NASA TECHNICAL NOTE

m

z c

C F I L E C O P Y

APOLLO EXPERIENCE REPORT - SPACECRAFT HEATING ENVIRONMENT A N D THERMAL PROTECTION FOR LAUNCH THROUGH THE ATMOSPHERE OF THE EARTH

by Robert L. Dotts

Manned Spacecrufi Center Houston, T e x a 770.58

N A T I O N A L AERONAUTICS A N D SPACE 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 1973

https://ntrs.nasa.gov/search.jsp?R=19730010176 2018-07-10T00:41:50+00:00Z

1. Report No.

NASA TN D- 7085

16. Abstract

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

An accurate definition of the aerothermodynamic environment of the Apollo spacecraft was neces- sary to ensure the structural 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- perature levels and thermal gradients induced in the structure of the spacecraft by the boost-phase aerothermodynamic environment could result in degradation of the structural properties of the spacecraft, degradation of the thermal-control coatings, and excessive thermal stresses in the structure. The techniques that were used to define the aerothermodynamic environment of the Apollo spacecraft during the boost phase and to predict the structural temperatures of the space- craft are discussed in this report. Also, the wind-tunnel and radiant-heating tests that were used to support the analytical predictions a re discussed; the analytical predictions are discussed; the accuracy of the boost-phase heating-analysis techniques is shown by comparing the techniques with flight data. The analytical techniques for predicting heating characteristics and structural tem - peratures of the spacecraft were adequate for predicting the temperatures of the Apollo spacecraft during the boost phase. Also, the applicability of similar techniques for future spacecraft design and analysis is indicated.

4. Title and Subtitle APOLLO EXPERIENCE REPORT SPACECRAFT HEATING ENVIRONMENT AND THERMAL PROTECTION FOR LAUNCH THROUGH THE ATMOSPHERE OF THE EARTH

7. Author(s)

Robert L. Dotts, MSC

9. Performing Organization Name and Address

Manned Spacecraft Center Houston, Texas 77058

12. Sponsoring Agency Name and Address

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

15. Supplementary Notes

17. Key Words (Suggested by AuthorIsJ)

5. Report Date

6. 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 Covered

Technical Note 14. Sponsoring Agency Code

Launch Vehicle Configurations Aerothermodynamics Thermal Protection Aerodynamic Heating

19. Security Classif. (of this report)

None

18. Distribution Statement

20. Security Classif. (of this page) 21. No. of Pages 22. Price

None 23 $3.00

APOLLO EXPERIENCE REPORT

SPACECRAFT HEAT1 NG ENVl RONMENT AND THERMAL

PROTECTION FOR LAUNCH THROUGH THE

ATMOSPHERE OF THE EARTH

By Robert L. Dotts Manned Spacecraft Center

SUMMARY

An accurate definition of the boost-phase aerothermodynamic environment of the Apollo spacecraft was required for the design of a thermal protection system that would ensure the structural integrity of the spacecraft during the boost phase of an Apollo mission. Without adequate protection, the temperatures and the temperature gradients induced in the spacecraft structure by the boost -phase aerothermodynamic environment would result in an unacceptable degradation of the spacecraft structural properties and thermal-control coatings, significant thermal stresses in the structure, and excessive temperatures in the pyrotechnic charges attached to the structure.

The total Apollo boost-phase thermal protection system of the spacecraft is dis - cussed in this report, and the techniques that were used to predict the boost-phase aerodynamic-heating environment of the spacecraft are discussed briefly. In general, conservative design approaches were used; however, localized failures of several lunar module adapter panels during the Apollo 6 (AS-502) mission caused an extensive reeval- uation of the component that resulted in thermal redesign. Extensive analyses and ground tests were performed on this component during the investigation of the AS-502 anomaly. The results of the analyses and tests are discussed in detail. In addition, analytical predictions of structural thermal response to the boost -phase environment are compared with flight data.

I NTRODUCTI ON

The Apollo spacecraft is subjected to substantial boost -phase aerodynamic heating during launch through the atmosphere of the earth. An accurate definition of the effects of this boost-phase aerodynamic -heating environment on the Apollo spacecraft w a s required to design a thermal protection system that would ensure structural integrity of the spacecraft and subsequent operation of all spacecraft systems and components affected by this thermal environment.

The design of the boost-phase ther- mal protection system of the Apollo space- craft involved definition of the boost-phase aerodynamic-heating environment; ther- mal analysis of the various spacecraft ele- ments during the boost phase to determine the temperature histories of the elements; design of thermal protection systems for the spacecraft elements that would expe- rience excessive temperatures during the boost phase; and ground tests of the vari- ous elements of the thermal protection system and flight tests of the integrated thermal protection system to verify the adequacy of the design.

The general boost-phase design con- siderations of the thermal protection sys- tem for the primary Apollo spacecraft components are discussed in this report. Also, the techniques that were used to define the aerodynamic-heating environ- ment are discussed; the basic thermal- analysis techniques and the ground and flight tests that were used to verify the adequacy of the design of the thermal pro- tection system a r e presented.

SPACECRAFT BOOST-PHASE THERMAL- PROTECTION-SYSTEM CONFIGURATION

AND DES I GN CONS I DERATl ONS

The primary spacecraft elements that require boost-phase thermal protec- tion are shown in figure 1. The elements include the launch escape system (LES); the boost protective cover (BPC), which covers the command module (CM) during launch; the service module (SM); and the spacecraft/lunar module adapter (SLA), which covers the lunar module.

-Launch escape system

Command module

Figure 1. - Spacecraft launch configuration.

2

Launch Escape System The purpose of the LES is to remove the CM rapidly from the vicinity of the

booster during a boost-phase abort. The LES (fig. 2) consists of a solid-propellant- rocket-motor assembly that contains three rocket motors attached to the forward por- tion of the CM by a tower structure. The CM BPC is attached to the LES tower structure and provides thermal protection for the CM during the boost phase. The entire LES, including the BPC, is jettisoned near completion of a normal boost phase o r an abort.

The legs and crossmembers of the LES tower were covered with Buna-N rubber insulation that was sized to res t r ic t the maximum temperature of the horizontal and vertical titanium leg members to 600" F and the diagonal leg members to 800" F dur- ing the design boost -abort phase. The thermal response of a typical vertical leg mem - ber (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 material was used to protect the tower-jettison motor, part of the LES motor, the power-systems and instrumentation wire harness, and the structural skirt from the boost -phase/abort thermal environment. A typical thermal response for the LES motor casing (location B in fig. 2) during the AS-503 design boost-phase/abort environment (ref. 1) is shown in figure 4.

Power-systems and instrumentation wire harness

insulation

Figure 2. - Launch escape system.

3

I I

I I

I I I I , I /

Command Module

I I I I

1 , I , I 1 I 1-

The BPC prevents the CM external thermal-control coating from exceeding 250" F and provides protection from coating contamination during the boost phase. Also, the BPC protects the CM from the high-temperature exhaust of the LES motor during a boost-phase,abort. A side view of the BPC is shown in figure 5. The BPC is composed of two sections: a hard cover that is attached to the LES tower legs and a soft cover that is connected to the trailing edge of the hard cover. fiber -glass/honeycomb sandwich that is covered with cork thermal-protection material (0.3 inch thick). A typical thermal response for the BPC during the AS-503 design boost-phase/abort environment is shown in figure 6. The temperatures shown are for

The hard cover is a phenolic

Fiber-glass ring Honeycomb-cored laminated-

Hard-cover seal 7r f iber-glass panel r L a m i n a t e d

LM near-snieia

0.030-in-thick corkboard ablator fiber-glasslhoneycomb sandwich

-ES jettison ;at 312000 fl I

;at 24000 fl

50 I ! I

1 I I I l l I 1 1 1 1 . 1 0 50 100 150 200 250 300 350 400 450 500 5% 6W

I I

I

I Elapse t ime, sec

4

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

the outer and inner facesheets of the hard-cover fiber -glass/honeycomb sandwich (ref. 1). The soft cover consists of a 0.008-inch-thick Teflon-impregnated glass cloth that is reinforced on the outer surface with a 0.0095 -inch-thick nylon fabric. Cork thermal-protection material is bonded to the outer surface of the nylon fabric.

Service Module

The SM (fig. 7) is a cylindrical unit that is composed of an outer shell and an inner concentric core. The resulting annulus is divided into six bays by radial beams. The outer shell is fabricated of 1-inch-thick aluminum-honeycomb panels and is cov- ered with cork thermal-protection material (0.020 to 0.155 inch thick), except in the regions where the environmental control system (ECS) and the electrical power system (EPS) radiators are located. The cork provides protection from boost heating and SM

Service propulsion

Figure 7. - Service module.

reaction control system (RCS) plume- impingement heating. The cork thickness was sized to maintain the SM honeycomb structure below the design temperature limit of 400" F during boost. The four RCS modules, which protqude from the surface of the SM shell approximately 10 inches, cause flow disturbances that result in increased boost-heating rates in the vicin- ity of the modules. on a representative portion of the SM sur - 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 of

the SM. defined in te rms of angular position meas- ured in degrees from the +Y axis. The thicknesses result from designing for the heating variation over the surface (ref. 1) and from maintaining a maximum tempera- ture of 400" F at all surface points while minimizing the cork weight,

The cork thicknesses

The position X = 200 inches S

The circumferential position is

5

0.155

0.06

0 0.03 0.03 o ~ o ~ ~

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

Radial I beam 1 +Y axis 10'1

... ........... ....... ........... ........ ........... .............. ........... ........ ..................... .................... Bare surface Note Al l dimensions are in inches ......................

0.02

0.04 0.02 \ 0.03

umbi l ical Q- i I 0.06

0.155

I 0.06

l"h"i: n e! 0.03

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

Radial beam 2

I R + Z axis l9O"l beam

~ = 376

= 355

* = 200

I

Figure 8. - Block 11 SM (AS-503 and subsequent spacecraft) cork thermal- protection thickness.

6

Spacecr aft/L u n ar Mod u I e Adapter

The SLA structure is a truncated conical shell (fig. 9) that is fabricated of 1. 7-inch -thick aluminum -honeycomb panels with an outer cover of cork thermal- protection material (0.030 to 0.20 inch thick). Numerous items, such as hinges, an- tennas, lights, access doors, and structural joints, protrude from the surface of the SLA (fig. l o ) , causing flow disturbances and localized areas of increased aerodynamic heating. The items were protected by cork o r fabricated of material capable of with- standing the localized temperature extremes (ref. 1). The longitudinal position on the SLA is defined, in inches, by X

end of the SLA; XA = 838 inches corresponds to the top of the SLA. The circumferential

position is defined in terms of angular position measured in degrees from the +Y axis.

The position X - 502 inches corresponds to the lower A' A -

Initially, the cork thickness was sized to maintain the SLA honeycomb structure below the design temperature limit of 490" F. The 0,030-inch-thick cork (shaded su r - faces in fig. 10) was added to Apollo 8 (AS-503) and subsequent adapters because of the AS-502 anomaly. The AS-502 SLA had a localized structural failure of several panels during the boost phase; as a result of the subsequent investigation, the 0.030-inch-thick cork was added to the bare surfaces of the SLA to reduce the temperatures and tempera- ture gradients in the SLA structure.

X A = 838 in.

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

7

I + Y axis tool

Note All dimensions are i n inches + Z axis (90")

-0.03-in:thick cork on AS-503 and subsequent spacecraft

Figure 10. - Cork thermal-protection thickness on the SLA.

8

THE S M AND SLA BOOST-PHASE AERODYNAMI C-HEAT1 NG ENV I RONMENT

A boost-phase aerodynamic-heating computer program that was originally devel- oped and used for Gemini spacecraft boost-heating analysis was used in defining the Apollo spacecraft boost -phase aerodynamic -heating environment. gram characterized the flow of air over the vehicle with a normal shock at the nose and an isentropic expansion to local conditions at each specific body station. Eckert's reference-enthalpy method (ref. 2) w a s used in the computer program with the Blasius skin-friction coefficient (ref. 3) for laminar flow and the Schultz-Grunow skin-friction coefficient (ref. 4) for turbulent flow. The local pressure P was determined by using

the ratio PL/PT and the total pressure P behind the normal shock. The local ther- T modynamic state point was 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, Reynolds L' number, vehicle angle of attack, and local body station. Wind-tunnel tests on scale models of the Apollo spacecraft launch configuration were performed to determine the pressure distribution over the vehicle surface for various Mach numbers and Reynolds numbers at 0 angle of attack. Angle -of -attack effects on SM and SLA heating were negligible because of the small angle of attack (nominally less than 1") for the Apollo vehicle during the heating portion of the nominal boost-phase trajectory. The wind- tunnel model and the pressure-measurement locations a r e shown in figure 11. Curve fits of the pressure-ratio data for the SM and the SLA are shown in figure 12. Varia- tion of PL/PT with Reynolds number was small; therefore, the data could be corre- lated uniquely with the Mach number. For a given vehicle trajectory and SM o r SLA location, time histories for smooth-body heating rate were generated for a range of surface temperatures. The heating rates were used in a thermal-analysis computer program as a function of time and vehicle-surface temperature to calculate transient temperatures at various locations on the SM and the SLA. body heating rate for a typical AS-202 SLA location are shown in figure 13. The spe- cific temperatures used to generate the heating-rate curves of figure 13 were chosen to bracket the expected SLA surface temperatures.

Time histories for smooth-

Flow direction

Pressure-sensor locations

L-I -

Centerline 3- ___

Figure 11. - Wind-tunnel-model configuration and pressure -measurement locations; the model is 0.045 scale.

9

1.0-

.a

. 6

-

-

cL+ . 4 - \ A C L

d 3 ", E . 2 -

5

m c 0 .- -

. l o -

e .M-

E .M- z

s = .A

0 - - .M- - 0 0

= ._ -

.02 -

.01 I I I I

Mach number 0 2 d 6 a

Figure 12. - Ratio of local pressure to stag- nation pressure pL/pT as a function of

Mach number for the SM and SLA.

The increased heating that was caused by the various flow disturbances on the su r - faces of the SM and the SLA was accounted for by using protuberance factors (ratios of the smooth-body heating rate to the local heating rate). The factors 'were applied as constant multipliers to the smooth-body

Local syface temperature = 40' F

L // Local surface

, temperature i 540" F

1040' F

I ; I 1 1 1 1 I I J

' u 20 40 60 80 100 120 140 160 180 200 Elapsed lime. sec

Figure 13. - Time histories for smooth- body heating rate for the AS-202 SLA (station XA = 775 in.).

heating rates throughout the boost-phase trajectory. This conservative approach of using constant factors rather than varying the factors with Mach number was used to ensure adequate design safety margins for the protuberance-affected areas of the space- craft. Subsequent Apollo flight data indicated that this approach was overly conserv- ative; therefore, Mach-number-varying protuberance factors were used later to verify the adequacy of the SM and the SLA for off-nominal trajectories. The predicted protuberance-factor contours that result from the RCS modules and other protuberances on the surface of the SM are shown in figure 14 (ref. 5). The predicted protuberance- factor contours for two sections of the SLA surface are shown in figures 15 and 16 (ref. 5). The protuberance-factor contours were estimated by using the test data ob- tained from reference 6 and wind-tunnel data for the Apollo spacecraft configuration.

10

I'

X s - 316 ir

Xs = 355 i r

xs = 200 I

Rad[, I

Radial beam 2 I I

+.? axis Radial beam 3 I

earn 1 +Y axis

Figure 14. - Typical interference -heating map of protuberance -factor contours for a section of the SM.

11

Radial locations : 0", 90", 180", 270"

Lines of constant protuberance factor

SLA hinge

-.._

@ - 2.0

I

1

-Cableway

Thruster cover

XA = 583 in. --I- --

1.0 Protuberance factor

-Lines of con stan t protuber- ance factor

- 502 in. fl - *A

Cent& l i n e

Figure 15. - Local interference -heating map of protuberance -factor contours for SLA hinges, thruster cover., and cableway system.

12

I I I

I

I

I 1

-L z

. XA = 583 i n

-Lines of constant Protuberance factor

Scimitar antenna -

I

i Protuberance factor

* 502 in . fi - ’A

I Radial locations * 45”. 135’, 21y, 31.50 I

Radial locations - 45: 225’

Figure 16. - Local interference -heating map of protuberance -factor contours for SLA/LM attachment cover and scimitar antenna.

BOOST-PHASE THERMAL RESPONSE OF THE SM AND SLA STRUCTURES

Thermal-Mathematical Models and Assumptions

The thermal response of the structure was determined by using the surface- heating-rate histories as boundary conditions for one - and two-dimensional thermal- mathematical models of the structure. Most of the SM and SLA structures are comprised of aluminum -honeycomb -sandwich panels that can be analyzed adequately by using a one -dimensional model. The thermophysical properties of the honeycomb panels were obtained from reference 7. A typical one-dimensional thermal model is shown in figure 17. One-dimensional mathematical models, which a r e similar to the model shown in figure 17, and two-dimensional mathematical models of more complex SM and SLA structural elements (joints, bulkheads, et cetera) were used for all tem- perature predictions. Extensive radiant -heating tests were performed to ensure that the one- and two-dimensional mathematical models were adequate. Also, the tests were used to determine the performance of the cork thermal-protection material that was used extensively on the surfaces of the SM and the SLA.

13

Thermal- mathematical- model nodes-

300 280 260 240 220

-- 180 e m - c 5160- 5 140 E120 e 100

80 60- 40-

Aerodynamic heat flow Radiative heat flow

(+&cork

- - - - -

- - - - -

L A d i a b a t i c surface

Figure 17. - One -dimensional model of an aluminum -honeycomb panel.

GrGu nd Testi ng

One -dimensional SLA honeycomb - panel thermal tests. - Thermal tests were performed on 12 - by 12 -inch, 1 . 7 -inch - thick SLA honeycomb panels to verify the analytical techniques that were used to predict temperatures. Radiant-heating lamps were used to simulate the boost- phase heating history. Lamp voltages were programed so that the bare-surface (with- out cork) honeycomb outer -facesheet test temperatures matched the analytically pre - dicted temperature history for the AS-503 and subsequent -spacecraft design trajec-

tory. The bare-surface SLA honeycomb-panel thermal response for the AS1503 sim- ulation is shown in figure 18. The inner -honeycomb-surface analytical predictions that resulted from using a one-dimensional thermal model were slightly higher than the measured temperatures. Cork-covered honeycomb panels were subjected to the AS-503 radiant-heating history to assess the cork performance and to determine the accuracy of the analytical predictions. The predicted and measured thermal response for a panel covered with 0.030 -inch-thick cork and subjected to the AS-503 radiant-heating history a r e shown in figure 19. The analytical results that were obtained by using thermophysical-property data for virgin cork agreed well with the test data. The pre- dicted and measured responses for a panel covered with 0.050-inch-thick cork and sub- jected to the AS-503 heating history a r e shown in figure 20. the adequacy of the analytical temperature -prediction techniques for both bare-surface and cork -covered honeycomb panels in typical Apollo launch environments.

The tests.demonstrated

- Analysis 0 Test data

r Outer facesheet

0 P o

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

Figure 18. - Thermal response of a bare-surface SLA honeycomb panel to an AS-503 trajectory simulation.

240

- Analysis 0 Test data r Outer facesheet

20 E-zL- 0 20 40 60 80 100 120 140 160 180 200 220

Elapsed time, sec

Figure 19. - Thermal response of a cork-covered 0.030-inch-thick SLA honeycomb panel to an AS-503 trajectory simulation.

I 14

- Analys is 0 Test data

240 r o u t e r facesheet

5 120 + loo

I n n e r facesheet 80 t c-kwj ab ;o 4b Qo io lbo 1; 140 1; IS0 2& 2k

Elapsed t ime. sec

Figure 20. - Thermal response of a cork-covered 0.050-inch-thick SLA honeycomb panel to an AS-503 trajectory simulation.

Two -dimensional SLA-joint thermal tests. - Nine test samples were cut from a structural-test SLA at the locations shown in figure 21. perform radiant-heating tests to assess the temperature gradients in the SLA at the end of first-stage boost and to verify the ade- quacy of the analytical-prediction tech- niques. Each panel was subjected to the AS-503 radiant-heating profile that was used in the one -dimensional radiant -heating tests; however, only the data that were obtained from test joint 3 will be discussed in this report. Cross-sectional and top views of the test joint are shown in fig- ures 22 and 23; the thermocouple locations also are shown in figure 23.

- The samples were used to

The joint was configured with the AS-501 cork pattern (a 0 . 1 -inch-thick layer over only the pyrotechnic joint) and subjected to the heating history of the AS-503 design trajectory. The thermocouple temperatures for the joint after an elapsed time of 150 seconds (the approximate termination of first-stage boost) a r e shown in figure 24. The two-dimensional thermal-mathematical model (fig. 25) was used to predict the joint temperatures for the same heating profile. The analytical predictions also are shown in figure 24. plexity of the joint and the uncertainties in the thermal-mathematical model (thermo- physical properties, contact resistances, and so forth).

.

The correlation is very good, if consideration is given to the com-

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

15

Aluminum- Aluminum- honeycomb honeycomb core ( lk - in . core ( W i n .

Figure 22, - Cross section of test joint 3 (longitudinal-separation joint .between stations X = 710 in. and X = 750 in.). A A

ate: Thermocouples 1 to 24 are located on the outer faces heet. Thermocouples 25 to 46 are located on the i n n e r facesheet.

0 0 0 0 . 0

1 2 3 4 5 6

25 26 27 28 29 M

-

0

7

31

0

8

47

-

/-Bolt heads-

0

0 0

9

32

0

0

0

0

11

35

0

10 33

. 12

34

0

15

37

0

13

36

0

14

38

0

0 0

16

39

0

0

0 3 in.

0

17

48 . 18 40

0 . 0 . . . 19 x) 21 22 23 24

41 42 43 44 45 46

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

16

340 320 300- 280 2* 240 220

," 200- 180-

B 140- 5 120-

loo 80 60- 40- 20-

c

I - Analys i s 0 Outer- facesheet data I 0 I nner- facesheet data

- - - b -1 -

- -

Distance f r o m center of j o i n t , in.

Figure 24. - Temperature distribution for test joint 3 covered with an AS-501 cork pattern and subjected to an AS-503 trajectory simulation.

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

Inner facesheet Note All dimensions are in inches

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

17

A 0.030 -inch -thick cork layer was bonded to the remaining bare surfaces of the panel, which corresponded to a pro- posed AS-503 SLA cork configuration. The radiant -heating simulation was again per - formed on the joint to a s ses s the effect of the added 0.030-inch-thick cork on the thermal gradient in the joint. The test results for the configuration are shown in figure 26. The addition of the 0.030-inch- thick cork layer reduced the gradient in the outer facesheet from 140" to 90" F. The results of the two-dimensional SLA- joint thermal tests were used to define the temperatures and temperature gradients in the SLA structure at the termination of first-stage boost. The data were used as inputs for analyses of structural and ther- mal s t ress , which were performed to assess the structural adequacy of the SLA for the AS-503 and subsequent spacecraft.

0.1 -in: t h ic k 0.03-in-thick cork- Cent!rline r c o r k

340 320

20

.. I

0 Outer-facesheet data I 0 inner-facesheet data

o l k o : I ; k ; 1 I 1 I6 ; 1 ;: k d 'I ; 6 ;0;1 Distance from center of joint, in.

Figure 26. - Temperature distribution for test joint 3 covered with an AS-503 cork pattern and subjected to an AS-503 trajectory simulation.

Fl igh t Data fo r t h e SM and SLA as Compared w i t h Analyt ical Predic t ions

Spacecraft AS-202. - The temperature history for SM sensor SA7916T is shown in figure 2 ' l . The analytical predictions, which were based on the actual launch traiec-

7 Maximum predicted response 260 r I M i n i m u m predicted response

L L 7

40 80 120 160 200 240 280 3M Elapsed time, sec

Figure 27. - Temperature history for SM sensor SA7916T (outer su r - face, X - 350 in. , and a circum- ferential position of 253") during the launch of spacecraft AS-202.

S -

tory, a r e slightly conservative. mum predicted response is based on full solar exposure on the panel and radiation cooling to the earth; the minimum predicted response is based on no solar exposure or radiation cooling to space. The maxi- mum and minimum predictions should bracket the expected response. The tem- perature history and predicted responses for SLA sensor AA7931T are shown in fig- ure 28. The spike in the measured data at an elapsed time of 145 seconds was caused by the heating from the forward-firing retrorockets of the first stage at separa- tion; this heating was not accounted for in the analysis. analytical predictions for SLA inner -surface sensor AA7932T are shown in figure 29. As observed in the radiant -heating tests, the analytical predictions were higher than the actual temperatures encountered on the SLA

The maxi-

The temperature history and

18

M i n i m u m predicted r

200 290 Measured-data band ax imum predicted response

160 ~ 250 urn predicted respons

L L

LL 210 g 120

e c

+ 130 M i n i m u m predicted response 40

2 c

l 5 170 80 Measured-data band 0) CL a I-

5 I

I

Note Spike in data caused by

forward-f ir ing retrorockets at separation

I 0 40 80 120 160 Mo 240 280 320 Elapsed t ime. sec

90 heating from f i rs t -stage

I I I I I I I I 0 40 80 120 160 200 240 280 320

501

Elapsed time, sec Figure 29. - Temperature history for SLA sensor AA7932T (inner su r - face, X = 775 in., and a circum-

ferential position of 124 ") during the launch of spacecraft AS-202.

~

A Figure 28. - Temperature history for SLA sensor AA7931 T (outer sur - face, XA = 775 in. , and a circum-

ferential position of 124 ") during the launch of spacecraft AS-202.

inner surface, The higher predictions Maximum predicted

probably were caused by the adiabatic- inner -surface assumption that was made for all the analytical predictions.

Spacecraft AS-501. - The majority of the SM surface was covered with cork for the Apollo 4 mission, and only the inner surface of the SM w a s instrumented with thermocouples. As a result, the maximum measured temperature for the AS-501 SM during boost was 90" F. The temperature history and analytical predictions for SLA sensor AA7864T a r e shown in figure 30.

launch phase, the correlation is good. A review of atmospheric balloon data from the NASA John F. Kennedy Space Center indi- cated that the atmospheric temperature up to approximately 50 000 feet (0 to 80 sec- onds) was considerably colder than was indicated by the 1962 standard atmosphere (ref. 8), which was used in the analysis. The unusually cold atmospheric conditions resulted in an initial cooling after lift-off.

Measured-data band

Except for the first 80 seconds of the Elapsed lime. sec

Figure 30. - Temperature history for SLA sensor AA7864T (outer su r - face, x = 730 in. , and a circum- ferential position of 174") during the launch of spacecraft AS-501.

A

19

CONCLUDING REMARKS

The analytical capabilities for predicting spacecraft heating characteristics and structural temperatures were adequate for prediction of the Apollo spacecraft temper - atures during boost. Given the pressure distribution over the surface of the vehicle, the analytical techniques provided reasonably accurate predictions of the heating rates and thermal responses of the spacecraft during boost. having local pressure disturbances, the utilized techniques had dubious value. Pressure disturbances, caused by external hinges, umbilicals, RCS modules, e t cetera, result in localized a reas of higher heating. For the SM and SLA, the problem of protuberance- affected areas was solved by applying an experimentally determined heating-rate factor to the smooth-body heating rates. cedure were overly conservative for the Apollo -type protuberances.

For areas of the spacecraft

The analytical predictions resulting from this pro-

In the future, to reduce the thermal-protection weight penalty associated with overconservatism, attention should be given to obtaining better techniques for predict- ing the effect of protuberance heating on spacecraft. The use of protuberance factors that are functions of Mach number should eliminate much of the overconservatism. Although the techniques that were used for design of the boost thermal protection sys- tem for the Apollo spacecraft were conservative (because no structural element o r component exceeded the design temperature), the Apollo 6 (AS-502) anomaly led to increased thermal protection of the SLA. structural margins and to reduce thermal stresses below the originally defined require- ments. The analytical-prediction and testing techniques that were developed for the Apollo Program should enable the thermal analyst to design future spacecraft with less severe thermal gradients in the structure during boost and with lighter weight boost thermal protection systems.

This addition was made to increase the

Manned Spacecraft Center National Aeronautics and Space Administration

Houston, Texas, October 26, 1972 914-50-20-12 -72

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REFERENCES

1. Anon. : Thermal Data Apollo Block II Spacecraft. Vol. I, Structural Thermal Response, 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 mit 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 Heating Environment, 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 Pressure Measurement on a Flat -Plate Surface and Heat -Transfer Meas - urements on Attacked Protuberances in a Supersonic Turbulent Boundary Layer at Mach Numbers of 2.65, 3.51, and 4.44. NASA TN D-1372, 1962.

7. Anon. : Evaluation of the Thermal Properties of Materials. Vol. 11, Data Handbook Final 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.

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