Apollo Experience Report Spacecraft Heating Environment and Thermal Protection for Launch Through...

8/8/2019 Apollo Experience Report Spacecraft Heating Environment and Thermal Protection for Launch Through the Atmosph

1/23

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

m

zc C F I LEC O P Y

APOLLO EXPERIENCE REPORT -SPACECRAFT HEATING ENVIRONMENTAND THERMAL PROTECTION FOR LAUNCHTHROUGH THE ATMOSPHERE OF THE EARTHby Robert L. DottsManne d Spacecrufi CenterHouston, T e x a 770.58N A T I O N A L A E R O N A UT 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 , D. C. M A R C H 1 9 7 3


2/23

1. Report No.NASA TN D- 7085

16. Abstract

2. Government Accession No. 3. Recipient's Catalog No.

An accura te definition of the aerothe rmodynamic environment of the Apollo spacecraf t was neces-sa ry to en sure the structur al integrity of the spacecraft during the boost phase of the mission.Without an adequate definition of the environment and a provision for thermal protection, the tem-peratur e levels and thermal gradients induced in the stru ctur e of the spacecraft by the boost-phaseaerothermodynamic environment could resul t in degradation of the str uctur al proper tie s of thespacecraft, degradation of the thermal-control coatings, and excessive ther mal stresses in thestructure. The techniques that were used to define the aerothermodynamic environment of theApollo spacecraf t during the boost phase and to predict the str uctu ral temper atur es of the space-craft are discussed in this report . Also, the wind-tunnel and radiant-heating tests that were usedto support the analytical predictions ar e discussed; the analytical predictions are discussed; theaccuracy of the boost-phase heating-analysis techniques is shown by comparing the techniques withflight data. The analytical techniques fo r predicting heating charac ter isti cs and st ru ct ural tem -pe ra tu re s of the spacecraft were adequate fo r predicting the temperatures of the Apollo spacecraftduring the boost phase. Also, the applicability of simila r techniques for future spacecraft designand analysis is indicated.

4. Title and Subtitle APOLLO EXPERIENCE REPORTSPACECRAFT HEATING ENVIRONMENT AND THERMALPROTECTION FOR LAUNCH THROUGH THE ATMOSPHERE OFTHE EARTH7. Author(s)Robert L. Dotts, MSC

9. Performing Organization Name and AddressManned Spacecraft CenterHouston, Texas 77058

12. Sponsoring Agency N ame and AddressNational Aeronautics and Space AdministrationWashington, D. C . 20546

15. Supplementary Notes

17. Key Words (Suggested by AuthorIsJ)

5. Report Date6. Performing Organization Code

March 1973

8. Performing Organization Report No.MSC S-350

10. Work Unit No.914-50-20-12 72

11. Contract or Grant No.

13. Type of Report and Period CoveredTechnical Note

14. Sponsoring Agency Code

Launch Vehicle ConfigurationsAerothermodynamicsThermal ProtectionAerodynamic Heating19. Secur ity Classif. (o f this report)None

18. Distribution Statement

20. Security Classif. (of this page) 21. No. of Pages 22. PriceNone 23 $3.00


3/23

APOLLO EXPERIENCE REPORTSPACECR AFT HEAT1 NG EN Vl RONMENT AND THERMAL

PROTECTION FOR LAUNCH THROUGH THEATMOSPHERE OF THE EARTH

By R o b e r t L. D o t t sM a n n e d Spacecraft C e n t e rS U M M A R Y

An accur at e definition of t h e boost-phase aerothermodynamic environment of theApollo spacecraft w a s required fo r the design of a the rmal protection s yste m that wouldensu re the s tru ctu ral integrity of the spacecraft during the boost phase of an Apollomission. Without adequate protection, the temp era ture s and the tempera ture grad ient sinduced in the spacec raft st ruc tur e by the boost -phase ae rothermodynamic environmentwould res ul t in an unacceptable degradation of the spacec raft st ructu ral prop erti es andthermal -contro l coatings, significant the rma l stresses in the structure, and excessivetemperatures in the pyrotechnic charges attached to the structure.The total Apollo boost-phase ther mal protection sys tem of the spac ecra ft is dis -cussed in this re port, and the techniques that were used to predict the boost-phase

aerodynamic-heating environment of the spacec raft are discussed briefly. In general,conservative design approaches were used; however, localized fail ures of se ve ral lunarmodule ad apter panels during the Apollo 6 (AS-502) missio n caused an extensive r eeva l-uation of the component that result ed in therma l redesign. Extensive ana lyses andground tests wer e performed on this component during the investigation of the AS-502anomaly. The res ul ts of the analyses and test s are discuss ed in detail. In addition,analytical predic tions of str uc tur al ther ma l response to the boost -phase environmentare compared with flight data.

INTRODUCTI O NThe Apollo spacecra ft is subjec ted to substantial boost -phase aerodynamic heatingduring launch through the atmosphere of t h e ear th. An accura te definition of the effectsof th is boost-phase aerodynamic -heating environment on the Apollo spacecr aft w a srequired to design a ther mal protection system that would en sure stru ctur al integrityof the spa cec raf t and subsequent operation of all spa cecraft syst ems and componentsaffected by thi s th ermal environment.


4/23

The design of the boost-phase th er -ma l protection syst em of the Apollo space-craft involved definition of the boost-phaseaerodynamic-heating environment; ther-mal analysis of the various s pacecr aft ele-ments during the boost phase to determinethe temperature histories of the elements;design of thermal protection sys tem s forthe spacecraft elements that would expe-rience excessive temperatur es during theboost phase; and ground te st s of the va ri -ous elements of the thermal protectionsyst em and flight tes ts of the integratedthermal protection system to verify theadequacy of the design.

The general boost-phase design con-side rat ions of the ther mal protection sy s-tem for the primary Apollo spacecraftcomponents are discussed in this repor t.Also, the techniques that were used todefine the aerodynamic-heating environ-ment are discussed; the basic thermal-analysis techniques and the ground andflight test s that were used to verify theadequacy of the design of the thermal pro-tection system ar e presented.SPACECRAFT BOOST-PHASE THERMAL-PROTECTION-SYSTEM CONFIGURATIONAND DE S IGN CONS IDERATlONS

The primary spacecraft elementsthat requi re boost-phase thermal protec-tion are shown in f igure 1. The elementsinclude the launch escape syst em (LES);the boost protective cover (BPC), whichcovers the command module (CM) duringlaunch; the service module (SM); and thespacecraft/l unar module adapte r (SLA),which covers the lunar module.

-Launch escape system

Command module

Figure 1. - Spacecraft launchconfiguration.

2


5/23

Lau nch Escape SystemThe purpose of the LES is to remove the CM rapidly from the vicinity of thebooster during a boost-phase abort. The LES (fig. 2) consis ts of a solid-propellant-rocket-motor assembly that contains three rocket motors attached to the forward por -tion of the C M by a tower st ruct ure. The CM BPC is attached to the LES towerst ru ct ur e and provides therm al protection fo r the CM during the boost phase. The

en ti re LES, including the BPC, is jettisoned near completion of a normal boost phaseo r an abort.The legs and cros smemb ers of the LES tower were covered with Buna-N rubb erinsulation that w a s sized to re st ri c t the maximum temper atu re of the horizontal andver tic al titanium leg me mbers to 600" F and the diagonal leg members to 800" F dur-ing the des ign boost -abor t phase. The therma l response of a typical ver tic al leg mem -be r (location A in fig. 2) during the AS-503 design boost-phase/abort environment(ref. 1) is shown in figure 3. External corkboard ablative thermal-protection m ate ria lwas used to pro tec t the tower-jettison mo tor, part of the LES motor, the power -systemsand instrumentation w i r e har nes s, and the str uctura l sk ir t from the boost -phase/abort

thermal environment. A typical the rma l response for the LES motor casing (location Bin fig. 2) during the AS-503 design boost-phase/abort environment (ref. 1) is shown infigure 4.

Power-systems andinstrumentationwire harness

insulation

Figure 2. - Launch escape system.3


6/23

IIII

I I I II /

Comma nd Mo du le

I II I1 , I , I 1 I 1

The BPC prevents the CM external thermal-control coating from exceeding 250" Fand provides protection from coating contamination during the boost phase. Also, theBPC protects the CM from the high-temperature exhaust of the LES motor dur ing aboost-phase,abort. A side view of the BPC is shown in figure 5. The BPC is composedof two sections: a hard cover that is attached to the LES tower legs and a soft coverthat is connected to the trailing edge of the har d cover.fiber -glass/honeycomb sandwich that is covered with cork thermal-protection materi al(0.3 inch thick). A typical therma l response for the BPC during the AS-503 designboost-phase/abort environment is shown in figure 6. The temperatur es shown are for

The hard cover is a phenolic

F i b e r - g l a s s r i n g H o n e y c o m b - c o r e d l a m i n a t e d -H a r d - c o v e r s e a l 7r f ib e r- gl as s p an el r L a m i n a t e dLM e a r - s n i e i a

0 . 0 3 0 - i n - t h i c k c o r k b o ar d a b l a to rf iber-glasslhoneycomb sandwich

-ES je t t ison;at 312000 flI ;a t 24000 fl50 I !

1 I I I l l I 1 1 1 . 10 50 100 150 200 250 300 350 400 450 500 5% 6WI II

Elapse t im e, sec

4

Figure 5. - Boost protective cover. Figure 6. - Thermal response of the hardBPC during the AS-503 design boost-phase/abort environment.


7/23

the outer and inner facesheets of the hard-cover fiber -glass/honeycomb sandwich(ref. 1). The so ft cover consis ts of a 0.008-inch-thick Teflon-impregnated gla ss cloththat is reinforced on the outer surface with a 0.0095 -inch-thick nylon fabr ic. Corkthermal-protection material is bonded to the outer surface of the nylon fabric.

Serv ice ModuleThe SM (fig. 7) is a cylindrical unit that is composed of an oute r she ll and aninner concent ric core. The resulting annulus is divided into six bays by radial beams.The outer shel l is fabricated of 1-inch-thick aluminum-honeycomb panels and is cov-ered with cork thermal-protection materi al (0.020 to 0.155 inch thick), except in theregions where the environmental control system (ECS) and the elec tri ca l power syste m(EPS) radiators are located. The cork provides protection fro m boost heating and SM

Service propulsion

Figure 7. - Service module.

reaction control system (RCS) plume-impingement heating. The cor k thicknes sw a s sized to maintain the SM honeycombstru ctu re below the design temperaturelimi t of 400" F during boost. The four RCSmodules, which protqude fro m the su rfaceof the SM shel l approximately 10 inches,cause flow disturbances that r esult inincreased boost-heating rates in the vicin-ity of the modules.on a representative portion of the SM su r -face a r e shown in figure 8. The longitudi-nal position on the SM is defined, in inches,by Xs.corresponds to the lower end of the SM;Xs 376 inches corresponds to the top ofthe SM.defined i n te rms of angular position me as -ured in degrees from the +Y xis. Thethicknesses re sul t from designing for theheating variation over the surface (ref. 1)and from maintaining a maximum tempera-ture of 400" F at all surf ace points whileminimizing the co rk weight,

The cork thicknesses

The position X = 200 inchesS

The circumf erentia l position is

5


8/23

0.155

0.06

0 .03 0.03 o ~ o ~ ~

::I\ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \ \

Radial Ibeam 1 +Y axis 101................................................................................ Bare surface Note A ll dimensions ar e i n inches.

0.02

0.04 0.02 \ .03

umbi l icalQ-i0.06

0.155

I 0.06l"h" ne!

0.03

\\ \ \ \ \ \ \ \ \ \ \ \ \m

Radialbeam 2 I R+Z axis l9O"l b e a m

~ = 376

= 355

* = 200I

Figure 8. - Block 11SM (AS-503 nd subsequent spacecraft) cork ther mal-protection thickness.6


9/23

Spac ecr aft/L un ar Mod u AdapterThe SL A structure is a truncated conical shell (fig. 9 ) that is fabrica ted of1 . 7-inch -thick aluminum -honeycomb panels with an outer cover of cor k ther mal-protection material (0.030 to 0 . 2 0 inch thick). Numerous items, such as hinges, an-tennas, lights, ac ce ss doors, and st ruc tur al joints, protrude from the surface of the

SLA (fig. l o ) , causing flow disturbances and localized areas of inc reased aerodynamicheating. The it em s were protected by cork o r fabricated of mat er ia l capable of with-standing the localized tempera ture ext reme s (ref. 1). The longitudinal posi tion on the SLAis defined, in inches, by Xend of the SLA; XA = 838 inches corresponds to the top of the SLA. The circumferentialposition is defined in te rm s of angular position measured in degr ees f rom the +Y xis.

The position X - 502 inches corre sponds to the lowerA' A -

Initially, the cork thickness w a s sized to maintain the SLA honeycomb structurebelow the design tempera ture l imit of 490" F. The 0,030-inch-thick cork (shaded s u r -faces in fig. 10) w a s added to Apollo 8 (AS-503) and subsequent adapters because of theAS-502 anomaly. The AS-502 SLA had a localized structu ral fail ure of seve ral panelsduring the boost phase; as a result of the subsequent investigation, the 0.030-inch-thickcork w a s added to the bar e sur fac es of the SL A to reduce the tempe ratu res and tempera-ture gradients in the SLA structure.

X A = 838 n.

-260 in.-Figure 9. - Spacecraft/lunar module adapter.

7


10/23

I+ Y axis tool Note A l l d i m e n s i o n s ar e i n i n c h e s + Z axis (90")- 0 . 0 3 - i n : t h i c k c o r k on A S - 5 0 3

and subsequent spacecraf t

Figure 10. - Cork therm al-protect ion thickness on the SLA.

8


11/23

THE S M A N D SLA BOOST-PHASE AERO DYN AMI C-HEAT1 NGENV I RONMENT

A boost-phase aerodynamic-heating computer program tha t was originally deve l-oped and used for Gemini spacecra ft boost-heating analysi s w a s used in defining theApollo spacecraf t boost -phase aerodynamic -heating environment.gr am characteri zed the flow of air over the vehicle with a normal shock at the nose andan isentropic expansion to local conditions at each spec ific body station. Ecker t'sreference-enthalpy method (ref. 2) w a s used in the computer program with the Blasiusskin-friction coefficient (ref. 3) for laminar flow and the Schultz-Grunow skin-frictioncoefficient (ref. 4) fo r turbulent flow. The local pr es su re P w a s determined by usingthe ratio PL/PT and the total pr es su re P behind the normal shock. The local the r-Tmodynamic st at e point w a s then defined using the entropy behind the normal shock and

The computer pro -

L

The ratio PL/PT for the SM and the SLA is a function of Mach number , ReynoldsL'number, vehic le angle of attack, and local body station. Wind-tunnel tests on scalemodels of the Apollo sp acecra ft launch configuration were perfor med to determine thepre ss ur e distribution over the vehicle su rface for var ious Mach numbers and Reynoldsnumbers at 0 angle of attack. Angle -of -attack effects on SM and SLA heating werenegligible because of the sm al l angle of attack (nominally le ss than 1" ) for the Apollovehicle during the heating portion of the nominal boost-phase tra jec tory. The wind-tunnel model and the pressu re- measurement locations a r e shown in figure 11. Curvefits of the pre ssu re- rat io data for the SM and the SLA are shown in figu re 12. Varia-tion of PL/PT with Reynolds number was small ; there fore, the data could be corre-lated uniquely with the Mach number. For a given vehicle traj ect ory and SM o r SLAlocation, tim e his to ri es for smooth-body heating rate were generated for a range ofsurface tempe rature s. The heating rates were used in a thermal-analysis computerprogram as a function of t ime and vehicle-surface temper atu re to calculate transienttemperatures at various loca tions on the SM and the SLA.body heating rate for a typical AS-202 SLA location are shown in figure 13. The sp e-cific tem per atu res used to generate the heating-rate curves of figure 1 3 were chosen tobrack et the expected SLA surface temperatures.

Time his tories for smooth-

Flowdirection

Pressure-sensor locations

L-I Centerline3- ___

Figure 11. - Wind-tunnel-model configuration and pres su re -measurement locations;the model is 0.045 scale.9


12/23

1.0-.a. 6

--

cL+ . 4 -\C Ld3",E . 2 -

5mc.--

. l o -e .M -E .M-zs=A0-- .M --0=._-

.02 -

.01 I I I IMach number0 2 d 6 a

Figure 12. - Ratio of local press ure to stag-nation pre ssu re pL/pT as a function ofMach number for the SM and SLA.

The increas ed heating that was causedby the various flow disturbances on the s u r -faces of the SM and the SLA was accountedfo r by using protuberance fac tors (ratio s ofthe smooth-body heating rate to the localheating rate). The fac tors 'were applied asconstant mult iplier s to the smooth-body

Local syfacetemperature = 40' F L/ /ocal surface, temperature i 540" F

1040' F

I ; I 1 1 1 I I J' u 20 40 60 80 100 120 140 160 180 200

Elapsed lime. sec

Figure 13. - Time histories fo r smooth-body heating rate fo r the AS-202 SLA(station XA = 775 in.).heating ra tes throughout the boost-phase trajectory. This conservative approach ofusing constant factor s ra th er than varying the fa ct or s with Mach number was used toensure adequate design safety margins for the protuberance-affected areas of the space-craf t. Subsequent Apollo flight da ta indicated that th is approach was overly conse rv-ative; there fore, Mach-number-varying protuberance facto rs were used later to verifythe adequacy of the SM and the SLA for off-nominal tra jec tor ies . The predictedprotuberance-factor contours that res ult fro m the RCS modules and other protuberanceson the surface of the SM are shown in figure 1 4 (ref. 5). The predicted protuberance-factor contours fo r two sections of the SLA surface are shown in figu res 15 and 16(ref. 5). The protuberance-factor contours we re estima ted by using the test dat a ob-tained from refer ence 6 and wind-tunnel da ta for the Apollo spacec raf t configuration.

10


13/23

X s - 316 ir

X s = 355 ir

x s = 200 IRad[,

IRadial beam 2 I I+? axis Radial beam 3Iearn 1 + Y axisFigure 14. - Typical interference-heating map of pro tube rance -factor contours fo r asect ion of the SM.

11


14/23

Rad ia l l oc a t i ons :0", 90",180", 270"

L i n e s o f c o n s t a n tp r o tuber anc e f ac to r

S LA h i n g e-.._

@-.0I

1

-Cableway

T h r u s t e rc o v e r

X A = 58 3 in .--I---

1.0 P r o t u b e r a n c efac to r

-L ines ofc o ns tanp r o t u b e r -anc efac to r

- 502 in .fl- AC e n t & l i n eFigure 15 . - Local interfe ren ce -heating map of protuberance - factor contours f or SLAhinges , thruster cover. , and cableway sys te m .12


15/23

-Lz

. X A = 58 3 i n- L i n e s o f c o n s ta n tProtuberance factor

S c i m i t a r a n t e n n a

Ii rotuberance factor* 502 i n .fi- AIRadial locations * 45. 135, 21y, 31.50 IRadial locations - 45: 225

Figure 16. - Local inte rference -heating map of p rotuberance -factor contours forSLA/LM attachment cover and sci mi ta r antenna.BOOST-PHASE THERMAL RESPONSE OF THE SMAND SLA STRUCTURESThermal -Mathemat ica l Models and Assumpt ions

The thermal response of the structure was determined by using the surface-heating-rate histories as boundary conditions for one- and two-dimensional the rma l-mathematical models of the st ructure. Most of the SM and SLA structures arecomprised of aluminum -honeycomb -sandwich panels that can be analyzed adequatelyby using a one -dimensional model. The thermophysical prope rti es of the honeycombpanels wer e obtained from refer ence 7 . A typical one-dimensional the rmal model isshown in figur e 17. One-dimensional mathematical models, which a r e s im ila r to themodel shown in figure 17, and two-dimensional mathematica l models of m or e complexSM and SLA str uc tu ra l elements (joints, bulkheads, et cetera) were used for all tem-pera tur e predictions. Extensive radiant -heating tests were performed to ensure thatthe one- and two-dimensional mathematical models we re adequate. Also, the te st swere used to determine the performance of the co rk thermal-protection mate ria l thatwas used extensively on the su rfaces of the SM and the SLA.

13


16/23

T h e r m a l -m a t h e m a t i c a l -model nodes-

30028026 0240220

-- 180e m -c160-5 40E1 2 0e 100

806 0 -40 -

Aerodynamic heat f low Radiat ive heat f low(

----------

L A d i a b a t i cs u r f a c e

Figure 17. - One -dimensional model ofan aluminum -honeycomb panel.

G r G u nd Test ingOne -dimensional SLA honeycomb-panel thermal tests. - Thermal tests wereperfor med on 12 - by 1 2 -inch, 1 . 7 -inch -thick SLA honeycomb panels to verify the

analytical techniques that wer e used topredic t temper ature s. Radiant-heatinglamps wer e used to simulate the boost-phase heating history. Lamp voltages wer eprogramed s o that the bare-s urface (with-out cork) honeycomb oute r -facesheet testtemper atures matched the analytically pre -dicted temper atu re his tory for the AS-503and subsequent-spacecraft design trajec-tory. The bare-surface SLA honeycomb-panel thermal res pons e for the AS1503 sim -ulation is shown in figure 18. The inner -honeycomb-surface analyt ical predic tions thatresult ed from using a one-dimensional the rmal model were slightly higher than themeasured tempera tures . Cork-covered honeycomb panels wer e subjected to the AS-503radiant-heating history to as se ss the cork performance and to determine the accuracyof the analytical predictions. The predicted and measured thermal response fo r a panelcovered with 0.030 -inch-thick cork and subjected to the AS-503 radiant-heating historya r e shown in figure 19. The analytical re su lts that we re obtained by usingthermophysical-property data for virgin cork ag reed well with the test data. The pre -dicted and measured resp onses for a panel covered with 0.050-inch-thick cork and sub-jec ted to the AS-503 heating histo ry a r e shown in figure 20.the adequacy of the analytical t empera ture -prediction techniques fo r both bare-s urfaceand cork -covered honeycomb panels in typical Apollo launch environments .

The tests.demonstrated

- nalysis0 Test data r u t e r f a c e s h e e t

0P o

0 20 40 60 80 100 120 140 160 180 200 220Elapsed t ime , sec

Figure 18. - Therma l respons e of abare-surface SLA honeycombpanel to an AS-503 t ra jectorysimulation.

240- n a l y s i s0 Test data r O u t e r f a c e s h e e t

20E-zL-20 40 60 80 100 120 140 160 180 200 220Elapsed t ime, se c

Figure 19. - Thermal response of acork-covered 0.030-inch-thickSLA honeycomb panel to anAS-503 trajectory simulation.

I14


17/23

- nalys is0 Test data

240 r o u t e r acesheet

5 120+ loo I n ne r f aces hee t80t c-kwj

ab ; 4b Qo io lbo 1 140 1 IS0 2& 2kElapsed time. secFigure 20 . - Ther mal response of acork-covered 0.050-inch-thickSLA honeycomb panel t o anAS-503 rajectory simulation.

Two -dimensional SLA-joint t her maltests. - Nine test samples were cut from astructura l-test SLA at the locations shownin figure 21.perform radiant-heating tes ts to ass es s thetemperature gradients in the SLA at the endof f irst -stage boost and to verify the ade-quacy of the analytical-prediction tech-niques. Each panel was subjected to theAS-503 radiant -heating profile that wasused in the one -dimensional radiant -heatingtests; however, only the data that wereobtained from test joint 3 will be discussedin this repo rt. Cross-sectional and topviews of the test joint are shown in fig-ur es 22 and 23; the thermocouple locationsalso are shown in f igure 23.

-The samples were used to

The joint was configured with the AS-501 cork pattern (a 0 . 1 inch-thick layerover only the pyrotechnic joint) and subjected to the heating hi story of the AS-503 designtrajectory. The thermocouple tempera ture s for the joint after an e lapsed time of150 seconds (the approximate termination of fi rst- stag e boost) a r e shown in figure 24.The two-dimensional thermal-mathematical model (fig. 25) was used to pred ict thejoint tempe ra tur es for the s ame heating profile. The analytical predictions also areshown in fi gure 24.plexity of the joint and the uncertainties i n the thermal-mathematical model (thermo-physical propertie s, contact resistances, and so forth).

.

The correlation is ver y good, i f consideration is given to the com-

Figure 21. - Spacecraft/lunar module adapter test-joint locations.15


18/23

Aluminum- Aluminum-honeycomb honeycombcore ( l k - i n . core ( W i n .

Figure 22, - Cro ss section of test joint 3 (longitudinal-separation joint .betweenstations X = 710 in. and X = 750 in.).A A

ate: Thermocouples 1 o 24are located on the outerfaces heet. Thermo couple s25 to 46 are located on thein ne r facesheet.

0 0 0 0 . 01 2 3 4 5 625 26 27 28 29 M

-

07310847

-

/-Bolt heads-

00

0932

0

0

0

011350103.1234

0153 70133601438

00

01639

00

03 in .

01748.1840

0 . 0 . . .19 x) 21 22 23 2441 42 43 44 45 46

Figure 23. - Top view of te st joint 3 (longitudinal-separation joint between Sta-tions X = 710 in. and X = in.) with thermocouple locations.A A16


19/23

340320300-2802*24022 0

," 200-180-

B 140-5 120-loo8060-40-20 -

c

I - n a l y s i s0 O u t e r - f a c e s h e e t d a t aI 0 I n e r - f a c e s h e e t d a t a

---b-1---

D i s t a n c e f r o m c e n t e r of j o i n t , in .Figure 24. - Temper ature distribution for test joint 3 covered with an AS-501 cork

pattern and subjected to an AS-503 rajec tory simulation.

r 0 . l - i n - t h i c k cork

Inner facesheetNote A ll dimensions are in inches

Figure 25. - Thermal-mathematical-model nodes for test joint 3 covered with an AS-501cork patter n and subjected to an AS-503 trajectory simulation; the diagram is not toscale.17


20/23

A 0.030 -inch -thick cork layer w a sbonded to the remaining ba re sur face s ofthe panel, which corresponded to a pro-posed AS-503 SLA co rk configuration. Therad ian t -heating simulation w a s again per -form ed on the joint to as se s s the effect ofthe added 0.030-inch-thick cork on thethe rmal gradient in the joint. The testresults for the configuration are shown infigu re 26. The addition of the 0.030-inch-thick cork layer reduced the gradient inthe oute r facesheet from 140" to 90" F.The results of the two-dimensional SLA-joint thermal tests were used to define thetemperatures and temperature gradientsin the SLA struc ture at the termination offirst-s tage boost. The data we re used asinputs for analyses of stru ctur al and ther -mal str es s, which were performed toassess the st ruc tur al adequacy of the SLAfo r th e AS-503 and subsequent spacecraft.

0.1 in: t h c k 0.03- in- thickco rk- C en t! rl ine r c o r k340320

20

..I 0 Outer-facesheet dataI 0 inner-facesheet datao l k o : I ; 1 I 1 I6 ; 1 ;: d 'I ; ;0;1Distance from center of joint, i n .

Figure 26. - Temper ature distributionfor test joint 3 covered with anAS-503 cork patt ern and subjectedto an AS-503 tra jec tory simulation.

F l i g h t D a ta f o r t h e SM a n d SLA a s Co m p a re dw i t h A n a ly t ic a l P r e d i c t i o n sSpacecraft AS-202. - The temperature history for SM sens or SA7916T is shownin figure 2 ' l . The analytical predictions , which were based on the actual launch tra ie c-

7aximum predicted response

260rI Minimum predicted responseL L

40 80 120 160 200 240 280 3MElapsed time, sec

Figure 27. - Temperature history f orSM sensor SA7916T (outer sur-face, X - 350 in. , and a circum-ferenti al position of 253") duringthe launch of spacecraft AS-202.

S -

tory, a r e slightly conservative.mum predicted respon se is based on fullsolar exposure on the panel and radiationcooling to the e ar th; the minimum predictedresponse is based on no so lar exposureor radiation cooling to space. The maxi-mum and minimum predictions shouldbracke t the expected response. The tem-perature history and predicted responsesfor SLA sensor AA7931T are shown in fig-ure 28. The spike in the measu red data atan elapsed time of 145 seconds was causedby the heating fro m the forward-fir ingret roro cke ts of the first stage at separa-tion; this heating was not accounted for inthe analysis.analytical predictions for SLA inner -surfacesens or AA7932T are shown in figure 29. Asobserved in the radiant -heating te st s, theanalytical predictions were higher than theactual te mpe ratu res encountered on the SLA

The maxi-

The temperature history and

18


21/23

Min imum predicted r20090 Measured-data band axim um predicted response160250 urn predicted respons

L L

LL 210 g 120e c

+ 130 Minimum predicted response 40

2c

5 170 80 Measured-data band0)L a-5

Note Spike i n data caused byforward-fir ing retrorocketsat separation

0 40 80 120 160 Mo 240 280 320Elapsed tim e. sec

90 heating from fi rst -stage

I I I I I I I I0 40 80 120 160 200 240 280 320501 Elapsed time, sec Figure 29. - Temperature history forSLA sensor AA7932T (inner su r-face, X = 775 in., and a circum-ferent ial position of 124 ") duringthe launch of spacecraft AS-202 .

Aigure 2 8 . - Temperature history forSLA sensor AA7931 T (outer su r -face, XA = 77 5 in. , and a circum-ferenti al position of 124 ") duringthe launch of spacecraft AS-202.

inner surf ace , The higher predictions Maximum predictedprobably were caused by the adiabatic-inner -surface assumption that w a s madefor all the analytical predictions.

Spacecraft AS-501. - The majorityof the S M surface w a s covered with cor kfo r the Apollo 4 mission, and only the innersurface of the SM w a s instrumented withthermocouples. A s a result, the maximummeasured temperature for the AS-501 SMduring boost was 9 0 " F. The temperaturehistory and analytical predictions for SLAsensor AA7864T a r e shown in figure 30.launch phase, the correla tion is good. Areview of atmosphe ric balloon data from theNASA John F. Kennedy Space Center indi-cated that the atmospheric tempera ture upto approximately 50 000 feet (0 to 80 sec-onds) w a s considerably colder than w a sindicated by the 1962 standard atmosphere(ref. 8), which w a s used in the analysis.The unusually cold atmos phe ric conditionsresulted in an initial cooling after lift-off.

Measured-data band

Except for the first 80 seconds of the Elapsed l ime. secFigure 30 . - Temperature history fo rSLA sensor AA7864T (outer sur-face, x = 730 in., and a circum-

fe renti al position of 174") duringthe launch of spacecraft AS-501.A

19


22/23

C O N C L U D I N G REMARKSThe analytical capabilities fo r predicting spacec raft heating char act eri stic s andstr uctu ral temp erature s were adequate for prediction of the Apollo sp acecr aft temper -at ur es during boost. Given the pr es su re distribution over the surface of the vehicle,the analytical techniques provided reasonably accu rate predict ions of the heating rates

and thermal respo nses of the space cra ft during boost.having local pre ss ure disturbances, the utilized techniques had dubious value. Pr es su redisturbances, caused by externa l hinges, umbilicals, RCS modules, et ce ter a, res ultin localized a reas of higher heating. For the SM and SLA, the problem of protuberance-affected area s w a s solved by applying an experimentally determined heating-rate fac torto the smooth-body heating rates.cedure were overly conservative f or the Apollo -type protuberances.

For areas of the spacecraft

The analytical predictions resulting fro m this pr o-

In the future, to reduce the the rmal-pro tection weight penalty as soc iated withovercons ervatism, attention should be given to obtaining bette r techniques for predict-ing the effect of protuberance heating on spacecraf t. The use of protuberance fact or sthat are functions of Mach number should elimina te much of the overconservatism.Although the techniques that were used fo r design of the boost ther mal protect ion s ys-tem for the Apollo spacec raf t were conservative (because no structur al element o rcomponent exceeded the design tempera ture ), the Apollo 6 (AS-502) anomaly led toinc reased therma l protection of the SLA.stru ctur al margins and to reduce therma l stresses below the originally defined requi re -ments. The analytical-prediction and testing techniques that were developed fo r theApollo Program should enable the the rmal analyst to design future space craft with lessse ve re thermal gradients in the st ruc tur e during boost and with lighte r weight boostther mal protection syst ems.

This addition was made to increase the

Manned Spacecraft CenterNational Aeronautics and Space AdministrationHouston, Texas, October 26, 1972

914-50-20-12 -72

20


23/23

REFERENCES1. Anon. : Thermal Data Apollo Block II Spacecraft. Vol. I, Structural ThermalResponse, SD 67-1107, secs. I to IV, Space Div., N. Am. Rockwell Corp .,

Nov. 20, 1968.2. Eckert, E. R. G. : Engineering Relations for Friction and Heat Transfer to Sur-faces in High Velocity Flow. J. Aeron. Sci., vol, 22, no. 8, Aug. 1955,pp. 585-587.3. Blasius, H. : Grenzschichten in Flussigkeiten mi t Kleiner Reibung. Z. Math. U.Phys. vol. 56, 1908, pp. 1-37. (Available as NACA TM 1256.)4. F. Schutz-Grunow: A New Resistance Law for Smooth Plates. Luftfahrt Forsch.,vol. 17, 1940, pp. 239-246. (Available as NACA TM 986.5. Anon. : Thermal Data Apollo Block II Spacecraft. Vol. 11, Aerodynamic HeatingEnvironment, SD 67-1107, Space Div., N. Am. Rockwell Corp., July 1, 1968.6. Burbank, Paige B. ;Newlander, Robert A. ; and Collins, Ida K. : Heat-Transfer

and Pre ss ur e Measurement on a Flat -Plate Surface and Heat -Transfe r Meas -ure men ts on Attacked Protuberances in a Supersonic Turbulent Boundary Layerat Mach Numbers of 2.65, 3.51, and 4.44. NASA TN D-1372, 1962.7. Anon. : Evaluation of the Ther mal Propertie s of Materials. Vol. 11, Data HandbookFinal Report, AVSSD 0197-66-RR,Space Systems Div., Avco Corp. , June 28,1966.8. Anon. : U. S. Standard Atmosphere, 1962. U. S. Govt. Printing Office.

Apollo Experience Report Spacecraft Heating Environment and Thermal Protection for Launch Through...

Documents

Transcript of Apollo Experience Report Spacecraft Heating Environment and Thermal Protection for Launch Through...