Apollo Experience Report Thermal Protection Subsystem

8/8/2019 Apollo Experience Report Thermal Protection Subsystem

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N A S A T E C H N IC A L N O T E N A S A TN D-7564

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APOLLO EXPERIENCE REPORT -THERMAL PROTECTION SUBSYSTEMby Jumes E . Puulosky und Leslie G, t . LegerLy12d012 B . Johlzson Space CenterHonst0~2,Texus 77058N A T I O N A L A E R O N A U T I C S A N D S PA CE A D M I N I S T R A T I O N W A S H I N G T O N , 0. C. J A N U A R Y 1 9 7 4


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~ - - . - .1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.D-7564

4. T i t l e and Sub t i t l e 5. Report DateAPOLLOEXPERIENCEREPORTTHERMAL PROTECTION SUBSYSTEM

James E. Pavlosky and Leslie G. St. Leger, JSC

January 19746. Performing Organizat ion Code

JSC S-38310. Work Unit N o.

I 7. Aut hor (s ) I 8. Performing Organization Report N o.

17. Key Words (Suggested by Author(s )* Reentry ShieldingAblative Mate rial

* Thermal Protection' Reentry Tra jectori es' spacecraft Structures

18. Dis t r ibut ion Statement

Cat. 31

I 9. Performing Organization Name and Address

19. Secur i ty C lass i f. (of th is repor t ) 20. Security Classif . (o f this page) 21. O. of PagesNone None 24

Lyndon B. Johnson Space Cente r

22. Price'$2.75

1 1 . Cont rac t o r Gran t No .Houston, Texas 77058 13. Type of Report and Per iod Covered

12. Sponsoring Agency Name and Address

National Aeronautics and SDace Administration 14. Sponsoring Agency CodeWashington, D. C. 20546 1The JSC Direc tor waived the use of the Inte rnat iona l System of Unit s (SI) for this Apollo ExperiencReport beca use, in his judgment, the use of SI units would impai r the useful ness of th'e re po rt o rresult in excessive cost.

15. Supplementary Notes

16. Abstract

The Apollo command module was the first manned spacecraft to be designed to enter the atmos-phe re of the ear th at lunar -ret urn velocity, and the design of the thermal protection subsystemfo r the resulting entry environment presented a major technological challenge. Brief desc rip-tions of the Apollo command module thermal design requirements and thermal protection con-figuratio n, and som e highlights of t he ground and flight testing use d for desi gn verifica tion ofthe system are prese nted. Some of the significant events that occu rre d and deci sions that weremade during the progr am concerning the thermal protection subsystem a r e discussed.


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CONTENTS

Sect on PageSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1THERMAL PROTECTION DESIGN REQUIREMENTS . . . . . . . . . . . . . . . 2THERMAL PROTECTION SUBSYSTEM DESCRIPTION . . . . . . . . . . . . . . 4FABRICATION. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

9ROUND AND FLIGHT VERIFICATION TESTING . . . . . . . . . . . . . . . . .SIGNIFICANT EVENTS AND DECISIONS . . . . . . . . . . . . . . . . . . . . . . 12

Therma l P rotect ion Subsystem Fabrication Concepts -Ti les Compared With Monolithic Ablator . . . . . . . . . . . . . . . . . . . 12Materi al Selection for Heat-Shield Substructure . . . . . . . . . . . . . . . . . 13The rmal Protec tion Subsystem Weight History . . . . . . . . . . . . . . . . . 13Experimental Research Flights. . . . . . . . . . . . . . . . . . . . . . . . . . 15Ablator Backup Program . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16Boost Protectiv e Cover . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17Water-Impact Capability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18REFERENCES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20

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TABLES

Table PageI APOLLO ENTRY DESIGN PARAMETERS . . . . . . . . . . . . . . . . 3

I1 APOLLO TPS FLIGHT VERIFICATION . . . . . . . . . . . . . . . . . 10I11 SUMMARY OF ENTRY CONDITIONS FOR OPERATIONAL LUNARMISSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

FIGURES

Figure123456

10

1112

13

Block I and 11design entry trajectories . . . . . . . . . . . . . . . .Command module ablator thickness . . . . . . . . . . . . . . . . .Heat-shield substructure assem blies . . . . . . . . . . . . . . . . .Structura l arr angemen t of Apollo TPS . . . . . . . . . . . . . . . . .Boost protective cover for Apollo command module . . . . . . . . .Penetr ations in command module heat shield

(Block 11) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Fabrication of forward heat shield sub stru ctu re . . . . . . . . . . .Injection of ab lator into honeycomb cel ls . . . . . . . . . . . . . . .Space-flight- mission simulation tes t (space craft 2TV- 1) inthermal-vacuum chamber at JSC . . . . . . . . . . . . . . . . . .

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9Entry te st conditions for flight test verification ofApollo CM heat shield . . . . . . . . . . . . . . . . . . . . . . . . 11History of ab lato r dens ity and TPS weight . . . . . . . . . . . . . . . 14Measured and predicted heating ra te s for Proj ect FIRE Iduring entry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16Spacecraft R-4 ablator material flight test . . . . . . . . . . . . . . 16

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APOLLO EXPERIENCE REPORTTHERMAL PROTECT1 O N SUB SYS TEM

B y J am es E. P a vlo sk y a n d L e s l k G . S t. L e g e rL y n d o n B . J o h n s o n S pace C e n t e r

S U M M A R YThe Apollo thermal protection subsystem was designed to prot ect th e commandmodule during en try at lunar- return velocities. Discussed in this report are the major

activities as sociated with the design, development, and flight tes ting of the subsystemand the significant technical and management decisions that evolved during the program.An extensive ground-based t est development program in plasma a r c heated facilitiesw a s conducted to charact eri ze the ablator thermal performance, followed by full- scal ethermal-vacuum te st s at the NASA Lyndon B. Johnson Space Center and a comprehen-sive unmanned flight test program. Continuous efforts wer e made during th e develop-ment of the Apollo thermal protection subsystem to reduce the weight of the subsystem;3these efforts led to the reduction in ablator material density from 66 to 31 lb/ft ,but the total subsystem weight always trended upward. The the rma l protection subsys-tem perfo rmed well on all operational lunar missions , and no anomalies requi ring post-flight investigation wer e recorded. It is concluded that an adequate technology nowexi sts to perm it the efficient design of ablative heat shields for entry at lunar -re turnvelocities.

I N T R O D U C T I O NThe Apollo Program was the third in a se ri es of manned space progra ms under-taken by the United States, and the Apollo command module (CM) w a s the fi rs t mannedspacec raf t designed to retur n to the ea rt h from the moon. Proj ec t Mercury and theGemini Program, which preceded th e Apollo Pro gram, wer e earth-orbital mission sthat resulted i n atmospheri c entry velocities of 26 000 ft/sec (inertial). The lunar -

ret urn traj ectory of the Apollo spacecraft resulted in an atmospheric entry velocity of36 333 ft/ sec (inertial), which creat ed an aerodynamic heating environment approxi-mately four times as severe as that experienced by the Mercury and Gemini spacecraft.In addition, the deep-space and lunar environments imposed stringent thermal-c ontrolrequirements on the Apollo spacecraft.The prec ise definition of the entry thermal environment and the design of a heatshield to protec t the entry module from the greatly incr eased heating environment wer ear ea s of pr ima ry technical concern in the design of the Apollo CM. A brief summary


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of the design, development, and testing of the Apollo th erm al protect ion subsyst em(TPS) and a discussion of the s ignificant technical and management decisions thatevolved during the prog ram a r e given in this re port . The Apollo spacec raft flew onnumerous earth -orb ital and lunar-landing missions, and the re were no problems oranomalies associated with the TPS. The suc ces s of the syst em can be attributed to thesomewhat conservative design philosophy that was adopted and to the rigo rous analyticaland test certification requ irem ents that we re imposed. The rep ort is concluded withsom e observations on design requireme nts and ablator perfo rmance.

THERM AL PROTECT1 ON DES GN REQUIREMENTSThe Apollo CM TPS was designed to prot ect the CM during entry into the e ar thatmosphere at lunar- retu rn velocities. The induced the rmal environment resultingfro m such an entry nec essit ates the installation of a heat-shield mate rial on the CM.Such material must be capable of sustaining (without excess ive ero sion ) the tem per a-tur es caused by the high heating rates on the blunt fac e of the en try veh icle and mus tals o provide insulation to minimize excessiv e sub struc ture tempera ture s.Early in 1961, a set of e ntry tr aj ec to ri es was developed by North American Rock-well (NR) that w a s based on a lift-to-drag ratio (L/D) of 0 .5 at a t ri m angle of attackof 33", a lun ar- ret urn speed of 36 333 ft/s ec, and an entry along a flightpath inclined40" to the equator. Fro m this set of trajectories, two trajectories that formed a pre-scribed entry corrid or were selected to size the heat shield. The tw o trajectories(fig. 1 ) initially se lected fo r TPS design we re called Block I. The ove rshoot boundary,designated HSE- A, was limi ted to a 5000-nautical-mile rang e and use d aerodynamiclift for a "skip-out" flightpath that maximized the heat load. The unde rshoot boundary,designated HSE-6, wa s of short duration to maximize the heating rate and was predicate don a 20g deceleration lim it based on a biomedical constraint. In addition, the Block I

design required that the abl ator sizing include the effects of the heating environmentduring ascent flight (with no specific tem pera ture limit imposed) and the ther mal envi-

Figure 1.- Block I and I1 design entrytrajectories.

ronment in space. In space, the temp era-ture requirement s for the ablator werelimited to a cold temperatur e of -260" Fand a hot tem pe ra tu re of 250" F.imum initial bondline temp erat ure at thestart of entr y wa s specified as 250" F, andthe limiting temper ature s at splashdownwere 600" F at the interf ace of th e ablato rand the stainle ss-st eel honeycomb str uct ureand 200" F fo r t he aluminum-honeycombpressure-vessel structure.

The max-

In the f a l l of 1963, the development ofthe CM was such that numerous weight-saving refinem ents in th e design and newoperation al logic in the pro gra m dictatedthe need f or a design change. Th is change,call ed the Block I1 design, provided anopportunity to resize the TPS. A guidance

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T HERMAL PROT ECT ION SUBS YST EM D ESCRI PT IO NThe externa l shape of the Apollo CM (fig. 2), like the Mercury (ref. 1)and Geminispacecraft (ref. 2), cons ists of a blunt entry fa ce with a conical afterbody tha t was de-

signed to minimize convective heating during atmosphe ric entry. The cente r of grav ityof the CM is offset fro m the axis of symm etry to gene rate the nec ess ary lift to satisfyentry corridor and range requirements.

The TPS comp rises the ent ire oute r shell of the CM and consis ts of an ablatorbonded to a stainless-steel struc ture that is fabricated in thre e su bassemblies (fig. 3).In addition to protecting the CM fr om the ther mal environment, the oute r shel l tr an s-mi ts the aerodynamic loads to the prim ar y structure during boost and entry and tran s-mits the hydrodynamic p res sur es to the prim ary stru cture during a water landing.Because of the uncertai ntie s concerning the flow field during CM elitry into the atmos -ph er e of the ear th, the magnitude of radiat ion heating, and the analysi s of ablator-to-metal junctions, the decision was made to fabr icate the entire C M TPS fro m ablativematerial. This decision was made, even though temp eratu re p redictions indicated thatrer adi ati ve metal' shingles would provide sufficient therm al protection fo r much of theCM conical afterbody.

The ablative mater ial se lected for the TP S is designat ed Avco 5026-39G and con-si st s of an epoxy-novalac r es in re info rced with quartz fib er s and phenolic microballoons.The density of thi s .mat eria l is 31 lb/ft . The ablator is applied in a honeycomb matrixthat is bonded to a stain less- stee l substructu re. The phenolic honeycomb is first bondedto the stain less -steel shell with HT-424 adhes ive, and then the abla tor is inserted intothe individual honeycomb ce ll s with a hypodermic device that is similar to a caulking

3

gun.

2500" F 7

S u r f a c et e m p e r a t u r e s

Re la t i ve w ind-

F o r w a r d h e a t s h i e l d

C r e w c o m p a r tm e n t64-------a-- h e a t s h i e l dAf t h e a t s h i e l d

Figure 3 . - Heat-shield substructur eassemb ies.Figure 2. - Command module ablatorthickness.

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The thickness of the ablato r va ri es with the local the rma l environment (ref. 3)and corresponding temperat ure profile, as shown in figure 2. Two typical c ro ss sec-tions of the TPS and pri mar y s tru ctu re are shown in f igu re 4. Section A - A representsth e stagnation heating a r e a where the total heat load is a maximum and req uir es anablator thickness of 2 .7 inches. Section B-B cuts through the leeward side where theheating rates a r e lowest and the ablator thickness is 0.7 inch. The space between theouter TPS shell and the cabin stru ctu re is filled with a low-density (3 . 5 lb/ft ) fibrousinsulation, TG15000. This insulation is used to reduce the heat tran sfe r between theouter shell and the cabin wall during space flight and, in part icul ar, during entry intothe ea rth atmosphere.

3

2 . 1

2.0

0.81. 5

1 i n .

i n .

--+Aerodynamic --flow ---

Section A -AAft heat shield

side

side

Section B -BForwa rd he a t s h ie ld9 n .5 i n .33 i n ..5 i n

Figure 4. - Stru ctural arrang ement of Apollo TPS.To accommodate th e heat-shield deformations that occ ur because of the thermalex trem es in spac e and entry heating, the conical section of th e heat shie ld is attachedto the aluminum cabin stru ctu re by means of a syst em of fibe r-gl ass slip str ing ers .Thi s attachment syst em (fig. 4) provides strain isolation between the inner and oute rst ru ct ur es and reduces heat conduction from the heat shield to the cabin. The thermal

control re quirements for the spacecraft in outer space necessita tes a relatively lowther mal absorptance-to-emittance rati o of 0.4 or the surface of the CM . This lowratio is achieved with a pressure-sensitive Kapton polyimide tape that is coated withaluminum and oxidized sili con monoxide and that is applied over the entire external

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su rf ac e of the ablator . The installa tion of a boost protective co ver ove r the conicalportion of the C M pre ven ts contamination of the thermal-c ontrol coating and the C Mwindows by aerodynamic heating during boost and by the tower jettison engine plume.The boost protective cover, which is attached to the launch escape tower and is jetti-soned with the tower befo re orbita l insertion, con sis ts of a la ye r of co rk bonded to afib er- gla ss cloth backing. Deta ils of the boost protect ive cover are shown in f igu re 5and are discussed in reference 4.In addition to the bas ic th ermal environment design considerations, the Apolloheat shield also has numerous penetrations and protuberances fo r the installation ofcomponents such as windows, reac tion control engines, antennas, and vent s, as shownin f igure 6.tion s such a s the re ces sin g of the components and the us e of densified ab lat ors in loca l

adjacent areas.Each of these discontinuities in the TP S required special design considera-

Cover

I n n e r structure I r e 0Brazed stainless-

C-band antenna

S-band antenna

Figure 5. - Boost protective cover fo rApollo command module.

Figure 6. - Penetrat ions in commandmodule heat shield (Block 11).

FA B R I CAT1ONThe Apollo TPS consis ts of an ab lato r in a honeycomb matrix bonded to a stainless-stee l substructure. The substructure is made up of thr ee suba ssem blie s (fig. 3) , which

are refe rred to, respectively, as the aft heat shield, the cre w compartment heat shield,and the forward heat shield. Each subassembly contains seve ral brazed sandwich panelsthat are welded together by N R using a tungsten iner t gas pro cess . A typical weld-asse mbly sequence fo r the forward compart ment heat shield is shown in figure 7. Thissubassembly co nsis ts of four lar ge braz ed panels and four tower-well fittings, which arefi rs t welded together and to which the fo rwar d ring and panel and the aft ring are addedto complete the assembly.

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B r a z e d q u a r t e r D a n e 1

Welded cone assemblyW e l de d q u a r t e r p a n e l a s s e m b l i e s 14)

Towe r-we l l f i t t ing

F o r w ar d r i n g

F o r w ar d r i n a

J-F ina l we ld a s s e mbly

Af

Welded assembly

Figure 7. - Fabri cation of forward hea t shield substructu re.A total of 4 1 brazed sandwich panels constitutes a shipset for each CM. The sepanels w ere manufactured by the Aeronca Corporation under subcontract to the N R Cor-poration. The first severa l sh ipsets were made of P H 1 5 - 7 M 0 stai nle ss steel; however,the lat er shipse ts were constructed of PH14-8MO sta inle ss st ee l because of the bett ercryogenic toughness of the 14-8 materi al.After the panels a r e welded by NR, the three subassemblies a r e sent to the AvcoCorporation fo r application of the ablator. The st ructure is first cleaned by scrubbingwith an abras ive detergent slu rry , and a primer coating is applied before the bonding

of the fi be r- gl as s honeycomb with HT-424 tape adhesive. The fiber-glass honeycombcor e sections are then fitted in place over the tape, and the edge members a r e posi-tioned at the sam e time. The assembly is vacuum bagged, and the adhesive is ovencured at 325" F for 1 hour . Inspection of the bonding of the honeycomb to the st ru ct ureis made by a nondestructive ult rasoni c transmiss ion evaluation. Any unbonded a reasa r e repaired, and then the assembly is ready for application of the abl ator into thehoneycomb. Thi s operation, te rmed "gunning, '' is the injection of the Avcoat 5026-39Ginto each cel l of the honeycomb by means of a specia l gun developed for that purpose.The cylindrical c artr idg es containing the ablator a r e dielec trically heated to 1 6 0 " Fand are inserted in the gun. When the nozzle is positioned over the honeycomb cell, a7


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solenoid-controlled air valve injects a blast of air into the cartridge and this entrainsthe ablator and carries it into the cell, filling it fro m the bottom to the top. Th er e a r eapproximately 370 000 cells in the honeycomb. A photograph of the gunning of thehoneycomb cells on the forward heat shield is shown in figu re 8. When all the cel ls ar efilled, the assembly is vacuum bagged and the abl ator is oven cured fo r 1 6 hours at200" F; then, it is postcured f or an additional 16 hou rs at 250" F. Then, the ent iresurface is machined on a numerically controlled tu rr et lathe to the design-thicknessrequ irements. The thick ness of the ablator is mea sur ed by an eddy-current techniqueat prese lected points, during machining as a pro ces s control, and after machining as afinal acceptance meas ureme nt. The machined TPS is radiographed to detect any defectsin the ablator, and rep airs a r e made i f necessary. Then silicone rubber ga skets a r einserted i n al l door openings, and various details (such as bolt plugs, molded ablatorpa rt s for the abort-tower wells, and fiber-glass she ar and compression pads) ar ebonded in place. After completion of these operations, the main ablator is checked formoisture content. A la ye r of thin, epoxy-based pore sea le r and a moisture-protectiveplastic coating then are applied to the surfac e to ens ur e sealing of th e porous ablator.


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A f t e r thi s operation, the final weight and center-of-gravity measu remen ts are made,and the heat-shield subassemblie s are returned to NR fo r installation on the spacecraft .Before the CM is shipped fro m the pri me contractor site to th e NASA John F. KennedySpace Center (KSC), the plastic coating is stripped off and the thermal-control coatingwith an adhesive backing is attached to the CM.

G R OU N D A N D F L I G H T V E R I F I C A T I O N T E ST IN GIn conjunction with detailed therma l and structu ral analyses, the Apollo sp acec raft

heat shield was c ertified f or manned operations by mean s of a n extensive ground and un-manned flight test prog ram. Th re e full-scale command modules (CM 004, CM 008, and2TV-1) were used f or the ground test program. The C M 004 vehicle (without the ablator,but complete in all other st ruct ural and thermal resp ects) was subjected to a radiant-lam p heating test that duplicated, i n real time, the predicted bondline temp era tur es ofthe h eat s hield fo r the Block I overshoot design entr y trajectory. The objectives of thistest (February 1966) were to evaluate the thermal-structu ral behavior of the TPS sub-str uct ure (including the windows and hatches and the s train-isola tion attachment s yst embetween the heat-shield subs truc ture and the cabin stru ctu re) and to verify the therm alcapability of the TF15000 insulation to limit t h e tem per atu re of the aluminum wall ofthe cabin to a maximum of 200" F.

Thermal-va cuum testing of spa cecr aft 008 and 2TV-1 was conducted in the 65-foot-diameter thermal-vacuum chamber at the NASA Lyndon B. Johnson Space Cent er (JSC).A photograph of space craf t 2TV-1 inside thi s cham ber, with one side of the spa cec raf tlighted by the simulated sun, is shown in figure 9. The se tests subjected the space craftto the tem per atu re ex tre mes and the vacuum conditions expected in spa ce flight. Thesolar t hermal energy was simulated withcarb on-a rc h ea te rs (fig. 9), and liquid ni-trogen in the chamber wall provided deep-space cold simulation.data on the expansion and contrac tion of thegaps between the heat-shield compartment s,the integr ity of the ablator when cold soaked,and a quantitative evaluation of t he distortionof the crew compartment heat shield.

These te sts provided

Figure 9. - Space-flight-mission simu-lation test (spacecraft 2TV- 1) inthermal-vac uum chamber at JSC.

Flight test ver ifi cat ion of the ApolloTPS was conducted on four unmanned space-craft. The test flight parameters are givenin table 11. The first two Apollo test flightsdemonstrated the perf orman ce of the TPSduring entry into the ear th atmosp here fromearth orbit. The first of thes e flights, mis-sion AS-201, use d the Saturn IB launch ve-hicle with spa cec raf t 009. The unmannedspacecraft was launched fro m KSC (Feb-rua ry 26, 1966) on a suborbital ballisticflight that gave an ent ry ine rtia l velocityof 26 482 ft /se c at a 400 000-foot altitude.

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T A B L E II. - APOLLO T PS FLIGHT VERIFICATION

AS-201 AS-202 AS-501 AS- 0 21B o o s t e rSpacecraftDate launchedM is s io n des cr ipt io n

Saturn IB009Feb. 26 , 1966Superc ircu la rent ry wi t hhigh heat rate

Inert ia l velocity at4 0 0 000 f t , f t / s ec . . . . .Relat ive velocity at400 000 f t , f t /sec . . . . .

Inertial flightpath angleat 400 000 It, deg . . . . .

Range f lown, n . m i . . . . .E n tr y t i m e , s e c . . . . . . .M a x im um hea t ing ra t e ,

Btu/f t2-sec . . . . . . . .T o t a l re f ere nce hea t ing2load, Btu/ft . . . . . . .

2 6 4 8 2

25 318

- 8 . 6 04 7 067 4

16 4

6889

Saturn IB0 1 1Aug. 25 , 1966Superc ircu la rentry withhigh heat rate

Saturn V01 7Nov. 9 , 1967Superc ircu la rentry at lunar-return velocity

Saturn V02 0Apr . 4 , 1 9 6 8L o wer v e lo c i t y beca us eof l a unch- v ehic l emalfunct ion

Entry condit ions

28 512

27 200

- 3 . 5 32 2 9 51234

8 3

2 0 8 6 2

36 545

35 220

- 6 . 9 319511 0 6 0

4 25

37 522

32 830

31 530

- 5 . 8 519351 1 4 0

19 7

27 824

a Al l m is s io ns w ere unm a nned.

The entry trajectory for th is fi rs t tes t flight was chosen to provide the highest heatingrates and, consequently, the highest ablator-surf ace tem per atu res and surface-recession rates that could be achieved with the Saturn IB boos ter. The TPS per for medwell during t h i s miss ion and qualified the subsystem fo r high-heating-rate en tr ie s fr omearth orbit .The second tes t flight was missio n AS-202, which again us ed a Saturn IB launchvehicle. This unmanned spacec raft was launched fr om KSC on August 25, 1966, f or asuborbita l flight with an entry tra jec tor y designed to give the maximum total heat load

that could be achieved fro m ear th orbit. Thi s flight qualified the TPS fo r this type ofatmos pher ic entry. In addition to qualifying the sy ste m fo r manned earth-orbita l en-tries, the AS-201 and AS-202 mis sio ns provided data fo r corre lat ion with the the rmalanalytical models used for thermal-perfor mance predictions at lunar-return entryvelocities.The Apollo 4 mission (also known as mis sio n AS-501) wa s the first flight test in-volving a Saturn V launch vehicle with a luna r module test article (LTA-1OR) and a

Block II configured C M (spacecra ft 017). The unmanned spac ecr aft was launched fromKSC on November 9, 1967, fo r a planned flight tim e of 8 hou rs 37 minutes. After tworevolutions i n ear th orbit, the S-IVB s tage was reignited for a simulated translunar

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injection burn. Shortly after the spacec raft separated fro m the S-IVB stage, the serv -ice propulsion subsystem was ignited fo r a short-duration burn to propel the spacecraftto an apogee alti tude of 9769 nautical miles . For approximately 4.5 hours during thecoast phase, the spacecraft w a s oriented with the CM hatch window pointed directlytoward the sun. This attitude cold soaked the thick ablator on the side opposite thehatch and achieved the maximum therma l gradient around the C M heat shield. Afterthe cold soak, the se rvi ce propulsion subsystem was reignited for a long-duration burnto accelerate the CM to entry conditions that represented the most severe combinationsof heating rate and heat load for the two extreme operational conditions that could possi-bly be achieved from a lunar- return trajectory. The entry traject ory resulted in anine rti al velocity at 400 000 feet of 36 545 ft/sec and a tr im L/D of 0.365. The CMlanded in the Pacifi c Ocean within 10 mil es of the predicted landing point, 1951 nauticalmile s down range f ro m the entry inter face at 400 000 fee t. A comparison of theent ry heating conditions of th is t es t flight (AS-501) with those expected during mannedoperational flights is shown in figure 10.

The actual mis sion flown w a s very close to that planned, and postflight inspectionof the re covered Apollo 4 CM indicated tha t the Block I1 TPS survived the simulatedlunar re tu rn entry environment satisfactorily. Sufficient flight data were obtained topermit a thorough evaluation of the thermal performance of the Block I1 TPS. Tempera-ture data wer e within design limi ts f or the flight. The response f rom the extensive in-strumentation was good, except for the heating rate measurements on the aft heat shield.Although the aft ablative heat shield w a s heavily cha rred and temperatur e data indicatedsurfac e temperature s approaching 5000" F, the measured surface recession w a s l e s sthan had been expected (based on ground test data) over all points on the aft heat shield.

The Apollo 6 mission (also known as mission AS-502) was the second mission touse a Saturn V launch vehicle with a lunar module test a rti cle (LTA-2R) and a Block IIconfigured CM (spacecraft 020). Thi s mission provided one further test of the TPS atlunar- return velocity. The only difference between the Apollo 6 mission and theApollo 4 mission in r egard to thermal confimration was that CM 020 had the Block IIthermal-control coating removed and it was-to be completely cold soaked during thecoast phase. The only C M change madefo r the Apollo 6 mission was that, for thefirst time, the unified crew hatch would bea pa rt of a CM undergoing test flight.

The unmanned spacecra ft wa slaunched fro m KSC on Apr il 4, 1968, for aplanned flight time of 9 hour s 57 minutes.Because of a malfunction of the S-IVB ineart h orbit, the serv ice propulsion subsys-tem had to be used t o achieve the pro-gramed apogee of 1 2 000 nautical Miles.Because of the resulting low fuel availabil-ity, the second firing of the s erv ice propul-sion subsystem was inhibited and the CMachieved an iner tia l velocity of only32 830 ft /se c at the entry int erface of400 000 feet. The heating conditions

Design 209 unde rs hoot , HR-1r

m AS-202 be s ign ov e rs hm1 ,HL-1AI I I Js 0 - m a M 35 40 45xld

(D 2Reference heat load, Btu/ f t

.--x

Figure 10. - Entry test conditions forflight te st verif ication of Apollo CMheat shield.

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achieved on this flight are shown in table I1 and figure 10. It can be seen that the ref-erenced heating ra te wa s only about half that experi enced on the Apollo 4 light, and the2total re ferenc ed heat load was 10 000 Btu/ft less.

Apollo 8/103 AS-503Apollo 10/106 AS-505Apollo 11,407 AS-506Apollo 12/108 AS-507Ap o ll ~ 3/109 AS-508Apollo 14/110 AS-509Apollo 15/111 AS-510Apollo 16/ ?12 AS-511

The unmanned flights provided te st ve rifi cat ion of the Apollo TPS for both ea rth-orbital and lunar-return missions. The measured data obtained fr om these flights andfrom the first tw o manned flights (AS-205 and AS-503) w er e used to corr ela te the analyt-ical models used for the requir ed certification analysis.

ft/sec35 00034 96835 02434 956

.~__._

34 88434 99634 92835 502

A summary of th e actual ent ry conditions fo r the manned lun ar landing miss ion s(Apollo 8 to 16 mission s) is given in table 111. A s indicated in the su mmar y, the maxi-mum down-range entry distance wa s 1500 nautical m ile s compared with the e stabli shedBlock I1 design requlrement of 3500 nautical mil es. These re su lt s indicated a crewpreference f o r a sh or ter down-range distance. The sho rter down-range entry distanceresulted in a maximum in tegrated hea t load of 26 500 Btu/ft , which is appreciably less2than the design r equ irement of 44 500 Btu/ft .

2

TABLE In. - SUMMARY OF ENTRY CONDITIONS FOR OPERATIONAL LUNAR MISSIONS

I EntrvMission/vehiclc relative,I Entry#elocity,inertial,ft/sec

36 22136 31436 19436 11636 2 1136 17 036 09636 090

Entryangle,inerticl ,

deg-6.48-6. 542-6.483-6.50-6.49-6.37-6. 51-6.49

D. 300.305. 300.309,290, 2 8 0.290.286

Range,n. mi .

12921295149712501250123411841190

Entry time,sec~-86887192981 5835853I7 881 4

Reference q,Btu/ft -sec2

(a)296296286285271310289346. ~

'Reference heating ra te .'Reference heat load.

S I G N I F I C A N T E V EN TS A N D D E C I 'S I O N ST h e r m a l P r o t e c t i o n S u b s y s t e m F a b r i c a t io n C o n c e p t s -

T i le s C o m p ar ed W i t h M o n o l i t h i c A b la t or

Reference Q,stu/rt2

(b)26 14025 72826 48226 22425 71027 11125 88127 939

The TPS concept submi tted initially by NI? to design and manufacture the Apollospa cec raft consisted of ablative tile s made fr o m phenolic-nylon mate ria l bonded to a

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honeycomb-sandwich su bs truc tu re made of aluminum. The substructure was to be builtby the pr ime cont racto r, and the design, fabrication, and installation of the ablativetiles we re to be accomplished by a subcontrac tor. In April 1962, a subcontractor wa schosen to supply an abl ative sy ste m consis ting of molded tiles (bjpically l-foot square)of Avcoat 5026-22 ab la to r bonded to a stainless-ste el substructur e. At approximatelythe s am e time (April 1962), recovered heat shields fro m Pro jec t Mercury wer e foundto have experienced debonding of the tiled ablative cen ter plug. This fact, together withthe uncertainty regarding the thermodynamics at the joints between the tiles, led to age ne ra l lack of confidence by both NASA and NR in the tile method of ab lator application.Consequently, NASA inst ruc ted NR to conduct an alternat ive fabrication study of theabla tor installation method being demonstrated successful ly at that ti me on the Geminispa cec raf t. The Gemini heat shield consisted of a fiber-glass honeycomb core filledwith an elastomeric ablator. A s a re su lt of this study, a lower viscosity Avcoablator (designa ted Avcoat 5026-39) wa s developed tha t could be applied i n a monolithicfashion to a phenolic fiber-glass honeycomb having a cell size of 3/8 inch. The fiber-gl as s honeycomb wa s first bonded to the stainles s-st eel sub str uct ure with HT-424 ad-hesive, and the indiviaual honeycomb cells were then filled with the abla tor. Initially,the cells were filled with the ab lator by tamping the dr y ablato r into the open cells,then curing the en tir e TPS inst alled on the vehicle. The tamping operation, however,caused considerab le coiicern with respe ct t o quality a ssu ran ce and the possibility ofdamaging the subs truc ture . Finally, the ablative mat eri al composition wa s modifiedso that it could be gunned in a mast ic for m (fig. 8) into the honeycomb ce ll s. Althoughthe monolithic ablator in a honeycomb ma tri x did provide a desirable fail-safe feature,it als o re sulte d in longer manufacturing schedules and req uired additional inspectionprocedures.

M a t e r i a l S e l e c t io n for H e a t - S h i e l d S u b s t r u c t u r eStainless steel was chosen in prefe rence t o aluminum for the TPS substructurebecause of the fail-safe cha rac ter ist ics provided by a higher-melting-point alloy in theevent of a localized los s of the ablator. The PH15-7M0 alloy wa s the alloy originallyproposed by NR because of its high tensile strength (Ftu M 200 000 psi at room tempera-

ture) and brazing compatibility. The initial heat shields we re fabricated fr om this alloy.However, furth er investigations revealed that the mater ial became brittl e at low temper-at ur es and the f rac tur e toughness was unacceptable. The temperature criterion at thistime f or spacecraft during space flight w a s i-250" F. Because of thi s fact , anothermater ial with better fracture toughness at -250" F was sought and the alloy PH14-8M0(vacuum melt ed) wa s s elected to r ep lace PH15-7MO. The PH14-8M0 exhibited out-standing fr ac tu re toughness throughout the tempera ture rang e of -250" to 600" F.However, it was a relatively new alloy and an extensive development pe riod w as re -quir ed to define the optimum welding and brazing pr oc es s specifications.

T h e r m a l P r o t e c t i o n S ub s y st e m W e i g h t H i s t o r yThe ablative mater ial initially selected in April 1962 fo r the Apollo TP S was3Avcoat 5026-22, which had a density of 66 lb/ft . The predic ted TPS weight with th ismaterial w a s 1684 pounds. Shortly ther eaft er, improvem ents (which included the addi-tion of microballoons) we re made to the materia l so that, by the end of 1962, a

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low-density ve rs ion of the ma te ri al (designated Avcoat 5026- 39, with a density of35 lb/ft ) wa s incorporated in the TPS. This represented a density reduction of 47 per-cent, but the corresponding reduction in predict ed sys tem weight was only 20 percent.The low rate of system weight reduction wa s caus ed by the inclusion of additional re -quir emen ts (primari ly the boost heating environment? which had been overlooked duringthe initial design phase. In the ye ar s that followed (afte r 1962), so me fur the r minorimprovements w ere made in the ablative materi al, which culminated in a mater ial den-sity of 31 lb/ft ; but the predicte d sy ste m weight for the TPS continued to have an up-ward trend (fig. 11). The upward weight tr end and the ca us es (which have been tr uehistorically of all aircraft and space craf t projects ) can be attributed to the followingfactors.

3

3

1. The continually i ncreasing number of protuberance s on the outside moldlineof t he vehicle and a resulting inc rea se in th e local heating environment2. The mor e refined analytical techniques that replaced earlier gross

predictions3. The addition of mo re rig oro us thermal-c ontrol cri te ri a as the spac ecraft pro-gram progressed

Because of management concern about the incre asing s pace craft weight, an attempt wa smade in 1964 to reduce the spacecra ft weight. The Block I1 design, which resu lted fro mthe se changes, showed a de cr ea se in TPS weight of approximately 200 pounds (fig. 11).Th is was achieved by (1) the elimination of the effects of boost heating environment bythe introduction of a boost protective co ver that wa s jettisoned with the launch esc apetower (fig. 5), (2) the reduction in the down-range req uir eme nt fr om 5000 to 3500 nau-tical miles (which provided a more realistic operational requirement ), (3) a reductionin the maximum initial entry tem pera turefrom 250" to 150" F (by the us e of an ex-ter nal thermal-control coating), and (4) theremoval of some protuberances. -= 80Two other weight-saving modifica-tions to the TP S we re incorporated in 1968and were based on recommendations by

NASA. The first modification w as the r e -moval of seve ra l layers of nylon from thesoft insulation blanket installed between theTP S subs truc ture and the aluminum cabin.The removal did not compromise the th er -mal insulation performance, wa s accom-plished without causing any schedule delays,and resulted in a weight saving of 64 pounds.The second recommendation (a simple man-ufacturing change) eliminated the applica-tion of the protec tive ename l paint thatacted as a moisture ba rr ie r. Because thether mal- con t rol coating subsequent ly

n20' 1962 I 1963 1%4 I 1965 ' 1966 I 1967 'Calendar year

lBoor700 r

1100Zml Predicted weight (Block IIBlock I (measured)Block II measuredll o 0 0 ' I I I 1 I 1 )1962 1963 1960 1965 1966 1967Calendar year

Figure 11.- History of abl ator densityand TP S weight.14


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applied to the ablator surface was also a good moisture se ale r, an ex tr a coating of thinepoxy-based pore sealer was applied instead of the enamel. Thi s sea ler , together withthe thermal-control coating, w a s sufficient t o keep the moi sture content in the ablatorbelow the specified 2 percent.

E x p e r i m e n t a l R e s e a r c h F l i g h t sThe test verification pl ans fo r the Apollo TPS included full- scale flight testing,which began in 1966. To gain confidence in the thermal p rediction and design methodsbeing used, the NASA Langley Resea rch Center (LRC) proposed two flight-researchexperiments. The first expe rimen t wa s the Flight Investigation of Reent ry Envi ronment(FIRE) proje ct, the objective of which was to measu re quantitatively the convective andradiative heating environment on a subsca le (2-foot maximum d iame ter) Apollo configu-ration at the cor rec t lunar-return entry velocity. The flight was justified because ofthe many approximations inherent in the aerothermodynamic theo ries and to substan-

tiate the magnitude of the therm al- radiation contribution to the total heating environ-ment. The FIRE payload was launched fr om KSC on Apri l 14, 1964, on a ballistictraje ctory that resulted i n an entry velocity (at a 400 000-foot a lt itude) of 37 900 ft/sec.Measurement s en so rs included beryllium heat-sink-type calori met ers for measuringthe total (convective plus radiative) heat flux and radiometers for measuring the ther-mal radiation fro m the shock layer . The data from these s ensor s were relayed toea rth by delayed telemetry a fte r the radio-frequency blackout period. Although the rewere a few flight anomalies, with r esp ect to perturbation i n the body motion, the FIREexperiment w a s a succ es s (ref. 5). A comparison of the measu red heating data on thesphe rical forebody of the vehicle with sev era l different theo retical predictions is shownin figure 12. These data provided confidence in the methods used fo r calculating theheating rates around the Apollo ent ry module.

The second subsca le experimenta l flight in support of the Apollo TPS developmentwas a lso proposed by LRC and had as its objective the verification of ablation per form-ance at the co rr ec t enthalpy and heating ra te s (corresponding to the Apollo undershootdesign (20g) trajectory). The justification for the experiment w a s that the high enthal-pie s as socia ted with the Apollo entry tr aject ory could not be duplicated in ground tes tfacilities. An added fifth stag e for the Scout booster was being developed by LRC, andthe five-st age Scout vehicle was proposed for the launching of t he test spacecraft (knownas the R-4). The entry nosecap used for the experiment was a spherical dish with aspherical radiu s of 17.4 inches and a diameter of 11.1 inches. The ent ry nosecap con-si st ed of Avcoat 5026-39G ablato r (1.25 inches thick) bonded to a stainless-steel stru c-tu re , which simulated the Apollo aft heat shield (fig. 13) . Instrumen@tion cons isted ofablation sen so rs and thermocouples. The R-4 spacecra ft was launched successfully bya five-stage Scout booster fr om Wallops Island, Va . , on August 18, 1964. To matchthe Apollo heating rat es at le ss than luna r-return entry velocity, the fifth stage of theScout vehicle was ignited late in the entry phase into the ear th atmosp here, so that highheating rates occurred at a lower altitude. A t the lower altitude, the heating rates ap-prox imate d those an Apollo vehicle would experience; however, the re sulti ng fr ee -st re am dynamic pr es su re s on the Scout vehicle were t hre e tim es higher than thepressures an Apollo CM would undergo on entry into the ear th atmosphere. The teleme-ter ed ablation-rate data fro m this flight indicated that the ra te s encountered were muchhigher than had been expected, par ticul arly during the lat er s tag es of entry , and, infact, result ed in complete erosion of the 1.25-inch-thick ablative mat eri al. The re was

I

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llOot800-?3 700-

crnC 600-Lrn._c1 500-S

Range of theoreticalpredictions - Convectiveplus radiative heating(circa 1962)

0 1I rc n

I I I I I h J1660 1665 1670 1615 1680 1685 1690nu," 1655 Time from laun ch. sec

Figure 12. - Measured and predicted heating rates for Project FIRE I during entry.

17-in-diamInsulationbulkheadrocket motor(Scout 5th stagelL C h a r r i n gablator nose,

1.25 in. thick

Figure 13 . - Spacecraft R-4 ablato rmateria l flight test .

much consternation as a result of thes e find-ings, and considerable analy ses and testin gwere done to convince the Apollo Programmanagement that the poor p erf orma nce ofthe R-4 te st s pacecra ft was a characteristicof the Avco 5026-39 ablator at high aerody-namic pres sur es, but that the high pre ssu resencountered in the R-4 test were not repre-sentat ive of the Apollo ent ry environment.From this, it can be concluded that, unlessa development flight can be made in a n en-vironment representative of the tru e envi-ronment, extraneous is su es can arise tocloud the resu lts and cause unnecessaryanxiety and work.

A b l a t o r B a c k u p P r o g r a mLate in 1962, concern grew'over the

increasin g weight of the Apollo TPS; i t wa s

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hoped that s ome other ablative fill er in a honeycomb matrix might offer a weight advan-tage. Accordingly, the spacecraft contractor w a s authorized t o conduct a comparativestudy of the ther mal performance and structural prope rt ies of five different ablativematerials (including the Avcoat 5026-39).Corning DC-325, which w a s used on the Gemini spacecra ft; a General Electric elasto-mer designated ESM-1000; the Emerson Electr ic Thermolag T-500-13; and a siliconemateri al with microballoons and eccospheres that was developed by LRC and w a s knownas purple blend. The re su lt s of the study, which lasted 6 months and included extensivetesting, showed there was no cl ea r thermal-performance advantage and, consequently,no cl ea r weight advantage to be gained by the us e of any one mater ial . In addition, itsoon became apparent from the thermostructural te st s that the low- temper ature require-ment of -250" F could not be tolerated by the el astome ric mat eria ls and, as a result,the Avcoat 5026-39 w a s retained as the mainstream material.

The four other mat eria ls wer e the Dow

By the end of 1963, a renewed concern grew ove r data that indicated the possibil-ity of an adve rse the rma l perfo rmance of the Avcoat ablator when it was subjected tohigh aerodynamic pre ssu res . Therefore, another backup progra m w a s undertaken forJSC by the Emerson Electric Co. , which used a lighte r version of the thermolag mate-3rial (known as T500-111) that had a density of 35 lb/ft . The material w a s well char-acter ized and was adaptable to the Apollo manufacturing scheme; however, noworthwhile thermal-performance advantage could be demonstrated. Also, the evalua-tion of the flight dat a fro m the Scout R-4 est in August 1964 finally eliminated the con-ce rn ove r the aerodynamic s hear sensitivity of t he Avcoat ma ter ial and the backupprogr am was terminated.

B o o s t P r o t e c t i v e C o v e rOriginally, the Block I TPS included approximately 0.12 inch of additional ablator

thickness to allow fo r the ch arr ing that would occur during vehicle exit flight. In Octo-be r 1963, the Apollo Program Manager agreed to a design change that incorporated aboos t pro tec tive cover over the conical portion of the CM (fig. 5). The boost pro tec tivecover w as attached to the launch escape tower and w a s jettisoned with the launch escap etower (ref. 4). The change resulted in the following design improvements.1. A reduction i n ablator thickness and weight2. Provi sion of a cover fo r the CM windows during boost, which eliminated pos-

sib le contamination of the gl as s by exhaust products f rom the tower jettison motors3. A thermal-control coating tha t could be applied to the outside surface of the

ablator to limit tempe rature extre mes during space flightThe temperat ure ext reme s without a coating we re of the or de r of k250" F; and test shad shown that, because of the difference in coefficients of t hermal expansion betweenthe ablator and stainless-steel substructure, the ablator cracked i f soaked at a tempera-ture less than -170" F.

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W a t e r - I m p a c t C a p a b i l i t yOriginally, the Apollo CM wa s designed to descend on land, and a deployable aftheat shield w a s attached to the CM pri mar y str uct ure by mea ns of energy-attenuatingst ru ts . In the spring of 1964, the pr im ar y mode of landing was changed f ro m land towater, and the deploying mechanism for the aft heat shield was deleted because it wasnot needed for w ate r landings. No modifications we re made to the heat-shield substru c-tu re to accommodate the wate r landing because it wa s believed that thi s condition wouldbe no mor e critica l than the 20 g design entr y condition. However, beca use the analyt-ical prediction of th e hydrodynamic pr es su re s and the st ruct ural resp ons e of the vehicl eto the se pre ss ur es was not understood completely, a qualification test program w a splanned. The test plan included the dropping of full-scale test articles into a watertank from a pendulum ri g to combine horizontal and ve rti ca l veloci ties. The first suchtest w a s conducted on October 30, 1964, with a test vehicle designated boilerp late 28.The test conditions simulated a nominal vertica l descent on thr ee parachutes(V = 28 ft/sec) in combination with a maximum horizontal wind design velocity of

28 . 5 knots (48 ft/sec). The test resulte d in extensive failure of the aft heat shield sub-str uct ure and of the cabin aft bulkhead, and the tes t sp ace craf t sank within 2 .5 minutesafter contact with the wate r. The subsequent invest igation showed that the re w a s a lackof understanding with r espec t to the landing c ri te ri a and the magnitude of the wate r-impact pres sure s. Therefore, the landing criterio n was established on a probabilitybasi s using a Monte Carlo random selection and a combination of th e se ver al variab les(wind velocity, wave slope, and s o for th). An extens ive 1/4-scale-model drop-te stprog ram was instituted at the spacecra ft contractor facility and at LRC to obtain realis-tic impact-pr essure data. The aft heat shield substru cture was redesigned to accom-modate the increased pres sure s, resulting in an increased TPS weight of approximately2 3 0 pounds.

V

C O N C L U D I N G REM ARKSThe Apollo spac ecra ft has flown on seve ra l earth -orbita l and lunar-landing mis-

sions, and the ther mal protection subsystem has perfo rmed well on all missions. Thesu cc es s of the syst em can be attribu ted to the co nserva tive design philosophy and rig-or ous development and verification testing that was conducted. Experie nce with lunarmissions has shown that the spacecraf t crewmemb ers pref er short-range entr ies andthat the command module can be guided prec isely to the middle of the entry corridor.This resul ts in a heating environment that is much less severe than that fo r which theth erm al protection subsystem was designed. Entry down range for the manned lun ar-landing missions (Apollo 8 to 16) was actually 1500 nautical mile s for lunar-returnent ry, compared with the design requ irement of 3500 nautical mile s. The resulti ngmaximum integrated heat load was 26 500 Btu/ft , compared with 44 500 Btu/ft whichwa s used in the design of the therm al protection subsys tem.2 2

The following are other maj or conclusions derived fro m the thermal protectionsubsystem design and application experience.1. The change fro m the proposed tiled abl ato r to a monolithic heat shield initiateda lengthy development of a new manufacturing pro ces s, and the ablato r had to be

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modified to pe rmit inser tion of the abla tor into the honeycomb core by gunning. How-ever , the advantage of having the fail-safe feat ures of a monolithic heat shield embeddedin a honeycomb matri x ov errode the disadvantages of extended manufacturing schedulesand more costly inspection procedures .2. The down-range distance selected f o r the Block I design overshoot traje ctor ywas 5000 nautical miles; this distance w a s reduced to 3500 nautical mi les for theBlock I1 design. However, the the rma l protection subsys tem could not be flight teste dbeyond 2500 nautical miles, and the astronau ts had difficulty ranging gr ea te r than1600 nautical mil es down range. A s a result, the operational choice ha s been in favorof ranges l e s s than 1500 nautical miles, which mea ns that the th ermal protection sub-

system w a s overdesigned.3. The vario us screening and backup prog rams during the Apollo the rmalprotection development showed that (a)curr ent ablato rs of the sa me general density

range (35 o 55 lb/ft ) have comparable thermal performance, (b) simple analyses donot for m a sound basis for ass ess ing the thermal protection subsystem weight, and(c) flight testing in environments not represen tativ e of design conditions, as in the caseof the five-stage Scout vehicle with the R-4 payload, can be misleading and cause un-warranted concern i f not interpre ted correctly.

3

4. Thermal performance of the ablation material is only one of s eve ral cr it er iarequired to develop a therm al protection subsystem. Viewing the the rma l protectionsubsystem as a whole, majo r changes were made in the sys tem to improve inspectionacc ess , th ermal str es s, manufacturing, center-of-gravity control, and performance atsingularities; however, majo r changes in the s yste m were not neces sary to obtain bet-ter thermal performance from the basic ablator.5. Because of aerothermodynamic uncert ainties associ ated with the many pene-trations, cavities, and protuberan ces that were re quired in the heat shield, many of thesingularities wer e reces sed into the ablator and other protuberances were located in theleeward regions of sep ara ted flow. A ll of these regions we re designed with a fairamount of conse rvat ism.6. An adequate technology ex ists t o permit the efficient design of abla tor the rma lprotection syst ems for entry speeds as high as those associated with luna r retu rn. Con-

sid era ble technological experience has been gained in the design, testing, and ana lys isof such a therm al protection system. Considering the criti cal natu re of the ther malprotection subsystem, the investment of time and money in extensive ground and flighttesting is considered to have been worth the effort.

Lyndon B. Johnson Space CenterNational Aeronautic s and Space AdministrationHouston, Texas, October 16, 1973914-50-00-00-72

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REFERENCES1. Swenson, Loyd S. ; Grimwood, James M . ; and Alexander, Char le s C. : Th is NewOcean, A History of Proj ec t Mercury. NASA SP-4201, 1966.2. Grimwood, James M. ; Hacker, Barton C. ; and Vorzimmer, Pe te r J. : ProjectGemini. NASA SP-4002, 1969.3. Lee, Dorothy B. : Apollo Experience Report- erothermodynamic Evaluation.NASA TN D-6843, 1972.4. Dotts, Robert L. : Apollo Experience Report- pacecraft Heating Environmentand Thermal Protection for Launch Through the Atmosphere of the Earth.NASA'TN D-7085, 1973.5. Dingeldein, R. C. : Flight Measurements of Reentry Heating at Hyperbolic Velocity(Project FIRE). NASA TM X-1053, 1965.

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Apollo Experience Report Thermal Protection Subsystem

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