to provide an approximate visibility of 65" down and 80" outboard. Thesewindows provided the required visibility during the lunar descent, lunar landing, andlunar stay phases of the mission. Each window cons is ts of two panes separated by acavity vented to the ext ernal spac e environment. A c ross section of the edge of thiswindow is shown i n figure 7. The outer (nonstructural) pane w a s a micrometeoroid- and

.

An overhead docking window is located on the left side of the vehicle directly overthe commander's head. The window provides the commander with the requ ired visibil-ity during the initial phase of the descent to the lunar surf ace and during the fina l phaseof the docking maneuver. The docking window has approximately 65 squa re inches (5 by13 inches) of viewing area. The construct ion of thi s rectangular window, shown in fig-u re 8, was s imi lar to that of the forwardwindows. One exception w a s that i ts innerst ru ct ur al pane was not floating; i t wasbolted to the cabin skin by a metallic edgemember bonded to the chemically temperedglass. An additional difference was thatthe inner pane of the docking window wascurved to match the 92-inch diameter of thecabin ins tead of being fla t as the forward

tective cover' lp olycarb onate)

Mach'ned'vlndow

Cushion

Localmachined boss

Protective cover

Outboard1Removed before launchFigure 6. - Lunar module window Figure 7. - Cr os s section of edge of LMlocation. forwar d window.

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QualificationThe LM windows wer e originally qualified for space flight at both the componentand vehicle level. The component level te st s consisted of both functional and st ru ct ur alte sts . Functional te st s showed the optical and elect rical aspec ts of the windows, where-as the structu ral te sts included tests for leakage, shock, acoustics, thermal-vacuum,

and humidity.The vehicle level t es ts were accomplished as part of the LM stru ctu ral vehicletes t program. In th is te st p rogram, the windows were exposed to design environmentconditions, including pr es su re , vibration, and landing loads. No problems concerningthe windows were encountered during the component or vehicle level te st s.

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Problem sForward window test failure. - In December 1967 during a factory checkout pr es -sure test on the LM-5 ascent stage cabin, a forward window failed at a cabin pressu re

of 5.1 psid. The pane was completely destroyed, which is consistent with the failuremode of tempered glass. The subsequent investigation revealed no abnormal conditions;it was concluded that the most likely cause of the failu re was an undetected flaw in thegla ss. This pr ess ur e tes t of LM-5 was the first time the window gla ss had been exposedto the co rr ec t st re ss distribution. Previous structural acceptance tests on the windowsconsisted of a mechanical flexure test (fig. 10) on the forward windows and a thermal-shock test on the docking window. The mechanical flexure test was accomplished threetimes on both sid es of each panel (each corner on rol le r 1). The flexure test stressedthe pane to a maximum of approximately 30 000 psi, which demonstrated a factor ofsafety of approximately 5 over normal operating stress. However, this test did nots t ress the corners of the pane, which werethe areas of maximum s tr es s in the vehicle.Because the docking window w a s curved, athermal- shock acceptance test was requiredinstead of a mechanical flexure test. Thest re ss level experienced during therm alshock was approximately 20 000 psi, whichdemonstrated a factor of safety of approxi-

Roller mately 7 . To reestablish confidence in thestructural integrity of the LM windows,many engineering evaluation tests were con-ducted. A summary of the tests conducted,including the purpose, condition, and re-sults, is given in table I. These t est s showedthe str uc tu ra l integrity of the window gl as sand the window-support st ruc ture. Thestruc tural integrity was sufficient i f glasswith cri tic al defects was not used i n thespacecraft. Two st eps were taken to guar-antee that defective gl as s was not installed.

Appl ied load

The first st ep was to modify the acceptancetest to include a pre ssu re test on each struc-tural pane in a te st fixture simulating theFigure 10. - Lunar module win-dow flexure te st s.

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TABLE 1.- LUNAR MODULEWINDOW TEST RESULTS - ConcludedTest

Ther mal shock onforward panel

Temperature cy-cling on forwardand dockingpanels

compatibilityOxygen

PurposeTo compare the stru ctural

integ rity between chemi-cally tempered glass andtempered soda-lime glass.

To demonstrate the abilityof the inner pane ls to sus-tain a thermal cycling ata constant 7 . 7 psid.To demonstrate the stru c-

tural integrity of a for-war d window whensubjected to combinedenvironment of oxygen,elevated temperatures,high humidity, andpressure.

ConditionPanels were heated in a kiln

temperature of 464" F fora period of 20 minutes.Panels were dropped into a64' F water bath then pres-surized to failur e.

Panels were subjected to ther-ma l cycling, hot and coldsoaks, fo r 220 hours be-tween -90' F and 350" F.

The panels wer e exposed to atemperature of 225" -I 5' Fand a 90 to 10 0 percent re l-ative humidity in a pureoxygen atmos phere . Thetest was conducted with adifferential pre ssur e of25 . 0 psid for 72 hours and11 . 6 psid fo r 144 hours.

ResultsChemically tempered gla ss

survived test . Soda-limeglas s developed spalls andcr ack s on edge. Chemi-cally tempered gl ass panelfailed at 22 psid. Onanother run, results wereidentical. Chemicallytempered gla ss failed at85 psid.Panels survived test, whichsimulated all phase s of theLM flight including lunarstay.Panels survived tests.

Panels destructed at90 psid and 82 psid.

space craft installation. The acceptance tes t pre ssu re imposed on the panes w a s 25 psidfor the forward windows and 15 psid for the docking window. These pr es su re s c or re -spond to proof factor s of 5 and 3 above operating pr es su re fo r the two windows, andthese proof factors are consistent with requ irements based on fr ac tu re mechanics tech-niques. The docking window test pr essu re was rest ric ted to 15 psid to prevent yieldingof the metall ic edge member . In addition to the improved acceptance tes ts , the secondstep taken w a s to redesign the window, gl as s protective covers to provide maximum pro-tection fr om acceptance until launch. Before the acceptance tests were conducted, apolycarbonate protective cover, retainer, and glass pane were installed as a subassem-bly in the pres sure- test fixture. The press ure test w a s then performed to the properpressure for 2 to 5 minutes. After the pre ssu re te st , the subassembly w a s removedfrom tile fixtu re and an additional protective cover w a s installed on the opposite sideto provide a handling and transportation protective cover assembly. The outer coversremained on each pane until the outer pane was installed. The two covers were not r e-moved fro m the vehicle until jus t before launch, which provided maximum protectionagainst inadvertent damage fro m workmen i n and around the vehicle.

No vehicle windows have failed since the revised acceptance test procedures andnew protective covers were implemented. Eight windows (seven forward and one dock-ing) have failed in acceptance tests. The range of failing pr es su re on the forward win-dow was 5 .0 to 24 . 8 psid . The only docking window fa ilure occ urred at 11 . 6 psid.

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Window heater electrical connection. - A forward window bus ba r arced during thefactory checkout te st of the elect rical heater on L M - 8 . Failure analysis revealed theglass was spalled under the bus ba r, causing the arc when electrical power was applied.A detailed inspection conducted on other L M windows revealed discolora tions that indi-cated spalling.

IV

X

x x

39 240

The configuration of the electricalconnection to the bus ba r f or t he windowheater is shown i n figure 1 1 . The bus ba rwas a s t r i p of silver paste fused into theglass along the edge of two sides of eachwindow. The electr ica l wi re was subse-quently soldered to the bus bar . To pro-vide strain relief to the solder connection,an overlay of epoxy was applied. After theaddition of the solder and epoxy to the win-dows, the window was exposed to variousmanufacturing processes that elevated thewindow temperature. Because of the d i f -ference in the thermal coefficient of expan-sion between the glass, solder, and epoxy,residual st re ss es were induced into theglass. Analysis indicated that the magni-tude of thermal stress was sufficient to re-lieve the compressive st re ss es in the outerlayer of the chemically tempered gla ss,thereby reducing the strength of the glass.An extensive test program was conductedto verify the cause of the spalling. TheMO R test matrix is presented i n table 11.

L.

'

Detail A

Bu s ba r

fi1 \

Figure 1 1 . - Lunar module windowheater electr ical connection.

v VIx x xx x x

XX

Xx x xx x x

74 230 53 130

T A B L E U . - L U N A R MODULE GLASS MOR TEST SPECIME N MATRIX

Group

ECC and bus barSolder wire to bus b ar9 - l b pu l l t e s t o n w i reAdd epoxy to so l dere d jo intHot soak at 250" F , 1 hrHot soak at 150" F, 1 hr

IX

x xXx xx x

1 l X l Xl ex t e s t t hree t i m es a t18 367 ps iCold soak at - 1 0 " F , 1 h rRemove epoxy and wireMOR a v era g e des t ruc ttest, ps i

x x88 4 9 0 4 8 210 40 15(

VI1X

X

X

X

5 71 0

VIUXXXXX

X

XX

87 02(-14


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pr og ra m included specimens with solder and epoxy applied that we re exposed to theflexure and tempera ture environment and then the solder and epoxy were removed.These specimens showed that the orig inal streng th was reestablished i f the solder andepoxy were removed.The electrical connection to the bus bar was redesigned to eliminate the solderand epoxy. The new elec tr ica l connection was achieved by a beryllium copper spring-

loaded elec tric al contact as shown in fig ures 12 and 13. The window panes in L M - 7 andsubsequent vehicles wer e removed f ro m the vehicles and the solder and epoxy were re-and reinstalled in the vehicles.. moved from the panes. Following the rework, the panes we re proof-p ressure tested

Beryllium copper leaf Beryl l ,um

Section A A

To power source

Window struc ture -

Contact area (goldor rhodium platedl

'Lus bar onI inne r windowOuter window

Beryllium copper lealspring (coated)

Fiberglasl

Bery lli um coppercoil springI (coated1

l n ~ e r indow

Figure 12.- Revised fo rward windowheater electrical connection. Figure 13. Revised docking windowheater ele ctr ica l connection.G U ID A NC E A N D N A V I G A T I O N W I N D O W S

Th er e were four single-pane windows in the C M optical unit assembly and onesingle-pane window in the L M alinement optical telescope. These windows wer e notconsidered par t of the st ruc tur al subsystem during the s tru ctu ral qualification of theCM and LM because they wer e located in the opt ics of the G&N sy ste m. These windowsdid not receive rigorous st ru ct ur al qualification and acceptance testing consist ent withthe other windows because they were defined as optics rathe r than windows. In genera l,there was a lack o r decrea se in st ruc tur al qualification and acceptance t est rigor whenglass was used in design by disciplines other than str uc tu re s. Fo r futu re pro gra ms,

'

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Some of the limitations of the fra ct ur e mechanics analysis a r e discussed i n ref-erence 3. Because fr ac tu re mechanics is not generally recognized as a method f o ranalysis of g lass, limited gl as s fra ct ur e mechanics data a r e available. Wiederhorn atthe National Bureau of Standards is one of the few experime nte rs who has publishedgl as s fractu re mechanics data (ref. 4). Fr ac tu re mechanics data can vary with glas scomposition and environment and should be m easured f or the specific composition and. environment being evaluated.The fr ac tu re mechanics analytical techniques and proof-test methods presently

. developed ar e provided in reference 2 . Some mat eri al data and confirmation that f ra c-tu re mechanics is a valid technique fo r analysis of glass a r e provided in refe rence s 2and 3. A l l proof te st s of g la ss should be perfo rmed in a dry or vacuum environment toreduce o r eliminate flaw growth during the proof te st .

C O N C L U S I O N SThe experience gained and the problems encountered in the design and verif ica -tion of Apollo spacecraft windows demonstrate that th er e is a genera l lack of under-standing of gl as s structur al ch ar ac ter ist ic s by design engineers and that th ere is a needfo r adequate str uctura l c ri te ri a. The following item s should be included in future win-dow struct ural requirements.1. A l l spacecraft windows and gla ss st ru ctu re s, including optics and instrumentwindows, should be structurally verifi ed.2 . The rationale f o r proving the s tr uc tu ra l integrity of gl as s should include aproof acceptance test to sc re en flaws. Fr act ur e mechanics should be used to analyzegla ss that must ca rr y st re ss , and acceptance proof tes ts should be perfor med based onthe fracture mechanics analysis.3. Coatings and bonded attachments tend to degrade the str uc tu ra l integrity ofgl as s and their u se should be avoided. However, when coatings and bonded attachmentsare required, the windows should be thoroughly qualified in flight configuration andenvironments.

Lyndon B. Johnson Space CenterNational Aeronautics and Space AdministrationHouston, Texas, July 2, 1973914-13-20-13-72

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Apollo Experience Report Spacecraft Structural Windows

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