conference on langley research related to the apollo mission

8/6/2019 Conference on Langley Research Related to the Apollo Mission

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LANGLEY RESEARCHRELATED TO APOLLO MISSION

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LANGLEY RESEARCH CENTERJUNE 22-24,1965

NATION AL AERONAUTICS AND SPACE ADM INIST RA TION .


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CONFERENCE O NLANGLEY RESEARCH

RELATED TO APOLLO MISSION ILANGLEY RESEARCH CENTER

JUNE 22-24,1965

I I

Scientific and Technical lnf ormation Divis ion 1 9 6 5NATIO NAL AERONAUTICS A N D SPACE ADMINISTRATIONWashington, D.C.


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CONTENTS

iii

SPACECRAFT TECHNOLOGY

Heating and Thermal Protec tion6 G

1. PROSECT FIRE STAGNATION HEATING MEASUREMENTS . . . . . . . . .By Dona L. Cauchon and Richard C. Dingeldeinl 4 G q J. RECENT EXPERLMENTAL STUDIES ON HEAT TRANSFER TOAPOLLO COMMAND MODULF, . . . . . . . . . . . . . . . . . . .By Robert A. Jones and James L. H u n t

3 . AERODYNAMIC HEATING CHARACTERISTICS OF APOLLOLAUNCH CONFIGURATION . . . . . . . . . . . . . . . . . . .By Robert L. Sta l l ings , Jr., and E a r l A. Price, Jr.4. l?GIGHT AND GROUND TESTS OF APOLLO HEAT-SHIELII MATERIALBy W i l l i a m A. Brooks, Jr., Stephen S. Tompkins,and Robert T. Swann

Command-Module Landing Dynamics. . . . . .SUIL5 . DYNAMICS OF DROGUE PARACHUTE PHASE OF APOLLO RECOVERYBy Sanger M. Burk, Jr. I / 8I / ' , 7 " < .6. LANDING IMPACT STUDIES OF APOLLO COMMAND MOWLE. . . . . . . . .' - I **clB y Sandy M. Stubbs

Lunar Surface and Landing Dynamicsk . r 41 - 1 5 v. PENETROMEIXR RESEARCH AND DEVEZOFNENT FORLUNARSURFACEEVAMATION . . . . . . . . . . . . . . . . . . . . . .By Alfred G. Beswick and John Locke McCarty i--. . . . . . . .. SURVEYOR SPACECRAFT TOUCHDOWN-DYNAMICS EXPERIMENT . .$a -=.B y Sidney A . Batterson

1

9

1929

4553

6169

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IL,. ' ?ID9. LANDING ST AB ILI TY FOR LUNAR-LANDING VE HIC LES . kL''-:[g . . 77\" B y W i l l i a m C . Walton, Jr. , R o b e r t W. H e r r , 4

and H. Wayne LeonardTESTS AT SIMULATED LUNAR GRAVITY . . . . . . . . . . . . . . 8710. TFCBNIQUE FOR CONDUCTING FULL-SC ALE LANDING-IMPACT ,q~By U l y s s e J . B l a n c h a r d

12,Ppacecraft Simulation

11. DISCUSSION O F REKDEZVOUS TECBNIQUES . - . p). . . . . 97!'>ry G e n e C . Moen

12. DYNAMIC SIMULATION OF L9M DOCKING WITH D O L L 0COMMAND MODuIlE I N LUNAR ORBIT . . . . . . . . . . I . . . . . . 07119

.k . . 125By H o w a r d G. H a t c h , Jr . , and Jack E. Pennington

$ P13. SIMPLIFIED MANUAL GUIDANCE TECHNIQUE FOR -1 1LUNAR ORBIT ESTABLISHMENT . . . . . . . . . . . . . y . . . . . . ./e A, !$- -11 1y G e n e W. S p a r r o w and G. K i m b a l l Miller, Jr.

1 1

,m. -- ;, 814. MANUAL CONTROL OF APOLLO HIGH-ALTITUDE ABORT . . . . i . . .! .By A l f r e d J . Meintel, Jr. , and K e n n e t h R . G a r r e n

15. PRELIMINARY TESTS WITFI THE L4NGIXY LUNAR LANDING RESEARCH FACILITYB y D o n a l d E. H e w e s

LAUNCH-VEHICLE TECHNOLOGY u\.I,$416. STATUS OF APOLLO-SATURN v DYNMC-MODELS PROGRAM . . . . . ['I. . 143B y Sumner A. L e a d b e t t e r d 4:

17 . COMPARISON OFFULL-SCALE AND MODEL BUFFET RESPONSEOFAPOLLO BOILERPLATE SERVlCE MODULE . . -By R o b e r t V. D o g g e t t , Jr.18. WIND-TUNNEL 7NVESTIGATIONS O F EFFECTS OF GROUND

By G e o r g e W. Jones, Jr . , and Moses G . F a r m e rON SATURN-APOLLO LAUNCHVEHICLES. . - - - .

192,"EFFECT OF ASTRONAUT CONTROL ON LAUNCH-VEHICI;E WINDB y R o b e r t K. Sleeper20.- SOME INITIAL RESULTS

! B y P h i l i p M. Edge,

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OPERATIONALTECHNOLOGYi G ( i 6 P PI. . . ./21. RESEARCH APPROACH TO RFENTRY COMMUNICATIONS BLACKOUTBy Paul W. Huber2 2 d FLIGHIl RESULTS OF GENTNI REENTRY WMMUIlICATICNS ExpERlMENT .By Lyle C. Schroeder

FLUID MXCHANICS ASPECTS OF THE GESIINI REENTRY ,I c j G d f ' I- -J,

I S F S23. C-CATIONS EXPERIMEXT . . . . . . . . . . . . . . . . e-... .

and Ja r re t t K. H u f f l n a nBy Ivan E. Beckwith, Dennis M. Bushnell,

24. ANTENNAS UNDER A B U T I O N MATERIALS . . . . . . . . . . . . . . . . . . .- *-. -2By W i l l i a m F. Croswell

APOLLORELATEDTECHNOLOGY25. mAR0RBm.. . . . . . . . . . . . . . . . . ,r a b . . . p] .11y Edmund A. Brummer26. APPLICATION TO APOLLO OF SOME ORBIT DETlTERMINATION Q -P 7 ti,"KFSULTSOFTEELLTNARORBITEIR . . . . . . . . . . . . . . . . . . .By Alto n P. Mayo

F :27. EXPLORER sATEL;LITE MEASUREMENTS 03' ME!I'EOROID { Q 38 Q-2PENETRATION RATES IN STRUCTURAL MATERIALS . . 4 4By Robert L. O'Neal

, zr7 I - \- " I i8. UTILIZATION OF APOLLO SPACECRAFT SYSTEMS FOR SUPPORTOF A MANNED ORBITAL RESEARCH LABORATORY . . - 2 *>!.By W i l l i a m C. Hayes, Jr." ' V D p].9. DYNAMTCS AND CONTROL ?33SEARCH APPLICABIX TOAPOLLO MTENSION SYSTEMS . . . . . . . . . . . . . . . . .By Peter R. Kurzhals, Claude R. Keckler, and W i l l i a m M. Pilandq q +)30. CAziBON DIOXIDE CONTROL FOR MANNED SPACECRAFT . . . . . . . . .

By Rex B. Martin31. WATER AkD WASTE MANAGEMENT SYSTEMS ? *32. CONTAMINAlvT COUECTION AND IDENTIFICATION . . . . . . . . . . . e - > 7

. . . . . . . . . .By Vernon G. Collins and Robert W. Johnson ~ -, [ , "3 r EBy Robert M. Bethea, I r i s C . Anderson, and Robert A . Bruce -,4

189205

217

239

267

275

283

293

303

313323

333

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-3. CONTAMINANTS FROM MANNED SPACECRAFT S m T I O N S . . . . . . .By E. Eugene Mason and C h a r l e s H. Wilson34. ~ F U l T E DEGENERATIVE LIFE SIPPORT SYSTEM FOR \

EXTENDEDMISSIONDURATIONS .B y War ren D. Hy-pes, Robert A. B r u c e , and Franklin W. B o o t h35. WNAR STAY TlME EXTENSION M O W .B y Charles I. Tynan, Jr.36. WATER-IMMEEISION TECHNIQUE FOR S m T I O N OF INGRESS-EGRESS,pqMANEWERS UNDER CONDITIONS OF WEIGaTLFSSNESS . . . bL'!. .B y Otto F. Trout, Jr.37. SPACETEZIEBING. . . . *6'?51 . . 391

\ I +< \ 1, i.y Gary P . B e a s l e y and R oy F. B r i s s e n d e n \

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1. PROJECT FIRE STAGNATION HEATING MEASUREMENTSBy Dona L. Cauchon and Richard C. Dingeldein

The results of the f i r s t f l i g h t of Pro jec t F i re (F l igh t 1) ind ica te t h a t astagnation t o t a l heating r at e approaching 1400 watts/cm2 was measured near peakheating. This value was of the order of 400 watts/cm2 higher than that deter-mined by adding th e r ad ia tio n above 0 . 2 3 measured onboard th e F i r e vehicle(corre cted fo r calorimeter absorptance) t o a representative calculated convec-t i ve hea ting rate.w a s e i th e r d i rec t ly o r ind i rec t ly respons ib le fo r t h i s addi t ional flux. It i sshown t h a t th e measured radia t iv e and t o t a l heat ing rates can be predicted withgood accuracy by using c er ta in ava ilable an al yt ic al methods..

The evidence i s strong t h a t vacuum-ultraviolet radiation

INTRODUCTION

Pro jec t F i re i s a fl ig h t experiment designed t o measure re en tr y hea ting a tTwo flights have been made along the Eastern Test Rangeyperbolic velocit ies.wi th reentry in th e v i c i n i t y of Ascension Island.Apri l 14, 1964, and the second, on May 22, 1963.about preliminary.results fram Flight 2, th e present paper i s concerned onlywith th e results of Flight 1.

The f i r s t f l i g h t was made on:Except fo r a concluding note

One af th e primary objec tives of Wojec t F i re i s t o d ef in e the t o t a lheating environment associated with the reentry of a large-scale Apollo-shapedvehicle a t a ve loc i ty of 37,000 f e e t p e r second. In this s tudy to ta l hea t ingi s defined as th e he ating t o t he body surface from t h e t w o principal modes oft r ans fe r ; namely, the convective heating and the radiation that i s absorbed bythe surface.forebody of t he vehicle .T hi s t o t a l h e at in g w a s determined by using calorimeters on the

In order t o understand more fu ll y th e mechanics of h eating f o r t he twoseparate modes, direct measurements of t h e gas radiance were also obtained byusing onboard radiometers.the measured radiative heating rates i s at t r ibu ted t o convective heat ing and t orad ia t ion absorbed by th e calorimeters, bu t which i s outsid e th e measurablewavelength range of the radiometers.

The difference between the t o t a l heating rates and

The results of Projec t Fir e offer f o r t he f i r s t time i n t h is severe envi-ronment t h e op portu nity t o compare actua l f l i g h t measurements on a large-scalebody with theoretical estimates.This paper will b ri e f ly tr e a t the reentry-package configuration and t h eFire trajectory with particular emphasis on the reentry portion.

a discuss ion of th e t o t a l and radiat ive heat ing rates deduced from t h e Fl igh t 1Then follows


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reentyy. Finally, calcula t ions w i l l be presented showing th e ext ent t o whichth e F i r e measurements are predic table by cer ta in avai lab le a nal yt ic al methods.

SYMBOLSq,: * .aw.mih&q*

Q stagnation heating ra te, watts/ cm2absorptance of beryUium calorimeters%e

h wavelength, microns (v )7 transmission of hot a i r i n wavelength range from 0.05 t a 0.21-1,defined as %b,with abs orp tio n/q W, transparent gasSubscripts :cow , convectiveGE ! General Electric radiation theoryRAD radia t ive

DISCUSS1 N AND RESULTSFigure 1 hows the P roject Fi re ree ntry package th at was acce lera ted t o an

The reentry package w as approximately 2 fee t i n diam-Because no single calorimeter could surviveentry veloci ty of about 38,000 fe et per second using an Atlas booster and anAntares I1 rocket motor.eter and weighed about 185 pounds.th e en ti re reentry, the experiment was broken down i n to th re e d is cr et e data-gather ing periods. Three separa te beryll ium layers ( ind ica ted in b lack i nf i g . 1) served as the calorimeters during these periods.mented ra di al ly and i n depth with thermocouples.lay ers , sandwiched between th e calorim eters, provided the he at pro tec tion afterthe preceding calorimeters had melted.ejected a t prescribed t i m e s during the reen try i n order t o conduct, i n a cleanenvironment, the experiments near peak-heating and during th e terminal heat ingportion of the reentry.makerial protected the main body of the reentry package.

Each layer w a s i n s t r u -Two phenolic-asbestos ablationThese ab la t ion layers w e r e a c t i v e ly

A final ab la t ion layer and appropriate i n s u l a t i n gThree to ta l radiometers were used t o measure the gas radiance - one a t t h estagnation region, one viewing a t an offse t locat ion on t h e fro nt face, and onelouking out f r m the afteYbody.fused-qwtz windows. The radiometers viewed the rad ia t i on thraughIt should be noted t h a t th e reentry-package shape during the second experi-ment per iod i s esse nt ia l ly t h a t d he Apolla command m o d u l e .

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Although Pro ject F i r e provided extensive hea ting data over the forebodyand afterbody, s p e c t r d coverage of the gas radiance, afterbody pre ssu re meas-urements, and data on radio-comrmmications at te nu at iu n or "blackout," t h i spaper i s confined t o a discussion of the heating rates i n th e s tagnation region.

The Fl i ght l h i s t o r y s tar ted with launch from Cape Kennedy, Flor ida(flg. 2) Following A t l a s stagin g and shroud separation, th e space craft, whichconsis ted of t he v elo ci ty package with i t s guidance shell, was separated fromth e A t i a s susta iner md u p i t c h maieu-ver 'wu&s zffeeted to place it ir: the p q e za t t i t u d e f o r l a te r f i r i n g of th e Antares I1 rocket motor downrange. After acoast of some 20 minutes, a t an alt i tude of about 1,000,000 feet, the space-c ra f t was spun up, th e velocity-package s h e l l separated, and the Antares I1 wasignited.reentry package was separated from the spent motor case a t about 470,000 feet .The reentry portion of t h e t r a j e c to r y i s defined as beginning a t 400,000 feet,some 7 seconds af ter separation.l i e s below 400,OOO fee t .

The rocket motor burned out a t approximately 7OO,OOO feet and the

The p a r t of th e f l ig h t of resea rch in te re s tThe Fire reentry trajectory beginning a t 400,000 feet i s shown i n figure 3.

The open-bar segments a t th e bottom ofVelocity and a l t i t u d e are plat ted agains t reentry t i m e from 400,000 feet .entry angle a t k00,OOO feet was -14.50.t h e figure ind ica te t he periods where the calorimeters obtained data; eachtotal. heating experiment was considered termi nated when t h e fr o n t surface ofthe calorimeter melted. The shorter darkened portions define the periods ofacceptable transmission char ac te ris tic s of t h e op ti ca l windows. Note th a t th e. F i r e vel oc it y rernained above t h e i n i t i a l Apollo en try vel oci ty (which i s about36,000 feet per second) through the second data period and was superorb i ta lthroughout almost t he en ti re experiment.i n figure 4.spread i n t he results re fl ec ts th e accuracy t o which t he temperature-time his -tories from the calorimeters could be reasonably determined.w a s due t o th e convective plus the absorbed radiative heating a t the beryll iumcalorimeter surface.heating.

The

Stagnation heating rates are plot ted agains t reentry time from 400,000 feetThe Project Fir e re sul t s are shown as heavily bordered bands; th eThe Fire heat flux

A value approaching 1400watts/cm2 i s indicated near peakWere i t poss ible t o measure a ll the gas rad iati on by the radiometers, th econvective and radia tive heating ra tes could eas il y be assessed separately bymerely sub tra ctin g th e absorbed radia tive component from th e t o t a l heat ing rate.

This i s not the case, however. Radiation i s sp ec tr al ly dependent and t h e wave-length cutoff d the quar tz windows essentially dictates limits on t h e amountof incident gas radiation a t the body surface t h a t i s actually sensed by theradiometers.a rep res ent ativ e convective heating ra te based on th e F i r e t r a j e c to r y wa s cal-cul ated by usi ng t'ne theor y of reference l. I n t h e absence of pe rtu rbi ng phe-nomena such as coupling effects with radiation, t h i s estimate of th e convectiveheat ing i s reasonable. The differe nce between t h i s estimate and t h e t o t a lheating results should represent a measure of the absorbed radiation over t h ecomplete spectrum. The discont inui t ies exhibi ted in th e ca lcu la t ions re f lec tth e effec t of considering changing body shape due t o e jec t ing the hea t sh ie lds .

I n order now t o dist ing uis h between t h e va r i m s modes of heating,

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The wavelength influence on the rad ia t ion data i s indicated in figure 5.The wavelength scale i s indicated logarithmically fram 0.05 t o lOp. The theoryind ica tes that the energy from h o t - a i r r a d i at i o n i s dis tr i bu ted mainly over th espectrum from 0.05 t o 2 p and t h a t beyond these bounds it i s negligible. Thequartz windows essentially l i m i t the radiometers to the range from 0.23 t o 4.wwhich i s indicated i n th e figure by t h e dark ba r segment. Therefore, t h e radi-ometers amply cover all t h e g as rad ia t ion a t t h e higher wavelengths.a t t h e low wavelengths, i n th e region denoted as vacuum ultraviolet below 0.2%,subs tan t ia l ly h igh l eve l s of radiation can be estimated.not be recorded by th e radiometers, but would be sensed by the calorimeterswhich absorb almost like a blackbody a t the short wavelengths.+acuun ul tra vio le t , has i n general been used to define the radiation belowabout 0.14.all ul t r avio le t radia t io n below 0.2 3.

reen t ry time frm 400,000 f e e t . The radiative heating i s indicated on a lo gscale.t o 4.w are indicated by the heavy li n e s fo r t h e t h r e e data periods.mum heating rate obtained by Fire was about 180 watts/cm2 i n the second dataperiod, pr io r t o the time of expected peak radi atio n.

However,This radiation would

The term,In th is repo rt , t h at descr iption has been extended to describe

Figure 6 i s a plot of the s tagnat ion rad ia t i ve heat ing rate againstThe experimental F ir e he ating rakes f o r t he wavelength range from 0.23The maxi-

In f igure 6 i s also indicated the range of values re su l t i ng f ram calcula-tions over comparable wavelength ranges made by using some of the more prominenttheor ies ( see re f s . 2, 3, and 4) and the F ir e reentry t radectory . I n the secondand th i r d data periods , the Fire r es ul ts tend toward th e lower bound of theo-r e t i c a l p r e di ct io n . The theory surrounding the f i r s t data period i s consider-ably more complex and le s s understood than f o r t he oth er perio ds. A t t h ebeginning of reentry nonequilibrium radiation and radiation-limiting phenomenacontribute strongly t o th e shapes of th e the or et ic al curves. It i s beyond thescope of t h i s prese ntation t o discuss t he assumptions used t o account fo r suche f f e c t s as truncation and coll ision l imiting; however, the se e f fe c ts have beenconsidered, and a t these l o w i n i t i a l values, t h e F i r e r e s u l t s a r e pr e di c ta b le t owithin a f e w watts/cm2.

I n an attempt t o assess th e poss ible radi a t io n below 0.23, one addit ionaltheory has been added i n figure 7.Fire t ra jectory denoted by this theory inc lud es an estim ate of t h e vacuum ultra-viole t contr ibut ion; this theory covers the spectnmr. of hot -a i r r ad ia t ion from0.05 t o l@. By noting again that the heat ing rate i s p lo t t e d on 8 logarithmicscale, the addi t ional radia t ion in the vacuum ultraviolet region can indeed beappreciated. This estimate of r a d i a t i o n i s based on the assumption of a trans-parent gas, that is , there i s no absorption in the shock layer. This assup-t i o n i s q u i t e v a l i d aver the conventional wavelength range covered by th e radi-ometers; however, i n th e u lt ra vi ol et por tion, this is no t the case. Much ofthe ene rw ind ica ted by th i s theory wi l l be absorbed within th e gas befare itreaches the body.t i o n of energy does t o t h e v a r io u s modes of heat transfer are not completelyunderstood at t h i s time.

The ca lcula ted rad ia t ive heat ing f o r the

The degree t o which it w i l l be absorbed and what t h i s absorp-

I f th e measured rad ia tio n from F ir e (i. ., that above 0.2%) , corrected fo rthe absorpt iv i ty of t h e beryllium calorimeter, i s added t o t he calc ula ted Cohen4


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convective heating (see f i g . k ) , a ra ther s m a l l mount of energy i s added com-pared wlth what is Eeeded t o account fo r the t o t a l heat ing leve ls measured onthe calorimeters. The evidence i s strong that the por t ion of energy which isnot accounted f o r i s due t o the vacuum-ultraviolet rad iati on, adjus ted for gasabsorption.t o th e body surface. On th e othe? hand, t he por tio n of the rad ia t ion that i sabsorbed i n t h e shock and boundary layers may w e l l reappear a t the surface inthe form of enhanced convective heating.

This addit ional radiation i s probably transmitted in par t d i re ct l y

What i s really s ign i f ican t , i s not so much the mode by which the heat i st r ansmit ted t o th e w a l l , b ut t h e f a c t t h a t , i n t h e r eg io n of peak heatin g, anaddi t io nal heat f lux of t he order of 4OO watts/cm2, roughly twice t he ra dia tionmeasured onboard i n th e conventional wavelength range, con trib ute d t o th e F i rereentry heating, and that vacuum-ultraviolet radi atio n appears t o be th e o rig i-n a l energy source.Figure 8 indica tes th e f a i r l y good accuracy with which the t o t a l calorim-e t e r h e a t in g rates f o r F l i g h t 1 can be predicted on the basis of' using avai l -able methods. (1)Cohen convec-t ive hea t ingf o r th e absorption of t he beryllium calorimeters and absorption within t he gas(designated by 7 ) .

The.calculated heating curve i s t h e sum of ;( ref . l),and (2)GE radiation theory ( ref . 2) spectra l ly adjus ted

It can be seen t h a t these methods ar e no t i n complete agreement with t h eProject Fi re resul ts , but cer t a inly appear t o be representa t ive of the hea tenergy a vail abl e t o th e body regardl ess of th e mode of t r ans fe r .which bet ter account can be made of the differe nces between t he the orie s and theF i r e results are ce rt ai nl y des ira ble, and may in some cases already be avail,-For instance, i t appears that the deviations i n the f i r s t data period mayw e l l be a t t r ib ut ed t o increased convective heat ing due t o vor t ic i ty in terac t ionsbetween t h e shock and boundary layer. The prevailing Reynolds numbers duringthe earay reentry are very low (less than 30,000), and a l l radiation, includingthe vacuum ultraviolet, i s a t re la t ive ly low l eve l s . I n th e second and th ir d

data periods, there e x i s t s a need f o r continued work, both t he or et ic al andexperimental , i n th e area of vacuum-ultraviolet ra di at io n, abs orp tio n phenomena,and interchan ge mechanisms between ra di at iv e and conve ctive he ati ng t o provideany f i n a l answers about some of the differences indicate d.

Methods by

' able.

The f l i g h t records of t h e Flight 2 reen t ry are being analyzed a t the p res -en t time; a preliminary assessment of the ra di ati on a;nd t o t a l heating measure-ments indicates that t h e results (using appropriate vel ocit y scaling) appear t obe consistent with those of Flight 1.

CONCLUDING REsrIAHIcs

It i s evident that Pro jec t F i re has indeed provided valuable anchor-pointdata fo r assessing the hea t ing associated with this reentry environment. Theindicat ion that th e Fi re re entr y can be predicted by cer tai n exi st in g methodsshould provide assurance f o r th e Apollo FYoject t h a t the reentry heating envi-ronment f o r t he command module can also be estimated w i t h good accuracy.

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REFERENCES

1. Cohen, Nathaniel B.: Boundary-Layer Sim ilar So lu tio ns and Co rre la tio nEQuations for Laminar Heat-Transfer Distribution in Equilibrium Air a tVeloci t ies up to 41,000 Feet Per Second. NASA TR R - l l 8 , 1961.

2. Nardone, M. C , ; Breene, R. 0.j Zeldin, S. S.; and Riethof, T. R.: Radlanceecies in High Temperature A i r . Tech. Infonu. Ser. R 6 3 S D 3 (Contract(694)-222), Missile and Space Div., Gen. Ele c. Co., June 1963.(Available from DDC as AD No. 40854. )3. Kivel, B.; and Bailey, K. : Tables of Radiation From High Temperature A i r .Res. Rept. 21 (Contracts AF 04(643)-18 and AF 49(638)-61), AVCO Res. Lab.,Dw. 1957.4. Meyerott, R. E.; Sokoloff, J.; and Nicholls, R. A . : Absorption Coefficientsof A i r . JWCRC-TR-59-296, U . S . A i r Force, Sept. 1959.

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PROJECT FIRE FLIGHT SEQUENCE OF EVE NT S

Figure 1

FLIGHT 1 REENTRY TRAJECTORYVELOCITY, ALTITUDE,40Fr, ~ ITY i40x104T

30 30

FLIGHT

ALTITUDE

EXPERIMENTS

16 24 32REENTRY TIME, SEC

Figure 2

FLIGHT 1 STAGNATION HEATING RATES OBTAINED FROMCALORIMETER MEASUREMENTSTIME'O AT ALTITUDE=400 .000 FT

FIRE DATA

Figure 3 Figure 4

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:TICAL

I

SPECTRAL RANGE OF RADIATION MEASUREMENTS

I J

I BERYLLIUM CALORIMETERS 4-I I I I I I.05 .I I IOWAVELENGTH, p

Figure 5

FLIGHT 1 RADIATIVE HEATING RESULTS AND THEORETICALPREDICTIONSTIME = 0 AT AlTITUDE * 400.000 T2,000-1,000

100;STAGNATION :RADIATIVE .HEATING RATE,W/cmz

IO :

FIRE RESULTS(23 O 4Q4I I I I12 16 20 24 28 32 36 40ELAPSED REENTRY TIME. SEC

Figure 7

Figure 6

ESTIMATE OF RIGHT 1 STAGNATION HEATING RATESTIME * 0 AT ALTITUDE = 400,000 FTCALCULATED TOTAL HEATING

qCoIIV+ A = O . Z l D D p )B0(qRAdtE + k = O . O 5 + 0 2 r j 7B.(qRAD)WrA /

STPGNATITIONHEATING RATE,W / c d iu

1 , / , , , ,'\,\0 16 24 32 40ELAPSED REENTRl TIME, SECFigure 8

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2. RECENT EWER=& STITDIES ON HEN' 'IRANSmRTO A P O U O COMMAND MODUIX

By Robert A. Jones and James L. IIunt

This paper presentseffects of protuberances some recent experimental results on the interferenceand reaction-control jets on the heat transfer to theApollo command module and some results of a basic investigation on the flowfield and heat transfer in the separated region on the afterbody.ments were made in the Langley Mach 8 variable density tunnel which is a con-ventional blowdown facility.increased the heat-transfer rate by factors as large as 2.5. Measurements ofthe separated-layer thickness on the afterbody indicate that this thicknessvaries with Reynolds number and is thicker at the lower Reynolds numbers. Theheat transfer in the separated region was found to be a function of theseparated-layer thickness. The results from several different types of groundfacilities a8 well as some results from flight 1 of Project Fire were comparedand it was found that an upper limit of heat transfer to the separated regioncould be defined.

The experi-It was found that the presence of the shear pads

Although much research has been done on the heat transfer to the Apollocommand module, several areas of uncertainty still exist.describe s a n e recent experimental work related to two such areas. One is theinterference effects of protuberances and reaction-control jets; the other isthe heat transfer in the separated region on the afterbody.made in the Langley Mach 8 variable-density tunnel which is a conventionalblowdown facility equipped with a model-injection mechanism for transienttesting.

This paper w i l l

These studies were

SYMBOLS

PdBthhS

specific heat at constant pressurediameter of face of modeltotal enthalpylocal measured heat-transfer coefficientheat-transfer coefficient at stagnation point

9


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kM

PW

pJ tpt'w, dS

TawTPcTtU

ZtU

6P

P

thermal conductivity of model wallfree-stream Mach number

10ca.l measured pressure.stagnation pressure of reaction-control jetstagnation press3 behind normal shock at free-stream Mach numberfree-stream Reynolds nunber based on body diametersurface distanceadiabatic wall temperaturephase-change temperaturetotal temperaturevelocity at edge of separated boundary layer based on a referencetemperaturecompressibility factor at stagnation conditionsangle of attack; thermal diffusivityseparated-layer thickness measured normal to free-stream flowdirect onviscosity based on reference temperature conditions at edge of sepa-rated boundary layerdensity based on reference temperature condition at edge of separatedboundary layer

PRO!INBERANC!ES AND REACTION CONTROLS

Photographs of the 0.026-scde model showing many of the protuberances areshown in figure 1.difficult to measure by conventional thin-skin calorimeter techniques becauseof the small size of the models which can be tested in hypersonic facilities.The small models make instrumentation with thermocouples difficult and, in addi-tion, since it is not known beforehand which area will be most affected it isdifficult t o determine where thermocouples should be placed.

The effects of these irregularities on the heat transfer are,

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To overcome these difficulties a new experimental technique developed atthe Langley Research Center m s used i n the present study.employs a very thin coating of a material which undergoes a visible phase changefrom an opaque solid to a clear liquid at accurately known temperatures. Themodel, which is made from a dark-colored low-thermal-conductivity plastic, issprayed with just enough of this material to fog its surface. This coating,which has the appearance of tiny opaque white crystals, is less than 0.001 inchthick. The cool coated model is suddenly exposed to the test stream and theprogreseion of the phase-change patterns is recorded by a time-study motion-picture camera. The temperature of the model surface at the location of thephase-change linea is assumed to be the same as the melting temperature of theparticular material used. A photograph of the tunnel test section showing themodel, camera, and stroboscopic flash light used to illuminate the model ispresented in figure 2.

This technique

The isothermal-coated model is placed in the model injection mechanismlocated directly beneath the test section; the t m e l is then started andbrought to the desired test condition; the camera and light are then turned on;and the model is rapidly injected into the test airstream. The useful test timeis usually from 1/2 to 10 seconds.removed, the coating is washed off with a special thinner, and then the model iscooled and repainted for the next test.cient is found by relating the time elapsed, from model exposure until a partic-ular phase-change pattern occurs, to the solution of the transient heat-conduction equation.values of the model thermal properties k and a are known as well as thevalue of the temperature ratio Tpc2mj thus for the time corresponding to anyparticular phase-change pattern, the heat-transfer coefficient can be read fromthis plot.the accuracy obtainable with it i s given in reference 1.

After completion of the test, the model isThe value of the heat-transfer coeffi-

One form of this solution is plotted in figure 3. The

A more complete description of this technique and a discussion of

Prints of three individual frames of the motion-picture film taken duringone test to determine heating rates in the vicinity of the shear pads on theface are given in figure 4.lines at which the phase change is taking place and consequently are lines ofknown constant heat-transfer coefficient.transfer distribution on the face of the model showing the effects of the shearpads and tension ties. The maximum heating rates near the windward pads were1.17 times the stagnation-point value. Measurements made in the same region ona smooth model indicate that the interference effect of the shear pads was toincrease the heating rate by factors as large as 2.5.

The lines separating the light and dark portion areFigure 5 is a map of the heat-

Figure 6 shows photographs of the phase-change patterns near one of theThis reaction-control jet is located inreaction-control jets and a map of the heat-transfer distribution obtained fromsuch patterns is given in figure 7.what is normally the separated afterbody region. In figure 6 he thrust of thejet is outward so as to roll the top of the model away from the observer. Thejet is a small contoured supersonic nozzle exhausting cool dry air. The designof this nozzle and the pressure at which it was operated were such as to matchthe exhaust expansion boundary of the reaction-control motors on the actualcommand module. The maximum heat-transfer coefficient measured in the inter-ference region of the jet was 0.15 of the stagnation-point value which

ll


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corresponds to an increase of about ll.times the heating rate in the same areawith the jet off.during entry the control jets are fired in short bursts and during a portion ofthe trajectory which does not coincide with peak heating.increase in heating due to the control jet does not greatly affect the heat-shield design. Perhaps of more concern is the effect of the control jets onthe af'terbody pressure distribution and the resulting changes in net forcecaused by the jets. The regions of increased pressure should correspondroughly to the regions of increased heating; therefore the present techniquemay be of m e in pressure-distribution studies.

This increase in heating rate is rather lmgej however,Therefore, this

The results discussed here as to the effects of protuberances and reactioncontrols on heat trmsfer are typical of the more complete results presented inreference 2.

SEPARATED AFTEIiBoDY HEAT !TRANSFER

Most of the data obtained in the separated region on the afterbody wereobtained by using sting- or strut-mounted models which, of course, disturb theafterbody flow field and make interpretation of data difficult. In aaition,no adequate theories or correlations for heat transfer in separated afterbodyregions exist at the present time.sure distribution, and heat transfer in the separated afterbody region was made.Therefore a study of the flow field, pres-Measurements of the separated-layer thickness at zero angle of attack areshown in figure 8 or two different model-aupport configurations.used to make these measurements is illustrated in figure 9. A cylinder ofapproximately 1/16-inch diameter was coated with a temperature-sensitive mate-

rial and placed in a hole in the afterbody surface so that it projected normalto the surface.ment of the separated shear layer on the coated cylinder resulted in a clearlydefined phase-change pattern which indicated the position of the shear layer.Only one cylinder was used for each test, but by varying the location of thiscylinder in subsequent tests, the streamlines of the shear layer shown in fig-ure 8 ere determined.thickness was thought to be negligible inasmuch as it had an insignificanteffect on the measured afterbody pressure.

The technique

When the model was exposed to the test airstream, the impinge-

The effect of the cylinder itself on the separated-layer

There are two interesting results indicated by the patterns of figure 8:(1)the separated-layer thickness varies with Reynolds number, the thickerlayer occurring at the lowest Reynolds number, and (2) the separated-layerthickness varies with the sting or strut used to support the model.four different strut configurations were studied.figure 8. The other two struts used consisted of a sting projecting straightback parallel to the center line of the model and one similar to the strut onthe left in figure 8 xcept that it was twice as thick.left of figure 8 ppeared to have the least interference effect on the separatedafterbody f l o w for zero angle of attack.

In all,Two of these are shown inThe strut shown on the

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Measurements of the separated-layer thickness at an angle of attack of 35OThe separated-layer thiclmess is much larger than thatre shown in figure 10.for an angle of attack of Oo and a general trend for thicker separated layersat the lower Reynolds numbers is evident.two highest Reynolds numbers is not understood.

The reversal of this trend at the

Figure 11 shows measured pressure distributions in the separated afterbodyregion forthermocouple-ty-pe gages which were located inside the model and connected tothe orifices with 112-inch lengths of tubing.rapidly and less than 1 second was required for a measurement; thus, the modelremained near room temperature.sure varies with Reynolds number; the higher pressures occur at the lowerReynolds numbers.thickness where the thicker layers occurred at the lower Reynolds number.culations of the afterbody pressure made by assuming that the flow expandsisentropically from stagnation conditions around to the measured separationangle were in close agreement with the measured pressure levels for an angle ofattack of Oo.

a = Oo and 3 These measurements were made with miniaturizedThese gages responded very

For both angles of attack the level of pres-This trend is similar to the variation in separated-layer

C a l -

Several attempts were made to correlate the heat transfer to the separatedOne possible correlation is shown in figure 12 for an angle ofIn this figure the Stanton number based on local flow conditions

afterbody by using the measured separated-layer thickness and pressures dis-cussed earlier.attack of Oo.external to the separated boundary layer is plotted as a function of the ratioof surface distance from the rear of the afterbody to the separated-layer thick-ness measured normal to the free-stream flow direction. The measured heat-transfer coefficients were taken from reference 3 which describes a heat-transferstudy for the afterbody of this same configuration which was a l s o made in thesame facility under similar test conditions.Reynolds numbers in the same manner as for figure 11.of figure 12 i s for a limited range of conditions, it does indicate that theheat-transfer rate in the separated afterbody region is sensitive to theseparated-layer thickness which, in turn, is a function of Reynolds number aswell as of the sting or strut configuration used in the test.

The different symbols denoteAlthough the correlation

In view of the fact that the afterbody separated flow field is affected bythe model support,. extrapolation of ground-facility results to flight conditionsis difficult.ities under different test conditions, an upper limit for heat transfer to thisregion can be defined. Figure 1 3 shows data from several facilities as w e l l asdata from flight 1 f ProJect Fire.local Stanton number based on conditions at the edge of the separated boundarylayer as a function of local Reynolds number based on the same conditions andsurface distance from the forward stagnation point. For most of these data, thelocal conditions were determined by expanding the flow isentropically fram stag-nation conditions to the measured afterbody pressure. In cases where the after-body pressures were not available, correlations were used to determine the pres-sure.handled somewhat differently.body is in chemical nonequilibrium; thus, in order to get the local Stanton

However, by comparing data obtained in several different facil-

These data are presented in terms of a

The set of data points for the Project Fire high-altitude case wereAt these flight conditions the flow around the


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numbers, the composition was assumed to be frozen at the stagnation equilibriumconditions and expanded isentropically to the measured afterbody pressure.It can be seen in figure 13 that a line representing the upper limit of

Such a line would correspond to approxi-hese different data can be drawn.mately 5 percent of the calculated convective heat-transfer rate for the stag-nation point at an angle of attack of 0 and could be used as an upper limit forheat -shield design.A comparison of the afterbody pressure data obtained during flight 1 fProject Fire with the afterbody pressures of the present study is given in fig-ure 14. For the low Reynolds rimer flight data, the band shown represents thescatter in the telemetered data.for the Reynolds number range between the points shown since the high dynamicpressure drove the gages off scale.same variation of afterbody pressure with Reynolds number occurred even thoughthe test conditions were very different.

There were no pressure data obtained in flightThis comparison indicates that much the

CONCLUDING REMARKSIn concluding, it is necessary to point out that although the new experi-mental techniques described herein have obtained test results which are helpfulin determining the Apollo heat-shield design, these data were obtained underconditions far different from those that will be encountered during the'Apolloreentry. Therefore, data from flights such as Fire, Mercury, Gemini, and theearly earth-orbital Apollo flights must be carefully analyzed and used for con-firmation of' the design.

REFERENCES

1. Jones, Robert A.; and Hunt, James L.: An Improved Technique for ObtainingQuantitative Aerodynamic Heat-Transfer Data With Surface Coating Materials.Paper No. 63-13., Am. Inst. Aeron. Astronaut., Jan. 1965.2. Jones, Robert A,; and Hunt, James L.s Effects of Cavities, Protuberances,and Reaction-Control Jets on Heat Transfer to the Apollo Command Module.NASA TM X-1063, 1965.3. Jones, Robert A , : Experimental Investigation of the Overall Pressure Dis-tribution, Flow Field, and Afterbody Heat-Transfer Distribution of anApollo Reentry Configuration at a Mach Number of 8.(Supersedes NASA TM X-699. NASA TM X-813, 1963.


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L-2458-1Figure 1 L-2458-2Figure 2

Figure 4 L-2458-4Figure 3


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Figure5

HEAT-TRA OL JETR

L-2458-6Figure 6

SEPARATED-L$E$ THICKNESS

Figure1

16

Figure8


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TECHNIQUE FOR MEASURING SEPARATED-LAYER THICKNESS

STANTONNUMBER.& 10-3-

n

W * % % !

Figure 9

AFTERBOW PRESSURE DISTRIBUTIONS

r a.350' d.mor a=O",o15b/ ".5X106

I 4 X106LII,,I,III,IIIII0 .I .2 .3 .4 .5 .6 .7 0 . I .2 .3 .4 .5 .6 .7s/d s/d

SEWRATED-LAYER THICKNESSa = 35 "

Figure 10

HEAT-TRANSFER DISTRIBUTION ON AFTERBODYU.0'

1 6 %

Figure 11 Figure 12


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STANTONNUMBER,

COMPARISON OF PROJECT FIR E AFTERBODY PRESSURE M TAWITH WINDTUNNEL MT AP P t. 0 5

.021 I O O.01 0 0."t.M* mu,m TtPR3 0 0 1,000PROJECTLlRE 1 8 .4 2 ,6 4 0 6 ,4 5 0I PROJECT FIRE 1 39.8 28.642 19,420

, 0 0 5 1 0 LRC VARDEN S. 8

104 2x104 5x104 105b , d x105 5x105 106 2 x 1 0 6

Figure 13

18

Figure 14


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3 . AERODYNAMIC HEATING CHARACIXRISTICS OFAPOLLO LAUNCH CONFIGURATION

By Robert L. Sta l l ings , Jr., and E a r l A. Price, Jr.

Hea t-tr ans fer measurements have been obta ine d i n wind-tunnel t e s t s ofO.dt5-scale models of the Apollo launch configuration.a t Mach numbers from 2.98 t o 4.44 and Reynolds numbers per foot from 2 x l a o6 x 106. The measurements were i n good agreement with e xi sti ng th eo ri es basedon l o c a l measured pressures i n regions not a ffe cte d by surfac e protuberances.In regions affected by these protuberances, l a rge hea ting ra t es were obtained;t h e maximum values were generally confined t o th e vic in i ty of the protuberancei n s t a l l a t i o n .

The tes ts were cond cted

dood cor re la t ion of the experimental heating data w a s obtained through thet e s t range of local Reynolds numbers for the region of t h e command and s e rv icem o d u l e s not a ffe ct ed by th e surfa ce protuberances.with f l i g h t heating ra te s obtained during the launch phase of the Apollo m i s -s ion A-101.tunnel results, it i s believed that the agreement could have been improved hadde ta i l e d p ressure data fo r th e f l ig h t configura tion been ava i lab le .

These data w e r e comparedAlthough f a i r agreement was obtained between f l i g h t and w i n d -

INTRODUCTION

The aerodynamic heating encountered by spacecraft during launch i s muchless severe tha n t h e heatin g associated with reentry; however, f o r th os e launchveh icle s reaching hypersonic fl i g h t speeds wit hin the ea rt h' s atmosphere,thermal protect ion f or th e spacecraf t and other components of the launch con-f i g u r a t i o n i s ge ne ral ly requi red. Due t o the w e i g h t penalty of th e heat-p ro tec t ion material, it i s highly desi rable t o be able t o predic t t he aerody-tlamic hea ting over the launch configuration i n o rder t o employ th e hea t pro-tection only where necessary.namic heati ng t e s t s were conducted i n the Langley Unitary Plan wind tunnel ona series of O.OLc5-scale models of th e Apollo launch confi gurati on. The tes tswere conducted through a range of Mach numbers from 2.98 t o 4.44 and a range offree-stream Reynolds numbers per foot from 2 x 1 06 t o 6 x lo6. Correlations ofth e experimental data are presented. The data are also compared with theoreti-c a l d i s t r ibu t ion s based on l o c a l measured pressures.tunnel t e s t s are compared i n t h i s presenta t ion with f l ig h t data obta ined dur ingthe Apollo mission A-101, which incor pora ted t h e BP-13 sp ac ec raf t and t h e%turn I launch vehicle (SA-6).

It w a s pr inc ipa l ly fo r th i s reason th a t aerody-

Results fran the wind-


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SYMBOLS

dhMooNSt, 2rRdR2%oS

Tw/Taw

service-module diametercoeffici ent of heat tr an sf erfree-stream Mach number. 5s Stanton number based on local conditionsservice-module r ad iu s, o r command-module ra diusfree-stream Reynolds number based on service-module diameterlocal Reynolds number based on surface length sfree-stream Reynolds number per foo tsurface length (see f ig. 4 )ra t i o of wal l temperature t o adiab at ic w a l l temperature

APPARATUS AND TEST CONDITIONS

The tests were conducted i n t he high Mach number t e s t section of t h eLangley Unitary Plan wind tunnel.t e s t Reynolds numbers with t he launch tra je ct or y values of t h e Apollo m i s -sion A-101 (BP-13) i s given i n f igure 1.on the service-module diameter i s plo t ted as a function of Mach number.curve represents th e launch traj ecto ry, and the crosshatched regions ind icat et h e Reynolds number range f o r th e wind-tunnel t e s t s .Reynolds numbers f o r th e tu nn el t e s t s are approximately 1/3 t h e f u l l - s c a l ef l i g h t Reynolds numbers. With increasing Mach number, which corresponds t o anin cr e as e i n a l t i t u d e and a reduction i n th e fl i g h t Reynolds numbers, clo se rsimulation i s obtained. M, = 4.44, t h e t r a j e c t o r yReynolds numbers f a l l within the tunnel range.ure, t h e maximum heat f lux for t h e e x i t t r a j e c t o r y was expected t o occur withint h e Mach number range of t h e tunnel invest igat ion.

A comparison of t h e wind-tunnel range ofThe free-stream Reynolds number basedThe

A t I = 2.98, t h e m a x i m u m

A t the highest Mach number,Although n ot shown i n th e f ig -

Figure 2 shows the basic configuration tested. It w a s constructed af th in-sk in s ta in less s t e e l and consi s ted of t h e follow ing components: escape tower,command module, serv ice module, lunar-excurs ion-module flare, and a small por-t i o n of t h e Saturn IVB stag e. The ov era ll len gth of th e model w a s 48.7 inchesand the maximum diameter w a s 11.7 inches . The model w a s instrumented with187 thermocouples l oca ted t o provide complete lon gi tud ina l and circumferentialheat d is t r ibut ions . "he thermocouple locations are ind icated i n the figure byth e small ci rcl es . Pressure instrumentation consis ted of approximately20


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50 pressure o r i f i c es loca ted only along a single ray on the command and servicemodules.pressure measurements could be applied a t an y circumferent ia l s ta t ion.when surfa ce protuberances are added t o th e bas ic configuration, the se measure-ments would only apply along the instrumented ray.

Due t o t h e symmetry of th e basic configuration, t h e results from t h eHowever,

The second configurationtested, which i s shown i n figure 3 , consisted oft h e components shown f o r the basic configuration plus the reaction controlmotors, umbi lical fa iri ng , and scimiter antenna.The hea t-tr ans fer co ef fic ien ts were determined from transient temperaturemeasurements using t h e thermocouple in st al la ti on i n th e m o d e l skin as a calo-rimeter.tunnel stagnation temperature obtained by bypassing th e tun ne l coole r system.A complete description of the Langley Unitary Plan wind tunnel i s given i nreference 1.

The tran si en t skin temperatures resu lte d from a sudden increase i n th e

RESULTS AND DISCUSSION

He at -tr ansfer measurements obtained on th e command and se rv ice modules ofthe bas ic conf igurat ion are shown i n f igure 4 .c ien t i s p lo t t e d as a function of the dimensionless su rface length s/r, wheres i s the surface length from the forward stagnation point on the commsnd m o d u l eand r i s the rad ius of t h e command module a t th e base, o r th e service-moduleradius.Reynolds numbers per foot of 2 x lo6, 4 x lo6, and 6 x lo6. Also shown i n t h ef igure for comparison with the experimental data are theo re t i ca l tu rbu len t d i s -tr ib ut io ns determined by t h e reference-temperature method of ref ere nce 2 andl o c a l measured pr es su re s. The instru mentation lo ca ti on s on th e cammand andgertrice modules for which data are presented are indicated by the small ci rc lesi n the sketch.

The lo ca l heat-tra nsfer coeffi-

!&e r e s u l t s are shown f o r Mach nunibers of 2.98 and 4.44 and for

The heating rates obtained on the command module a t a Mach number of 2.98and t h e lowest Reynolds number in cre ase with inc rea sing va lue s ofs/r This increase -is associa ted with t he decrease i n th e extent of thesheltering effect produced by the escape tower.more apparent on th e pressure d istr ibu tion s i n th i s region, as i s indicated bythe good agreement between theory based on measured pressures and the experi-mental hea t-transf er data. The decrease i n heating i n the base region of t h ecommand module i s due t o t h e propagation of a pressure drop back through thesubsonic portion of th e boundary layer; t h i s pressur e drop occurs a t the junc-ture of t h e command and s erv ice modules.on th e ser vi ce module a re l ess tha n theory for s /r < 3; howeyer, th e measure-ments obtained a t t h e t h r e e aft s ta t ions a re in good agreement with theory.The low heating rates on th e forward section of th e serv ice module ar e believedt o be pa rt ia l l y due t o ei th er one or both of th e following two factors: (1) h eseparation region that occurs a t t h e jun ctu re of t h e command and se rv ic e modulesand ( 2 ) conduction loss es t o an inte rn al support assembly loca ted under t he sk infor a t taching the react ion control motors.

s/r up t o1.6. This shel ter ing effect i s much

The measured heating rates obtained

The same t rends i n hea ting fo r the


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command and service modules were obtained f o r th e two Mach numbers over theReynolds number range.hea t ing ra tes with Reynolds number and.Mach number ar e i n good agreement wi ththeory.in cr ea sing Reynolds number and dec rease i n he at in g with in cr ea si ng Mach number.

I n general, th e var ia tio ns of t he magnitude of 'theThese variati ons consisted of t he expected increase i n heating with

The eff ect on the heating dis tr i bu tio n of adding surfa ce protuberances t oth e basi c configuration i s shown i n f ig ur e 5 .t h i s f igu re represen t data f o r th e basi c configuration, which were given inf i g u r e 4.t i o n .f igura t ions .presented were produced by the scimiter antenna and reaction control motors.The scimiter antenna was located adjacent t o th e ray of instrumentation, andthe longi tudinal center l ine of t h e reactio n contro l motor w a s located alongthe instrumentation ray.t o the reg ion of t he i r ins ta l l a t ion ; the protuberances generally cause anincrease i n heat ing.compared with the increas e i n heatin g upstream of th e rea ctio n con tro l motorsi s re la ti v e ly modest. The maximum increase i n hea ting ad jacent t o the sc imi te rantenna w a s approximately 40 per cen t, whereas th e maximum inc re as e ahead of t hereaction control motors w a s approximately 600 percent.t o es t imate theoret ica l ly these in terference heat ing ra tes because of th e lackof pressure instrumentation located i n th e same posit ions as th e thermocouples.The m a x i m u m heating ra te s upstream of t he r eac tion con tro l motors are approxi-mately the same as maximum values obtained on a f l at -p la te surface upstream ofa r igh t c i rcu la r cy l inder in prev ious t es t s conducted i n t he Langley U nitaryPlan wind tunnel.are s l ig ht ly less than those of the bas ic configuration.

The open symbols and curves i nThe solid symbols represent data obtained on the complete configura-

The largest ef fe ct s along the instrumented ray f o r which dat a areThe thermocouple loc atio ns and t e s t conditions a re th e same f o r both con-

The maximum ef fe ct of these protuberances i s confinedThe increase i n heat ing adjacent t o the sc imit er antenna

No attempts were made

The heating r at es i n the wake of t he rea ctio n con trol motors

The coeff ic ient of heat t ransfer i s valuable f o r d iscuss ing t h e r e l a t i v emagnitudes of' heating on a given model through a range of t e s t Conditions; how-ever, it i s of l i t t l e use as such when the results for a smal l-scale model i na wind-tunnel environment a re t o be ap plie d t o a ful l -sca le conf igurat ion i n i t sf l i g h t environment. This i s due t o the f ac t t ha t the var ia t i on of h wi thReynolds number i s sensi t iv e t o the method of varying &. For example, i f f o ra f l a t p l a t e i s increased by increasing th e m o d e l scale, th e value of hw i l l decrease; whereas, if %, i s increased by increa sing pressure, th e valueof h w i l l increase. A more sui t able parameter f o r th i s appl icat ion i s t h el o c a l dimensionless Stanton number, which f o r a f l a t pla te cons i s ten t lydecreases w i t h increasing Reynolds number a t a constant Mach number and w a l ltemperature. Furthermore, f o r turbule nt pipe flow th e va ria tio n of NStY2 withReynolds number can be minimized by multiplying the Stanton number by the localReynolds number ra i s e d t o a power of 0.2.

The parameter NSt,ZRzo*2 w a s evaluated from the heating rates measuredon t h e command and ser vi ce modules of t h e ba s ic co nf ig ur at io n fo r which de ta il edpressu re d at a were ava ilab le.shown i n fi gu re 6, where th e co rre la tio n parameter i s p lo t t e d as a function oft h e l o c a l Reynolds number f o r each t e s t Mach number.

The valu es ob ta ined on th e command module a r eThe experimental data are

22


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indicat ed by the open ci rc le s.b a t i c conditions, using th e reference-temperature method of refe renc e 2, f o rcomparison with f l i g h t data as w i l l be discussed subsequently.the f igure are adiabatic theoretical values determined by the reference-temperature method and indicated by the solid curve.butions f a l l within the data sc at te r through th e t e s t range of Mach numbers.Although there i s a slight increase i n t h e the or y with i nc rea sing Reynolds num-b er , t h i s e f f e c t i s small and con sis ts of approximately a 10-percent increase

These experimental data w e r e ad jus ted t o adia-Also shown i n

The theoretical d i s t r i -

3s rr res!!+ n f eE ir?crease in Rep-olda I?W!ber by Ell order of ?m&nit12de=An attempt w a s made t o compare these wind-tunnel co rr el at io ns with f l i g h t

(See f i g . 7.) The wind-results using heat fluxes obta ined on the command module of t h e BP-13 space-c ra f t dur ing the f l i g h t of the Apollo mission A-101.tunnel data shown by th e open ci rc ul ar symbols and t h e th eo re ti ca l valu es aref o r th e command module of t h e bas ic config uratio n and were pr esente d i n f i g -ure 6. The solid symbols represent values obtained on the command module of thef l i g h t vehicle a t th ree d if fe ren t loca tions . P ressure da ta fo r the f l i gh t t e s t swere not available a t th e wind-tunnel Mach numbers; t her efor e, pre ssu re da taobta ined f o r the bas ic configuration i n the wind-tunnel t e s t s were used as l o c a lconditions i n evaluating the values of the f l ig h t cor re la t ion paramete rNSt,2Rz0-2. Tw/Tawthan the wind-tunnel results.two sets of data, t h e f l i g h t results and wind-tunnel re su lt s were adj uste d t oadiabatic conditions by the reference-temperature method.t i on s required t o evaluate the f l i gh t corre lati on parameter, the agreementbetween th e wind-tunnel and f l i g h t r es ul ts i s f a i r .

The fl i g h t re su lt s were obtained a t much lower value s ofI n order t o make a valid comparison between the

Based on the assump-

Sim ilar comparisons f o r th e service module a r e shown i n fi gu re s 8 and 9.The values of the f l i g h t corr elat ion parameter presented i n f igu re 9 are basedon pressure measurements obtained on t h e se rv ice module of th e b a si c configura-t i on dur ing the wind-tunnel t e s t s .were adjusted t o adiab atic conditions f o r reasons discussed previously. Again,based on th e assumptions required t o evaluate the f l ig ht data, the agreementbetween experiment and theory i s f a i r .Both the wind-tunnel and f l i g h t results

CONCLUDING FGMARKS

Hea t-tr ans fer measurements obtained i n wind-tunnel t e s t s of 0.045-scalemodels of t h e Apollo launch config uratio n were found t o be i n good agreementwith ex ist ing the orie s i n regions not affe cted by surface protuberances. I nregions aff ecte d by the se protuberances, la rge heating rates were obtained; themaximum values were ge nerally confined t o th e v ic in it y of th e protuberancei n s t a l l a t i o n .

Good co rre lat ion of th e experimental heating d ata was obtained fo r th e t es trange of l o c a l Reynolds numbers f o r the region of t h e cammand and s er vi ce mod-ul es not affe cted by t h e surface protuberances. These data were compared withf l i g h t heating rates obtained during the launch phase of the Apollo missionA-101. Although fa i r agreement w a s obtained between flight and wind-tunnel


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resul ts , it i s bel iev ed t h a t th e agreement could have been improved had de ta i l edpressure data f o r t h e f l i g h t configuration been available.

REFERENCES

I. .-ion.: Manual f o r Users of the Unitary Plan W--?d Tunnel Facil t ies of theNatio nal Advisory Committee f o r Aeronautics. NACA, 1956.2. Sommer, Simon C.; a d Short, Barbara J.: Free-Flight Measurements ofTurbulent-'Boundary-Layer Skin F r ic t io n i n t h e Presence of Severe Aero-dynamic Heating a t Mach Numbers From 2.8 t o 7.0. NACA TN 3391, 1955.

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OF

Figure 1 Figure 2 L-2475-2

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HEATING DISTRIBUTIONS ON BASIC CONFIGURATIONTURBULENlR- THEORY

COMPARISON OF HEATIN G DISTRIBUTIONS ON BASICAND COMPLETE CONFIGURATIONSTURBULENT

Rm THEORY

slr s i r

Figure 4

CORRELATION OF WIND-TUNNEL DATA ON COMMAND MODULEADJUSTED FOR COOLING

Figure 5

CORRELATION OF WIND- TUNN EL AND FLIGH T DATAON COMMAND MODULEADJUSTLO FOR COOLING

0 W I N D N N N E L+ F L I G H T I M l S S l O N A - 1 0 1)- URBULENT THEORY

.ta Mm-3.71N s t j ? ; :E[02 ~ Om sa.. "

Figure 6

26

Figure 7


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CORRELATION OF WIND-TUNNEL M TA ON SERVICE MODULEAOJIISTFII FOP T M L I N T .0 BASIC CONFIGURATION

NRBULENT THEORY

Figure 8

CORRELATION OF WIND-TUNNEL AND FLIGHT WTAADJUSTED FOR COOLING- RBULENT THEORYON SERVICE MODULE0 WIND TUNNELFLIGHT IMlSS lON A-101)

--aNs,,& , ,

0IO6 5 Id 5

Rl

Figure 9


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0. FLIGHT AND GROUND TESTS OF

M O L L 0 HEAT-SHIELD MATERIAL "By W i l l i a m A. Brooks, Jr., Stephen S. Tompkins,and Robert T.

SUMMARY

A reen t ry f l ig h t t e s t and ground tes t s

Swann - -- -

have been made by t h e LangleyResearch Center t o evalu ate the Apollo heat-s hield ma teri al.revealed a dele teri ous effe ct of pressure a t levels above atmospheric pressure.The effect of pressure i s a rapid and irregular char erosion.char removal model involving pressure i s postulated, which yi eld s an al yti ca lre su lt s compatible with the f l ig h t data and ground-test re su lt s . Althuugh th eac tu al mechanism producing ra pi d erosion i s not presently detailed, it i sbeliev ed t o be coupled with t he extreme poros ity of t h e material .indicate that the postula ted pressure effect i s no t a considerat ion for thepresen t ly def ined Apollo reentry t r a jec tor i es . I t appears that oxidation w i l lbe the predominant char removal mechanism and the present Apollo heat-shielddesign i s therefore conservative.

These t e s t s haveA mathematical

Calculations

INTRODUCTION

The R-4 experiment of the Langley Research Center Scout Reentry HeatingProject w a s launched on August 18, 1964, with the primary objective of obtaininga reen try f l ig ht evaluation of t he Apollo heat-shield material. Many t e s t s oft h i s material have been performed i n ground f a c i l i t i e s a t the Langley ResearchCente r t o suppor t the f l i gh t t e s t .b r ie f ly the f l ig h t and ground te s t s and t o d i scuss the results .The purpose of t h i s paper i s t o d es cr ib e

SYMBOLS

A empirical cons tan tQ accelera t ion due to e ar th 's gravi tyH stream enthalpy

ra te of production of gases inje cte d in to boundary la ye rrhvpP pressure

t o t a l pressure behind shockp t , 2

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Q t o t a l heat i n p t6 heating ratei n ne t heat ing ra te t o su rface-9 dynamic pressureTS sukace temperat-t t imeVO voltageXC thickness l o s s ra te o f char mater ia lY entry angle w i t h respect t o loc al hor izon7 aerodynamic shearSubscript:m a x m a x i m u m

DESCRIPTION OF R -4 PAYLOAD AND INSTRUMENTA!I'ION

The R -4 payload, including a spherical rocket, w a s mounted on a standardfour-stage Scout launch vehicle. A photograph of t h e payload i s shown as f i g -ure 1.consisted of 1.23 inches of th e Apollo hea t-sh ield material bonded t o 5/8 inchof st ai nl es s- st ee l backup stru ctu re. There were 26 thermocouples and 1 2 abla-t i on sensors in th e ab la tion mate r ia l a t var iou s lo ca ti on s. The next componenti s the afterbody which was a 0.064-inch-thick magnesium-alloy s h e l l coveredwith approximately 0.5 inch of Teflon. A portion of the spherical rocket motorwhich served as th e f i f t h stage can be seen behind the afterbody section . A tthe bottom i s the adapter un it , which w a s attached t o th e Scout four th stage,and other interstage equipment.was about 3 fee t long.diameter w a s 11 inches.

A t the top i s the payload nose cap which was the primary experiment and

The payload, without t h e int er st ag e equipment,The payload base diameter was 18 inches and the forwardFigure 2 shows some of the construction and instrumentation details of thenose cap.indicated.of th e Apollo heat-sh ield mate rial.comb which had been bonded t o th e s t e e l s ub st ru ctu re .of forming t h e honeycomb around the relatively sharp corner and because of themore severe environmental conditions, a molded version of t h e same materialwithout honeycomb w a s fabr ica ted in to a ring which surrounds the centralportion.

I n th e sect ioned v i e w , t h e vario us components of t h e nose cap areThe ce n tr al po rti on of the nose cap w a s t h e t e s t area and consistsThis i s Avcoat 5026-39 gunned into a honey-Because of th e di ff ic ul ty

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The front view shows t h e dist r ibu tion of th e instrumentation over t h e nosecap.depths through the thickness and arranged i n ci r cl e s of t h r e e d i f f e r e n t radii.The thermocouples i n the outer ci rc le w e r e located so that they would providereadings a t 0.1-inch in te rv al s st ar ti ng from th e ex ter na l surface . Two of thes ethermocouples were a t the bondline between t h e a b la to r and the substructure.The thermocouples i n th e middle ci rc le were located a t O.05-inch in te rv al s t oa depth of 0.6. inch. The fou r thermocouples i n th e inn er c ir c le were loc atedst 3.1-1nch 1ilti2i-v-i5lS.temperature distr ibutions, t h i s arrangement of thermocouples al s o providesinformation on the temperature gradients along contours a t various depths.

The open symbols represent thermocouples which were located a t various

In d d i t i z l n t o pro>-mLri i i i - iomzt ion G i i the depth-kiisz

The pat tern of the abla t ion sensors i s similarly shown by the solid sym-bols . A t the center, there were four make w i r e abla t io n sensors located a t0.1-inch intervals from the f ron t surface. I n the inner c i rc le of abla t ionsensors, the re were fou r spring w i r e sensors located a t 0.1-inch intervals w i t hth e foremost sensor being 0.14 inch f romthe surface . In the outer c i rc le ofab lat io n sensors the re were fo ur l i gh t pipe sensors located a t approximately0.1-inch in te rv al s from th e fr on t surface.

A typical thermocouple assembly i s shown i n fi gu re 3. A plug was formedof t h e ab lat ion mat eria l and grooved up th e sid es f o r th e thermocouple w i r e .The pair of chromel-alumel thermocouple wires was separated a t the base of t h eplug.then threaded through quartz tubing.t h e plug.welded t o form the hot junction.dis turbance i n th e temperature f i e l d and t o minimize th e p oss ibi l i t y of e lec-t r i c a l s h o r ti n g by t h e char.hole i n th e back of th e heat shield.

A t th i s point , the insu la t ion was stripped from the wires which wereThe tubing w a s terminated a t the end ofThe bare wires were laid i n a groove across the end of the plug andThis design w a s se lec ted t o p rovide a minimumThe plug assembly was then inse r ted in to a bl ind

Figures 4, 5 , and 6 show schematics of the abla t ion sensors . These sen-A secondary obJective ofo rs were developed by the Langley Research Center.the f l ight experiment was t o p rovide f l i gh t da ta ess en t i a l t o the con tinueddevelopment of th es e type s of sensors.The sensing element of the m a k e wire sensor shown i n f igure 4 i s a p a i r ofopen- circuit ed 10-mil wires which a r e peened a t th e end and embedded i n thea b l a t i o n material.wire, the high e l ec t r ic a l conductiv ity of the char causes the c i r cu i t t o becompleted and thus an indication of the event i s provided. The c i r c u i t i s com-

p l e t ed i n t h e char l ay er a t some point between the poi nt a t which the charbegins t o form and th e exte rnal surface of th e char. Four of these sensorswere mounted i n a single plug a t the center of the nose cap.

When th e a bl at io n ma ter ia l chars t o th e peened ends of the

The spring wire abl atio n sensor i s shoim i n f i g u r e 5 . A 20-mil molybdenumtube i s a t tached t o a snap-action switch. A tungsten w i r e i s a t ta c he d t o t h ele af sp rin g of th e switch, passed through th e tube, and knotted a t the end ofthe tube i n such a manner that the switch i s held open. A s the ab la t ing sur-fa ce app roaches t h e end of t he senso r, th e temp eratur es become high enough t ocause t he molybdenum tube t o so fte n and, i n turn, t o re le ase th e wire and


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permit t h e s w i t c h t o c lo se .Because the temperat.ure g r d i e n t s a re steep near t h e surface of the ablator,th e end of the sensor i s gener ally a t th e s urface when the wire i s released andthus senses the surface location.

Thus a n indication of the event i s provided.

The l igh t pipe a bla tio n sensor shown i n fig ur e 6 u t i l i z e s a n o p t i c a l f i b e rwith a high m e l t i n g temperature t o transmit l i gh t t o a photo diode whose resis t-ance i s a func tion of th e in te ns i t y of the l ig h t .the l ight pipe and the diode, it i s p o ss ibl e t o g e t a sharp response i n diodecurrent a t temperatures corresponding t o those expected a t the abla tor surface .Thus t h i s type of sensor can be employed t o sense th e lo cat ion of t he surfac e.

By using a f i l t e r between

REENTRY ENVIRONMENT

A nominal reentry trajectory w a s selected t o provide reentry heat ing t h a twould satisfactorily simulate the heating which would be experienced by anApollo vehicle i n a 20g emergency reentry. The ac tua l reen t ry t ra jec to ry veryclosely matched the nominal trajectory and i s shown i n t h e a l t i tude-ve loc i typlane i n f igure 7. A t f i f t h-s t age burnout, the a l t i tu de was about 395,000 f e e tand the velocity was 27,800 fps.The labels on the curve identify certain f l ight events and the t ime of

This i s a majoroccurrence measured from launch.f o r 61 seconds as shown by the heavy portion of the curve.por t ion of the heating period which i s shown by the hatched portion of thecurve.end of blackout.

Note that real-time telemetry w a s blacked out

During blackout, a tape recorder s tor ed the da ta f o r playback a t the

Although t h e body motions during r een try ge ne ral ly were not lar ge , theywere somewhat greater than anticipated.heating and maximum pressure, th e sp in ra te inc reased f rom the in i t i a l 3 cpst o 5 cps and then slowly decreased t o 3 cps.th e angle of a t tack w a s l e s s t h an loo .heating and maximum pressure , angle-of-a ttack os ci l la t io ns increased t o a maxi-mum value of about l 5 O , and then decreased t o a le v el of about 1determined tha t t h e inc reased body motions d id not produce si gn if ic an t changesi n th e heating environment that the nose cap w a s expected t o experience.

During t h e t ime per io d between maximumUnt i l the t ime of m a x i m heat ing,In th e t ime period between m a x i m

It has been

Figure 8 shows calculated hi st or ie s of th e stagnation-point heating ra te,the t o t a l pressure, and t h e maximum aerodynamic shear stress for zero angle ofattack. The maximum t o t a l h e at in g rate was 775 Btu/ft 2-sec and occurred a t476 seconds.maxirmun pressure w a s 6500 l b / f t 2 , o r 3.1 atmospheres, and occ urred a t 487 sec-onds. The point on the nose cap a t which the aerodynamic shear w a s t h e g r e a t e s tw a s on the molded outer r i n g .value of 27 l b / f t 2 a t 482 seconds fo r laminar flow.t h a t over the instrumented portion of the nose cap the shear w a s l ess t han1 0 l b / f t 2 .

The heating pulse w a s approximately 120 seconds i n durat ion . "he

The history of t h a t shear ind ica tes a m a x i m u mI t should be pointed out


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Table I shows a comparison of th e ree ntr y environment f o r th e R -4 experi-ment and th a t expected f o r Apollo.planned 20g emergency re en tr y which involve s a s kipout resulting i n two heatpul ses . The numbers shown i n th e ta b le are f o r th e f i r s t he at in g pu ls e. How-ever, as can be seen a t th e bottom of the ta ble , th e ef fe ct of the second pulsei s t o add about 1 6 percent t o the heat input of t he f i r s t pulse and t o increasethe heat ing t i m e by a fac to r o f 2$.heat ing rate: t o t a l heat input, and heating t i m e f o r t h e R -4 ree nt ry matchedthose f o r th e Apollo 20g reentry very well, pa rt i cu lar ly f o r th e f i r s t heatingpul se. However, because of fundamental diff ere nce s between b a l l i s t i c andl i f t i n g r ee nt ry , t h e R-4 environment w a s more severe with regard t o pressur eand shear.w a s th ree times t h at expected f o r th e Apollo reentr y. The ma xi mum shear fort h e R -4 reentry w a s sl ig ht ly more than three times th a t fo r Apollo .

The Apollo reentry considered i s a recent ly

It can be seen th at the values f o r the

The maximum pressure experienced by the a bl ato r i n the R-4 reentry

RESULTS AND DISCUSSION

Typical Temperature HistoriesThe temperature readings for 6 of the 24 thermocouples t h a t were used areshown i n f ig ur e 9 as a function of time from launch.ra d ia l po si tio ns of thes e thermocouples; t h e darkened area represents t h e moldednose-cap ring. A l hermocuples were of chromel-alumel with a usable range ofabout SO00 F. Thermocouple data were commutated a t a rate of 5 readings persecond.

The s ke tch shows t h e

Data are shown for four depthwise locations measured from t h e o r ig in a lf ront-surface locat ion.a t t h e i n t e r f a c e between the ablator and the s tee l substructure.shown for each depth is the l a s t smooth on-scale reading.The thermocouple shown a t a depth of 1.25 inches wasThe l a s t point

The square symbols show readings at a depth of 0.2 inch, tak en a t t h r e eradial locat ions .uniform pen et rat io n of temperature over th e nose-cap sur fac e durin g th e ea rl yheating - before m a x i m hea tin g which corresponds t o th e po int where t h ereading f o r the thermocouple a t 0.3 inch goes off-scale.These and similar resu l t s a t a depth of 0.1 inch indicate a

Generally speaking, f o r depths up to about 0.35 inch, th e temperaturet r a c e s show a gradual response as indicated by thermocouples a t depths of 0.2and 0.3 inch shown i n fi gu re 9.response w a s very ra pid and, because o f t h e sampling rat e, th e maximum on-scalereadings were only a few hundred degrees.

A t th e gre at er depths, th e thermocouple

Note that the temperature of 850 F i s i d e n t i f i e d as th e p yrolys is tem-perature .of temperatures st ar t i ng a t l e s s t h an 500 F.Actually, pyr oly sis of t he ablatio n ma teri al take s place over a rangeTypical r es ul ts from thermal degradation stu di es a re shown i n fig ur e 10.The mass l o ss due t o thermal degradation i s shown as a funct ion of temperature

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f o r a n i n e r t and an oxidizing atmosphere.pressure were as shown.a t t h e lower temperatures bu t the g r e a t e s t rate of degradation occurs a t about850 F. This temperature can be ca l led a cha ract er is t ic p yrolys is temperaturebecause half of t h e t o t a l mass loss experienced has occurred a t t h i s point .The total mass loss was 53 percent. The 47 percent remaining as residue i sequal t o about 14 1b/ft3.20 t o 21 Ib/ft3.of pyrolysis gases i n the char during the arc- je t tests.

The temperature r ise rate and th e ga yI n helium, th e material experiences some degradation

Residues from arc-jet tests have been measured a tThe difference i n den si ty probably results from condensation

The te s t i n oxygen w a s conducted prim arily t o show the ef fe ct of oxida tion,Degradation becomes pronounced a t a lower temperature and only about half asmuch residue results.reinforcement. Therefore, it i s concludedthat near ly ha l f t h e r e sidu e t h a tresults when t h i s material i s ablated w i l l not be oxidized.As a matter of fact, the res idue i s p r im a r i l y t h e s i l i c a

It i s obvious from figure 9 t h a t t h e shallow thermocouples indicate charIt i s also obvious that, by us ing the charac te r i s t i c pyro lys i semperatures.temperature and temperature data such as shown, a time h i s to r y of th e penetra-t i o n of the pyrolysis zone can be constructed.A rapid temperature response i s an ind ica tion th a t the char and t h ereceding surface are approaching the thermocoyple junction.couples may be used as approximate surface sensors. This has been done byusing th e last smooth on-scale temperature reading a t each depth.believed t h a t the accuracy of this method of de tec t ing sur face loca t ion i sg r e a t e r for the lar ge r depths .

Thus the themo-It i s

Flight Surface Recession HistoryFigure lJ. shows a h i s to r y of surface recession as determined by thermo-The k n m lo ca ti on of each sensing element i souples and abla t io n sensors.p lo t ted aga ins t the t i m e a t which it funct ions - except f o r th e thermocouples,f o r which the t i m e of the last va lid on-scale reading w a s used.Data are shown by the open symbols f o r 23 of t h e 24 thermocouples used.One thermocouple exhibited an er ra t i c response ind ic ati ng malfunction and nosymbol i s shown.Three of t h e light pipe sensors produced data t ha t do not agree with the gen-e r a l t r e n d shown i n t h i s f igure. Therefore, it i s concluded t h a t they alsomalfunctioned and those data are not shown.

The ablation sensor responses are shown by the solid symbols.

The data c lea r ly de f ine the trend of t h e surface recession. Bet te r def in i -t i o n w a s obtained for the upper portion of the curve as a result of d i s t r i b u -t i o n of the sensors.in to th ree regions:rate is on th e o rder of 0.01 in./sec, a t r a n s i t i o n r eg io n w h i c h produces a kneei n th e curve, and a reg ion a f te r maximum heating where the recession i s on theorder of 0.1 in./sec.

Generally speaking, t h e recession history can be divideda region before maximum heat ing f o r which the recession

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I n the f i r s t region, th e receding surface i s bel ieved t o be located abwethe points derived from the thermocoxples - possibly by as much as 0.1 inch.Although t he di ffe ren ce between t h e data from th e make w i r e and spring w i r eabla t ion sensors indicates a char thickness of about 0.06 inch, there i s evi-dence that the char i s thicker.penetration determined from the temperature data i s also shown.represents the pyrolysis region between char and the v i r g in materials.t h a t the make wire sensors are functioning a t points much c lo s er t o t h e i nd i-cated surfa ce. Recently conducted ar c- je t t e s t s ind ica te tha t make wire sen-sors, similar t o those used i n the present f l i g h t t e s t , do not funct ion u n t i lthe temperature exceeds 2000 F.t o complete the e le c t r i c c i rc u i t u n t i l th i s t empera tu re i s reached..it i s concluded that the f l i g h t sensors did not make i n th e pyro lys i s reg ion.The differences between the data from the spring w i r e sensors and the interfacecurve i nd ic ate t h a t during ea rly heating th e char may have been 0.10 t o0.15 inch thick.got thinner.

The previously discussed hi st or y of p yroly sisThis l in e thusNote

The char i s not electrically conductive enoughTherefore

As heating increases, th e indica tion i s tha t the char l ayer

The knee of th e curve, which f a l l s between 474 and 482 seconds, correspondst o rap id ly changing environmental cond itions.generally decreased by 30 t o 35 percent.cent.most significant environmental change.sure i s an important consideration f o r the rapid recession of th i s mate r ia l .

The heating rate and enthalpyThe shear increased by about 20 per-However, t h e pressu re increa sed nea rly 100 perce nt and i s by fa r t h eOne i s thus l ed t o be lieve t h a t pres-

The thermocouple da ta in dic ate d t ha t t h e ero sio n became gr ea te r toward th eedge of the nose cap during t h e phase a f t e r maximum he at ing. The thermocouplesa t t h e greater depths are adjac ent thermocouples lo ca ted i n t h e oute r c i rc le o finstrumentation.

The readings of t h e bondline thermocouples, ind ica ted by th e l a s t twoHowever, more important, thesei r c l e s , also ind ica te a difference i n erosion.readings indicate that the s teel substructure w a s exposed as early as about486 seconds and th a t very li ke ly most of the ab la t ion material w a s consumed byabout 490 seconds. It i s therefore concluded that the payload survived theheating only because of t h e th ic k steel substructure which could absorb theremaining heat without f a i l i ng .The s o l id and dashed curves shown i n f ig ur e 11 result from analysis andar e discussed i n another sect ion.

Prediction Procedures for R -4 ExperimentsUsual ly , i n order t o get a complete picture of what happened i n a f l i g h t

Several different procedures have been employed t o analyze theThese methods are outlined as follows:t e s t , it i s necessary t o augment f li gh t d at a w i t h a theore t ica l ana lys i s ofperformance.R-4 reentry .

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Method A:MechanisticChar removal by oxidation

Method B:EmpiricalAssumed no charFicti t ious surface temperatureMethod C:Mechanistic

Char removal by oxidationAssumed carbon-nitrogen reactionMethod D:EmpiricalBased on correlations ( ~ s f ( in ) and = ~ ( T S ) )

Generally speaking, two ty pe s of ana lyse s were made.r e l a t i o n s which combine th e ef fe ct s of sev eral abl at ion processes.type, which can be c al le d a mechanistic procedure, attem pts t o account indiv id-ually for each process.

One type employs empiricalThe other

Method A i s a mechanistic procedure which considered diffusion-limitedoxidation as the only char removal mechanism.suff ic ie nt surface recess ion t o agree with th e f l i g h t da ta .This method could not produce

Method B could be cla ssi f ie d as empirical, mainly because a f i c t i t i o u s s u r -face temperature w a s assumed.produce t he correc t surfac e recession i n ground t e s t s a t constant heating ra te .The effect of t h i s method was t o t r e a t t h e charr ing abla tor as a high-temperature sublimer which forms no char. The s ur fa ce temperature of a char-ring abla tor responds t o var ia t io ns in heat ing ra t e . Therefore , subs tant ia lsurface-temperature va ria tio n could be expected i n th e R-4 reentry. Althought h i s method result ed i n gre ate r surface recession than di d method A, it f a i l e dt o produce enough erosion t o account f o r th e ab lat or f l i g h t performance.

A surface temperature was selected which would

Method C accounted f o r the individ ual char processes. I n order t o accom-p li s h an energy balance when ap pli ed t o ground tes ts , which poorly simulatedth e R -4 reentry, i t w a s necessary t o assume a char removal mechanism.mechanism chosen w a s a rea ctio n between the carbon i n the char and the nitrogeni n t h e stream. However, it w a s admitted that the required char removal mecha-nism could be mechanical rat he r tha n chemical. Thi s method pr ed ict ed t h a t th eablator would be consumed a t a t ime sl ig ht ly gre ate r than th at shown by thef l i g h t instrumentation.

The

Method D i s an empirical method. From ground t e s t s of t h e ab la to r made i nlow-pressure environments, a co rre la tio n between t h e surface temperature andthe net heat ing t o the surface was obtained. Then a cor re la t ion was obtainedbetween the rate of production of the gases in je cte d in to th e boundary la ye rand the surface temperature. An energy balance a t the su rface w a s then u t i l i zed36


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t o c a lc u la t e t h e char removal ra t e by assuming t h e same removal mechanisms a si n method C.th e nece ssit y of solving th e in te rn al conduction problem.R -4 reent ry, t h i s method pre dic ted t h a t the a bl at or would be consumed a fewseconds before the t i m e indicated by the f l i g h t instrumentation.

Spee ifieatio n of th e rate of production of pyrolysis gases avoidsWhen ap pl ie d t o th e

A gr ea t dea l of differe nce e xi st s between the re su lt s obtained by theseveral methods.point qui te w e l l .However, i n no ins ta nc e was the end point predicted while matching the remainderof the recession history curve.removal mechanism w a s not accounted for.

Some of t he se methods predic ted th e rec ess ion hi sto qy endOther methods matched a portion of the recess ion his tory .The implication i s t ha t the correct char

Ground T e s t ResultsArc-jet t e s t s m a d e a t Langley Research Center reveal t ha t, i n ce rt ai n high-

Figure 1 2 shows th e nature of t h i s erosio n.pressu re circumstances, t h e Apollo heat-s hield ma teri al experienced a rapid andi r re gu l a r e rosion .specimens of t h e Apollo heat-s hield mater ial te st ed a t the indicated heat ingrate and enthalpy are shown.the i n i t i a l conf igura tion . Note tha t , a s i n the f l i g h t nose cap, a molded ringi s used around th e ce nt ra l porti on which c ons ists of th e ab la tor i n honeycomb.

SectionedA t the top l e f t i s a n untested specimen showing

A t t h e t o p r i g h t i s a specimen which was su bj ec te d t o 1 atmosphere of pres-s u r e f o r 10 seconds. The recession was uniform and less than 0.1 inch. A charthickness of 0.1 inch w a s obtained. This re su lt can be accounted f o r by oxida-t ion theory .A t the bottom i s a specimen th a t was exposed t o 1.8 atmospheres of pres-sure f o r only 2.4 seconds because of excessive surface erosion. The recessiona t the cen te r w a s s l i g h t l y more than 0 .3 inch, or eight t i m e s as grea t as t h a texperienced a t 1 atmosphere. Note al so th at th e surface i s very irregular andthat the char thickness i s very s m a l l , a fact easily confirmed by examinationof the tested specimen. This was also confirmed by motion pictures which showt h a t th e specimen surface d id not ge t hot enough t o be luminous as it would i fthe char had been thick.The Langley Research Center ha s been conducting high- press ure t e s t s withth e object ives of corre la t ing the material erosion with pressure and determiningt h e phy sic al mechanism which causes erosion. Neither of the se obj ec tiv es has

as ye t been achieved. However, from ava ilab le t e s t res ul ts , it has been con-cluded th at , although aerodynamic shear has some effect , the eff

conference on langley research related to the apollo mission

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