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NASA CONTRACTOR NASA , " REPORT

'LOAN COPY: RETURN TO N I AFWL TECHNICAL LIBRARY PC: KIRTLAND AFB, U N. M.

*& ABLATIVEHEAT SHIELD DESIGN

Rolf W. Seijerth

Prepared by "MARTINMARIETTA COR?- E,U !'e:? Lvg Denver, Colo. 80201 > for LangleyResearch Center

NATIONALAERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. ECTM 4975 5.

I. TECH LIBRARY KAFB, NU

OObL287 ". 1. Report No. 3. Recipient'sCatalog No. NASA CR-2 57 9 - 4. Title and Subtitle II 5. ReportDate Ablative Heat Shield Design for Space Shuttle

- .. ____~ 7. Author(s) 8. PerformingOrganization Report No. R. W. Seiferth MCR-73-147 -

" . 10.Work Unit No. 9. Performing Organization Name and Address

Martin Marietta Corporation 11.Contract or Grant No. Denver Division Denver,Colorado 80201 NAS1-11592 13.Type of Report andPeriod Covered Executive Summary 2. SponsoringAgency Name and Address May 1972 - August1973 NationalAeronautics and Space Administration 14. SponsoringAgency Code Washington D.C., 20546

5. SupplementaryNotes

Details of thisExecutive Summary are inFinal Report, CR-132282 (Same Title) FINAL REPORT 6. Abstract

Ablatorheat shield configuration optimization studies were conductedfor the Orbiter.Ablator and RSI trajectoriesfor design studies were shapedto take advantage of the low conductance of ceramic RSI andhigh temperature capability of ablators.Comparative weights were establishedfor the RSI systemand for direct bond and mechanically attached ablator systems. Ablator system costs were determinedfor fabrication, installation and refurbishment. Cost penalties were assignedfor payload weightpenalties, if any. The direct bond ablator is lowestin weight and cost. A mechanically attachedablator using a magnesium subpanel is highlycompetitive for both weight and cost.

17. Key Words(Suggested byAuthor(s) ) 18.Distribution Statement Sp ace Shuttle CoatingsShuttle Space HeatAnalysesCost Shield Unclassified - Unlimited A blators Trajectory ShapingTrajectory Ablators Design Criteria ReusableSurface Environmenta1.TestingInsulation 1 SubjectCategory 33 19. Security Classif. (of this report)22. 20. Security Classif. (of this page) 21. No. of Pages Price' Unclassified Unclassified 28 $3.75

* For sale by the National Technical Information Service, Springfield, Virginia 22151 I

FOREWORD

This report was preparedby Martin Mari'etta Corporation, DenverDivision, Denver, Colorado, under Contract NAS1-11592, "Ablative Heat ShieldDesign for Space Shuttle."

Theprogram.was performed by members ofthe AbiRtive Systems Section,the Advanced Structures and .Materials Dc:r~;.;:"ment: the MechanicalDesign Engineering Department and the Systems Analysis Department.The program manager for Martin Marietta fromthe start of thecontract on May 17, 1972 toSeptember 5,'1972 was Mr. Daniel V. Sallis; Mr. Rolf W. Seiferth was programmanager fromSeptember 5, 1972 throughAugust 31, ,1973.

This finalreport covers the contract period from May 17, 1972through August 31, 1973.

Many personscontributed during the performance of thepro- gramand inthe preparation of this report. Martin Marietta Aerospace

D. V. Sallis,Program Managerthrough September 5, 1972

H. H. Chandler,Technical Director through September 5, 1972

W. R. Woodis,Trajectory Analysis

D. A. Schmitt,Aerothermal.Ana1ysis

W. D. Schwerin,Aerothermal Analysis

R. Stevenson,Aerothermal Analysis

D. G. Herbener,Thermal Analysis

iii S. C. Copsey, Thermal Analysis

E. T. Haire, Design Criteria

G. F. Molby, Stress Analysis

T. L. Hibbard, Mechanical Design

A. P. Lambatos, Weights Analysis

J. M. Toth, Jr., Design Verification

H. L. Draper, Design Verification

J, W. Maccalous, Design Verification

R, V. Kromrey, Design Verification

G, T. Kelly, Cost Analysis

W. L. Brown, Report Illustrator

C. F. Cancalosi, Schedule Coordination

iv TABLEOF CONTENTS

Page

Foreword ...... iii

Contents ...... V

Summary ...... 1 I INTRODUCTION ......

I1 DESIGNCRITERIA (TASK 1) ...... Factors of Safety ...... Backface(Structural) Temperature Limits . . AblatorStrain Limits ...... StrengthAnalysis ...... I11 FLIGHTENVIRONMENT (TASK 2) ...... 4 TrajectoryShaping ...... 5 Heating Rate Distribution ...... 5

IV HEAT SHIELDDESIGN (TASK 3) ...... 8 ThermalAnalysis ...... 9 StressAnalysis ...... 11 Total Heat ShieldConfiguration Weights . . 13

V DESIGNVERIFICATION (TASK 4) ...... 13 OpenGapTest ...... 14 Gap Sealer Tests ...... 14 Sealed Gap Tests ...... 14

VI WEIGHT AND COSTANALYSIS (TASK 5) ...... 15 PayloadWeight Penalties ...... 15 OperationalCosts. TPS ...... 16 TotalProgram Costs ...... 17

VI1 DISCUSSIONPROGRAM OF RESULTS ...... 20

VI11 CONCLUSIONS AND RECOMMENDATIONS ...... 21 Conclusions ...... 21 Recommendations ...... 22

REFERENCES ...... 23

V Figure

1 TrajectoryShaping ...... 6 2 EntryLocal Static PressureHistory at Body Lower CenterlineReference Location ...... 6 3 EntryHeating Rate .Ablator Trajectory (Without Factorsof Safety) ...... 7 4 EntryHeating Rate .RSI Trajectory(Without Factorsof Safety) ...... 7 5 AttachmentConfigurations Investigated ..... 8 6 SLA-561 AblatorDesign Chart ...... 10 7 TPS AttachmentConfiguration Weight vs Fastener Spacing ...... 12 8 Relative OperationalCosts of a TypicalThermal ProtectionSystem. 12.7 cm (5 in.) Aluminum Plate Subpanel ...... 18

Tab le

1 FactorsSafetyof ...... 3 2 ThermalCharacteristics Nonablatorof Components ...... 10 3 DesignLoading Conditions ...... 11 4 SubpanelGages Required for Direct Attachment Application ...... 12 5 Verification Test Program ...... 13 6 ThreeMissions ofthe 1981 Period ...... 16 7 OperationalCosts ...... 18 8 TotalProgram Costs ...... 19

vi EXECUTIVE SULW\KY REPORT

ABLATIVEHEAT SHIELD DESIGN FOB SPACE SHUTTLE

By Rolf W. Seiferth MartinMarietta Corporation DenverDivision

SUMMARY

State-of-the-artablative materials were usedto design a thermalprotection system (TPS) forthe Space Shuttle Orbiter. An "ablatortrajectory" was developedwithin L11e boundsof 2.5-g acceleration and 300 kW/m' (26 Btu/ft'-sec) heating rate at the referencepoint 15.24 m (50ft) aft of the fusr2lag.e nose 0x1 the bottomcenterline. An "RSI trajectory" was also developedfor designcomparison purposes. This trajectory was shaped to mini- mizeheating rate within the limits ofskipout during reentry. Heating rates and total heats were developedfor the total Orbiter. Ablativeheat shield designs were derivedfor numerous locations on theOrbiter using direct bond and mechanically attached concepts. A reusablesurface insulation (RSI) TPS was also developed forweight comparison purposes. Radiant heat tests were conductedon mechanically attached ablator specimens to verify designconcepts.

A costanalysis was preparedfor the variol~s heat shie1.d concepts.Weight was considered as a costfactor by determining a costper pound to orbitusing the "Preliminary Traffic Model for the Space Shuttle,"(Shuttle Utilization Planning.Office, NASA-MSFC, November 9, 1972). Ablator TPS operation was assumed forthe first fiveyears of Shuttleservice. Cost data werederived for the operational phase and for reliability, which was treated as a qualityassurance item. The sum ofthe weight costs, operational costs, and reliability costs was used to rate the various heat shield concepts and select this optimum ablator configuration.

The direct bond ablatorsystem had Che lowesl: weigh^ znd programcost of all thesystems examined. Xechanically attached plateswith ablator bonded to them are verycompetitive for bcth weightand cost.

A detailed description of this work is given in NASA CR- 132282,"Ablative Heat Shield Design for Space Shuttle," by Rolf W. Seiferth. I. INTRODUCTION

Ablators are a well-establishedsystem of thermal protection,. havingbeen used on such vehicles as Apollo,Gemini, Viking space- craft, X-15, Titan, PRIME, andothers. The need for ablator re- furbishmentfollowing each flight is a drawbackof thisthermal protectionsystem (TPS)and hasled NASA andindustry into the developmentof reusable surface insulation (RSI) ceramics. The RSI systemof thermal protection has been baselined for use on theShuttle Orbiter.

Much workneeds to be done to qualify the RSI forSpace Shuttleapplication and, to quote E. S. Love (ref. 1) fromthe Tenth Von Karman Lecture,"Ablators offer a confidentfall-back solution(temporary) for both leading edges and large surface areas, shoulddevelopment of the baseline approaches lag."

Inthe past, ablator systems have been bonded directly onto thestructures they are designedto protect. While thisapproach is both low inweight and costeffective, it hasthe drawback for theShuttle Orbiter of taking up criticalturnaround time for re- furbishmentbetween flights, and, during refurbishment, creates a problemof debris and dust control.

Thisprogram investigated Shuttle Orbiter TPS designconcepts usingavailable state-of-the-art ablators. An endobjective of theprogram was to obtain ablator TPS weightand cost estimates based on detailed,verified heat shield designs. RSI cost estimates were not part ofthis program. Direct bond ablator and RSI designs were preparedfor weight comparison purposes. A keypart of theeffort dealt with methods of mechanically attaching pre- paredablator panels onto the Orbiter. Radiant heat tests were conductedto verify the design concepts.

Theprogram was dividedinto five tasks:

Task 1 - Design Criteria;

Task 2 - FlightEnvironment;

Task 3 - Heat ShieldDesigns;

Task 4 - DesignVerification;

Task 5 - Weightand Cost Analysis.

2 11. DESIGN CRITERIA (TASK 1)

Design criteria were preparedto develop valid heat shield configurationsfor the total environment from prelaunch through reentryand landing. Criteria were preparedfor the trajectory definition andfor the thermal and stress analysis.

Factorsof Safety

Safetyfactors on loadsand pressures for prelaunch through deorbit are basedon engineering practice developed for boosters andspacecraft and for entry and atmospheric flight. The safety factors are those commonly usedin the design of aircraft (table

TABLE 1.- FACTORSOF SAFETY

Ultimate Ultimate Limit load Location design factorof condition heatingload (..safety design) times ( or nominal. ) Total Orbiter Through orbit 1.4 Limit load

~~ Total Orbiter Limit load Entry I 1.5 Leadingedges and Ascentand 1.15 Nominal lower forward entry heating surfaces load

Lower surfaces aft

Upper surfaces I 1.50

Backface(Structural) Temperature Limits

The backfacetemperature limits are asfollows:

Ablatorlinebond (500OF)533°K

Maximum structural temperatures At start ofentry 311°K (100°F) At completionof entry 450°K (350OF)

3 Ablator Strain Limits

Ablators have characteristically lowmechanical properties. Althoughvalues for charred materials are difficult to ascertain, thefollowing room temperatureproperties are typical:

Strength 480 kN/m2 (70 psi)

Modulus 20 700 kN/m2 (3000 psi)

Inaddition, the following induced strain limits havebeen developed for the Viking Project :

V irg in CharredVirgin

T e nsio n strain Tension 1.0% 0.6%

Co m pre ssion strain,Compression 1.O% 1.0%

StrengthAnalysis

The ablation material is notconsidered load carrying, but will beincluded in thermal and mechanical deflection analysis to determinethe strain in the ablation material. Thesubpanel will becapable of carrying design loads without the ablator, and with- outexceeding the following surface waviness deflection criteria:

H = 0.0125L limit

L = panel wave length

H = maximum deflection (wave height)

111. FLIGHT ENVIRONMENT (TASK 2)

Orbiterflight environments were definedfor the boost, orbit, andentry conditions. Airloads and thermal histories were estab- lished,including interference heating during ascent. Heating loaddistributions accounted for the variances in time at which transitionfrom laminar to turbulent flow occurs. Flow transi- tionnear the fuselage nose and wing leading edge occurs several hundredseconds following peak heating. Fully turbulent flow was assumed tooccur at a location twice thelength of thetransition onsetlength. Transition from laminar flow was consideredto oc- cur when: Ree = 10 El, (?)OS2 where

Ree = local Reynoldsnumber based on momentum thickness

% = local Mach number

= localunit Reynolds number

TrajectoryShaping

Designentry trajectories were establishedfor the ablator TPS and the RSI TPS. The ablatortrajectory was shapedto take advantageof ablator's high heating rate capability, as indicated infigure 1. The maximum heat r?te trajectorywithin the acceleration limit of2.5 g yields a q = 0.802 MW/m2 (70.7 Btu/ft2- sec)at the reference point. With a thermaldesign factor of safetyof 1.15, the design heat rate equals0.938 MW/m2 (81.3 Btujft 2 -set). Since s~,$-561* ablat9r is limited to a maximum heating rate of 0,692 MW/m (60 Btu/ft -sec), the entire lower surface of the Orbiter would require the higher densities of ESA- 3560HF* and ESA 5500*. Reducingthe entry angleand heating rate to 0.300 MW/m2 (26 Btu/ft2-sec)at the reference point permits use qf low density SLA-561 over98% of the Orbiter surface area. The q = 0.300 MW/m2 (26 Btu/ft2-sec)at the reference point was thereforedesignated as theablator design trajectory. The RSI trajectory was shapedto take advantage of the ceramic TPS low thermalconductivity for total heat insulation. By reducing entryangle to the skipout limit, heat rate is minimizedand total heat maximized resulting in an efficient RSI TPS design. Local static pressure histories were preparedfor the Orbiter for ascentand entry. The latter history is presentedin figure 2. Thesepressures are usedfor the ablator subpanel design.

Heating Rate Distributions

Heating rate distributionsnormalized to the reference point a at thelower fuselage centerline 15.24 m (50 ft) aft ofthe fuse- lagenose were preparedfor the ablator and RSI trajectories (figs. 3 and 4, respectively). * Elastomeric ablators developed by the Martin Marietta Corporation.

5 (70.7 Btu/ft2-sec)

.4Btu/ft*-sec)

(12.8 Btu/ft2-sec)

TIME FROM 12 I .92 km, 400 000 ft (seconds)

Figure 1.- TrajectoryShaping

800 I600 2400 TIME FROMENTRY (seconds)

Figure 2,- EntryLocal Static Pressure.History at Body Lower CenterlineReference Location

6 Figure 3.- EntryHeating Rate - AblatorTrajectory (WithoutFactors of Safety)

NOTE: I. e = 30°

Reference L0cotion”l 15.24m (50R) Aft of Nose d,,; 145 kW/m2 (12.8 Btu/ft“-sec)

Figure 4.- EntryHeating Rate - RSI Trajectory (WithoutFactors of Safety)

7 IV. HEAT SHIELD DESIGNS(TASK 3)

Thecontractual program investigated numerous locations on theOrbiter to establish heat shield configurations, weights,.and costs.The baseline method of ablator attachment was directbond. Alternativeconfigurations investigated were:

1)Ablator bonded to a subpanelplate and mechanically attacheddirectly to the Orbiter structure; 2)Ablator bonded to a subpanel honeycomb paneland mechani- callyattached directly to the Orbiter structure; 3)Ablator bonded to a subpanel honeycomb paneland mechani- callyattached through standoff fittings to the Orbiter structure.

The variousablator attachment configurations are shown in figure 5.

Eachconfiguration was examined for 12.7-cm (5in.), 25.4-cm (10in.) , 38.9-cm (15 in.) , and 50.8-cm (20 in.)attachment spac- ingfor weight and cost optimization. In addition, four materials ofconstruction for the subpanel were studied: aluminum2024-T81, Lockalloy,magnesium HM-21AY and graphitepolyimide.

DirectBond Plate Subpanel Baseline Aluminum 2024-T81 Lockalloy Magnesium HM-21A

Honeycomb StandoffHoneycomb Subpanel Honeycomb Subpanel Aluminum 2024-T81 Aluminum 2024-T8 I GraphitePolyimide Magnesium HM-21A GraphitePolyimide

DirectBond Through Strain Waior

Figure 5.- AttachmentConfigurations Investigated

8 ThermalAnalysis

Thermalanalysis for ablator sizing was carriedout with the Martin Marietta ThermochemicalAblation Program (TCAP 111). Data inputto this program include trajectory data, i.e., velocity, altitude,heating rate, andrecovery enthalpy; thermophysical propertiesfor the ablator material and backup structure materials;ablation kinetics; and geometry of the model being analyzed. Analysisresults include time-temperature distributions through- outthe model and a time-densityprofile through the ablative material. ,.

Analysis for sizing the RSI materialand for developing backup structuremodeling techniques was carriedout with the Martin Marietta ThreeDimensional Heat Transferprogram. Data inputand analysisresults are similar to TCAP I11 exceptthat no consid- erations are made foran ablation process. Both programs allow forvariations of conductivity with pressure as well as tempera- . ture.

Ablatorthickness design charts resulting from this effort are typifiedin figure 6. Therequired ablator thickness in this figure is basedon an entry start temperatureof 311°K (100°F) and a limit on maximum structuraltemperature after entry of 450°K (350°F). An additionalground rule assumes that the heating factor, F is theratio of maximum localheating rate-to-reference 4’ heatingrate at every point on the vehicle and is equalto a similar ratiofor total heats, i.e., - (41oJ;Ir.f) - Qloc/Qref* max The figure is usefulin that, knowingthe total heat capacity of thelocal structure, subpanel, bondlines, etc., to be protected, ($ rpCp , theablator thickness can be determined for any point onthe Orbiter 1 surface.

An example is given,based on a pointon the vehicle bottom centerlinethat is 15.24 m (50ft) aft of the nose. Here theheat- ingratio (fig. 3) is 1.00, amplifiedby the appropriate factor from from table 1 toyield a factorof 1.15. The nonablator components at thislocation and their physical and thermal descriptions are n presentedin table 2. Correspondingly >: -rpCP = 10.3 kJ/m2 OK i (0.00351Btu/in.2 OF). Enteringfigure 6 withthis argument and F. = 1.15yields a requiredthickness of SLA-561 of4.32 cm 9 (1.70 in.). 9 L I I I I I I I 1 I I 4 6 8 IO 12 14 (k J/~~oK) TOTAL HEAT CAPACITY, n qpicpi

Figure 6.- SLA-561 AblatorDesign Chart

TABLE 2.- THERMAL CHARACTERISTICS OF NONABLATOK COMPONENTS

P C I P Item Material I Thickness I Density HeatCapacity

cm in. kg/m3 lb/in. kJ/kg OK Btu/lb OF Orbiter 0.262 0.103 2768 0.100 0.941 0.225 Structure *luminumI I I I Ab lator RT V 0.076 0.030 1495 0.054 1.255 0.300 Bond Adhesive I 1 1 I Subpanel 0.079 0.031 2768 0.100 0.941 0.225

10 StressAnalysis

The plate andhoneycomb subpanels are designedby local aero- dynamic airloads(see table 3). Orbiterstructural strains are isolatedfrom the subpanels by using a mechanicalfastener in an oversizedhole. The strains anddeflections of thesubpanels must belimited to prevent induced strains in the ablator from exceeding 1%and to limit surfacewaviness to less than 0.0125 times theattachment spacing. Subpanel overall dimensions are limited by handlingcapabilities. A 107-cm (42 in.)by 107-cm (42 in.)panel size was selectedfor this program. Since total heatshield weight and cost optimizes with small attachmentspac- ing [less than 25.4 cm (10 in.) 1, theanalytical methods for in- ternalloads predictions were selectedfrom references 2 and 3.

TABLE 3.- DESIGN LOADING CONDITIONS

Designcriteria Configuration Designcondition Designload reference

~~~ ~ ~~ Platesubpanel Plate bending stiff- 6 = 0.0125 !L Servicelife mechanically ness strength analysis Airload(limit) = attacheddirectly Ablatorstrain limit 3.45 kN/m2 Service life tothe Orbiter of 1% (0.5 psi) ablator strain structure Airload(ult) = Environments 4.85 kN/m2 Dressures (0.7 psi) Honeycomb sub- Intracell buckling Environments panelmechani- offace sheet pressures callyattached directlyto the Orbiter structure Honeycomb sub- Intracell buckling Airload(limit) = Environments panelmechani- offace sheet 20.7 to 27.6 kN/m2 pressures cally attached (3 to 4 psi) Ablator strain limit Service life throughstandoff of 1% Airload = ablator strain fittingsto the (ult) 29.0 to 38.6 kN/m2 Orbiterstructure Panelflutter Environments (4.2 to 5.6 psi) acoustics I = CL3/E

The effectsof the studies are shown infigure 7; tab1.e 4 summarizesthe optimum fastener spacing and configurations for a large part of the Orbiter.

11 A- Direct Bond AbLator (Ref) B- Plate: x- Lockalloy y - Magnesium t- Aluminum C- Aluminum Honeycomb 0- Aluminum H/C + Standoff TABLE 4.- SUBPANEL GAGES REQUIRED FOR DIRECT E- Graphite a Glass H/C + Sta ATTACHMENT AfPLICATION F- RSI LI 1500 (Ref) Vehicle ,tation G-Mg H/Cand Attachment Subpanel h fastener Nose 584 to 2032 Graphite Poly HX: thickness Material (a). 230 to 800). identification spacing ' cap H-Mg H/C + Standoff cm (in.) cm (in.) cm (in.) 0.965 Aluminum 0.079 (0.380) (0.031)

12.7 "" 0.091 Plate thickness Magnesium (5.0) ("") (0.036)

"" 0.058 Lockalloy ("" ) (0.023) 0.025 0.013 Aluminum Honeycomb, each 25.4 (0.010) (0.005) face thickness (10.0) "" 0.020 Magnesium ("") (0.008) 0.041 0.020 Graphite (0.016) (0.008)

aAttachment fastener spacingdoes not apply for nose cap

1 J r I IO 110 I I (inched 0 25 50 FASTENER SPACING (cm)

Figure 7.- TPS Attachment Configuration Weight vs Fastener Spacing I

Total Heat ShieldConfiguration Weights

Thebaseline TPS weights were calculated in detail. A ratio process was usedfor the major components of the alternate designs.These yielded the total TPS weights shown infigure 7.

V. DESIGN VERIFICATION (TASK 4)

A test program was conducted to evaluate the mechanically at- tachedheat shield concept. The aluminum subpanel system at 25.4-cm (10-in.)fastener spacing was selected as beingrepre- sentativeof most of the direct mechanically attached configurations.Testing was performed inthree parts as indicatedin table 5.

TABLE 5.- VERIFICATIONTEST PROGRAM

- "~~~~ . ~~~ I Heat shield Test environmentltest Test objective Test specimens Results component facility -~ ~ . - .. ~~ Open gap/ Primary: Feasibility Reentry Heating/ 1 specimen, approximate Full heating profile not attachments of self-closing gap Structures Laboratory, size 1.12 x 0.56 m achieved (panel test) concept Radiant Heat Facility (44 x 22 in.) Indications of incomplete Secondary: gap sealing a. Motion of subpanel 0 Indications of inadequate plate under stud subpanel motion bolt 0 No abnormally high heat- b. Temperature distri- ing in counterbore bution around stud bolt counterbore ~~- ~ ~ ~~ Gap sealer Resiliency after high Dead load compression 17 specimens, Blanket material shoved (component temperature deforma- and uniform heating/ approximate size adequate resiliency test) t ions Advanced Structures 5.08 x 15.24 x 0.254 to 0 Rope material nonresilient and Materials Ceramics 1.29 cm (2 x 6 x 0.1 to Laboratory 0.5 in.) . ~__ ~ .- - -- -. -. __ Sealed gap/ Primary: Feasibility Ascent and Reentry 1 specimen, approximate Ascent heating - no ap- attachments of sealed-gap concept Heating/Structures size 1.12 x 0.56 m parent effect on panel (panel test) Secondary: Laboratory, Radiant (44 x 22 in.) Descent heating a. Motion of subpanel Heat Facility - some backface tempera- plate under stud ture > 450°K (350°F) bolt - high temperature in gap b. Temperature distri- - indications of inadequate bution around stud subpanel motion bolt counterbore - high temperature in counterbore

13 Open Gap Test

Ablatorpanels can be installed with a limitedwidth gap at thejoint between adjacent panels 0.318 cm for 107x107 cm panels (1/8in. for 42x42in.). Due tothe thermal coefficient of expan- sion of theablator, the gap closes with increasing temperature, limitingaerodynamic heating at thebottom of the gap. The entry heatingpulse was simulatedusing radiant heat lamps. Although heating was notuniform, some gapclosure with acceptable temperature limits at bottomof gap was noted.The attaching fasteners tendedto bind, inhibiting free thermal motion of the subpanel. Temperaturein the fastener cavity in the ablator was notexces- sive.

Gap Sealer Tests

Candidate materials were tested inovens under deformation at elevatedtemperatures. Springback or recovery of deformation followingtemperature reduction was measured.Five to ten per- centspringback was desiredfollowing heating. Only the Fiber- frax* blanket compressed normal to thefiber direction Was satisfactory.

Sealed Gap Test

Ablatorpanels can be installed with a sealer in the joint betweenadjacent panels. During ascent heating the gap closes partially,then opens while in orbit, closing again at entry. The selected sealer was Fiberfraxblanket, which is resilient enough totake these thermal cycles. The test panel was exposed toascent and entry heating using radiant heat lamps. The sealer didnot perform satisfactorily since temperatures were higher in thebottom of the gap than under the ablator. Fiberfrax blanket is stratified andthe thermal conductivity is thereforetrans- verselyisotropic with the low values across the gap and the high valuesparallel to the gap.

* Product of the Carborundum Company, Niagara Falls, New York.

14 VI. WEIGHT AND COST ANALYSIS (TASK 5)

Thisprogram has developed design concepts and weight data forseveral competing methods for ablator attachment to the Orbiterstructure. It remainsonly to select thebest attachment systemon a weightand cost basis. Competing heat shield weights were convertedto costs and added to operational costs to estab- lishthe total program costs. The ablatorconfiguration choice onthe basis of minimum programcosts thus includes the effect of heat shield weight.

PayloadWeight Penalties

To aid in the optimum ablatorheat shield selection, an ablatorheat shield was assumed forthe first five years or 151 flights ofShuttle operations--1979 through 1983. The payloads planned for that period were identified from the MSFC traffic model. Eachof the competing heat shield design weights was compared to thepayloads of the traffic model and payload weight penalties, if any were determined. To ascribe a dollarvalue to these weight penalties, a costper pound toorbit was derived. All program costs were apportionedagainst all thepayload weight. A DDT&E cost of $5150M* was apportioned to 445 flights for a cost per flightof $11.57" or $1747.5M for the first 151 flights. Added to this is an operational cost of $10.5M** per flight or $1585.5M forthe first 151 flights. The totalcost for the first five years of 151flights is $3333M. The first 151 flightscarry a totalof 2.07 Gg (4.56M lb)of payload. Dividing $3333M by2.07 Gg (4.56M lb)yields a unitcost to orbit of $1612/kg ($731/lb). Thisdoes not include the costs of the payloads themselves.

Thepayload weight penalties for each of the competing ablatorsystems was multipledby this cost per unit weight to orbit of$1612/kg ($731/lb) and added to the operational costs of each ablatorsystem to establish a totalprogram cost. The RSI heat shieldweight was used as a baseline to establish payload weight penalties.In many cases, thefull payload weight capability of theOrbiter was notused due to volume constraints. This reduced payload was used toestablish payload weight penalties (table 6), for three typical missions. * Estimated costs used during Shuttle Phase C-D pre-proposal briefings. ** Estimated cost given by Bastian Hello, RI, at AIAA/ASME/SAE 14th Structures, Structural Dynamics,and MaterialsConference, Williamsburg, VA, March 20-22, 1973. 15 TABLE 6.- THREE MISSIONS OF THE 1981 PERIOD

Flightno. (a) (payload no.) 15 (NCN-10) 18 (NE?.-44) 20 (NEO-16) Payloadloading factor 0.31 0.38 0.83 29484 kg 65 000 lbMission 20412 capability kg 45 000 lb 20412 kg 45 000 lb 9 140kg 20 150 lb Payload bay 7load 757 kg 17 100 lb 16942 kg 37350 lb 20344 kg 44850 lb Unused 12capacity 655 kg 27 900lb 3 470kg 7 650 lb RSI designweight 13 717 kg 30240 lb kg 30240 lb 13 717 kg 30240 lb13 717 Missionstandard weight 34 061 kg 75090 lb 26372 kg 58140 lb 17 187 kg 37890 lb 38.1-cm (15in.) fastener spacing,direct attach aluminumpate TPS system weight 22128 kg 48 782 lb 22 128kg 48 782 lb 22 128 kg 48782 lb Weightpenalty -0- -0- -0- -0- 4 941kg 10 892lb @ $1612/kg($731/lb) -0- -0- -0- -0- $7962 052 ($7 962 052) I 1 I I I -Note: All weightpenalties for all flights are added for the first 5 years(151 flights). For the case of 38.1 cm (15in.) spacing direct attach aluminum plate:

-YearPenalty $J 1979 -0- 1980 -0- 1981 $ 15924 1982 -0- 1983 348 484 3481983 -~ Total$364 408K

AverageWeight Penalty = $36:5,-08K = $2413K/Flight I (This TPS System) 1 aReference 1 nomenclature.

All thepayload penalties for all flights were added,resulting in $364M cost penalties for the aluminumsubpanel with 38-cm (15in.) spacing, the example in table 6.

OperationalCosts, TPS

Operationalcosts are thefabrication, refurbishment, and qualityassurance tasks incurred during the 151 flightprogram. Thesecosts include:

1) Ablator slab raw materials andfabrication including scrappage(this accounts for 3/4 of the operational cost);

16 Subpanel raw materials andassembly costs, including scrappage ;

Assemblyof ablatorslab to the subpanel;

Installationof ablator panel assembly--tools and labor;

Removalof used ablator panels--tools and labor;

Repair materials andlabor due to damagefrom handling duringpacking, shipping, storage and installation;

Bond line inspection;

Mechanicalfastener inspection;

Subpanelfabrication inspection;

Refurbishmentcleanliness inspection;

Inspectionfor damage followingablator installation;

Inspectionof repaired areas.

The totaloperations costs, table 7, show the direct bond heat shieldand the aluminum and magnesium subpanel mechanically at- tachedheat shields to be very competitive with the lowest cost 25.4-cm (10 in.)fastener spacing subpanels. This is dueto the reduction in ablative materials requiredand lower installation costsdue to smaller numbersof fasteners. An illustrationof theproportional effects ofelements of TPS operationalcosts is presentedin figure 8.

TotalProgram Costs

The inclusionof the heat shield weight as a payloadweight penalty cost has a direct effect ondetermining the optimum heat shieldselection (table 8). The direct bondsystem has the lowest total programcost, followed closely by the magnesium and aluminumsubpanel plate systems at 12.7-cm (5 in.)fastener spacing.The low operational cost magnesium and aluminum subpanels at 25.4-cm (10 in.)spacing are prohibitivelyheavy as is shown by the $34.9M and $76.9M payloadweight penalties. For the direct bond system,the effect of dust and debris and turnaround time was notevaluated.

L TABLE 7 .- OPERATIONAL COSTS

($W forFive years (151 Flighcs)

iblator Sub- PanelInstallation Fastener slabpanel assembly 6 removal Repair assurance Total Subpanel spacing - Direct N. A. 125.7 6.8 ______4.4 27.9 164.8 bond - Aluminum 12 0.7 1.2 8.7 21.9 8.7 1.2 120.7 6.6 9.2 168.3 plate - Nagnesium plate 122.1 1.8 8.7 22.4 6.8 10.0 __173.8 Lockalloy plate 120.7 31.1 10.4 25.6 6.8 10.6 205.2 12.7 - Aluminum cm honeycomb (5 in.)8.7 124.9 12.18.7 7.0 25.7 187.1 __ Magnesium honeycomb 121.8112.2 I 8.7 1 28.1 I 6.9 I 13.0 190.7 __ Graphite 12 2.3 15.0 8.7 25.7 6.8 12.0 6.8 composite 25.7 8.7 15.0 122.3 190.5 honeycomb - Aluminum 154.7 plate - Magnesium 156.6 plate - Lockalloy 7.7 207.5 plate - Aluminum 25.4 cm (10in.) 124.9 1 7.8 1 8.5 1 18.0 1 6.7 1 9.0 174.9 honeycomb - Plagnesium 121.8112.0 19.4 9.6 ~ 177.9 honeycomb 1:; 1 1 1:; 1 - Graphite 12 2.3 17.2 18.0 9.1 18.0 composite17.2 122.3 181.7 honeycomb __

\ lnstallation \

Figure 8.- RelativeOperational Costs of a TypicalThermal Protection System [12.7-cm (5 in.) Aluminum Plate Subpanel]

18 TABLE 8.- TOTAL PROGRAM COSTS

..- .. - -- . . ~ ~ ($M) for 5 years (151 flights) I I Configuration Heat Pay1oad Total Operational w eight, weight costs programcosts weightweight, p en alty costs penalty kg (lb)

l2 337 0 164.8 N. A. (27199) I I

- ." .. ~ ~ ______-" Aluminum 168 .3 178.9 168.3 plate (32577) Magnesium 14335 173 .8 177.4 173.8 plate (31602) 13650 205.2 205.2 205.2 (30092) Aluminum 14 587 I 194.4 187.1 I honeycomb (32158) Magnesium 190 .7 191.2 190.7 honeycomb (30882)

14059 composite 191.1 190.5 (30994)

Aluminum 154 .7 231.6 154.7 plate (37803) Magnesium 15 757 156 .6 191.5 156.6 plate (34738) Lockalloy (3114 549) 311 207 -53.3 210.8 plate "1 25.4 cm Aluminum (10 in.) honeycomb Magnesium honeycomb Graphite composite honeycomb

Consideringthese items qualitatively,the slightly higher cost of direct attached magnesiumand aluminum platesystems at 12.7-cm (5in.) fastener spacing are attractive alternatives.Mechanically attached heat shields must be selected prior to Orbiter CDR to per- mit inclusionof anchor nuts in the basic design. Adding this fas- tenerhardware after the original engineering release becomesin- creasingly difficult and at'the point where Orbiter hardware has beenfabricated without heat shield anchor nuts, adding them be- comes prohibitive. 19 VII. DISCUSSION OF PROGRAMRESULTS

Based on heat shield weight alone, the first six lowestweight ablatorsystems are ranked:

Direct bond ...... 12 337 kg(27 199 lb)

Lockalloysubpanel at 12.7 cm (5 in.) ...... 13650 kg (30 092 lb)

Magnesiumhoneycomb at 12.7 cm (5 in,) ...... 14 008kg (30 882 lb)

Graphitecomposite honeycomb at 12.7 cm (5 in.) ...... 14 059kg (30 994 lb)

Lockalloysubpanel at 25.4 cm (10in.) ...... 14 311 kg (31 549 lb) Magnesium plate at 12.7 cm (5in.). . 14 335kg (31 602 lb)

Basedon total program costs, the first six ablator systems are ranked :

Direct bond ...... $164.8M

Magnesium plate at 12.7 cm (5in.) ...... $177.4M

Aluminum plate at 12.7 cm (5in.) ...... $178.9M

Magnesiumhoneycomb at 25.4 cm (10 in.) ..... $182.9M

Graphitecomposite honeycomb at 25.4 cm (10in.). $186.4M

Aluminumhoneycomb at 25.4 cm (10 in.) ..... $186.9M

The only repeaters inboth lists are thedirect bondand the magnesium platesubpanel at 12.7-cm (5in.) attachment spacing.

Whilethe direct bond ablatorhas the lowest heat shield weightand program cost, the concerns with dust control and mini- mization of turnaround time duringthe refurbishment period makes thesecond alternative attractive. Magnesium HM-21A subpanel at 12.7-cm (5 in.)fastener spacing is less than 5% heavierthan the baselined RSI andonly 8% higherin total program cost than the direct bondheat shield.

20 VIII. CONCLUSIONS AND RECOMMENDATIONS

Conclusions

0 The design criteria forablative thermal protection systems on a SpaceShuttle Orbiter are comprehensiveand complete in scope.

0 A rangeof entry trajectories is availablethat fully uses anablative TPS--all within 2% g limitations. At oneend of this spectrum is a short time, highpeak heating rate entry that would demand considerableusage of dense ablator materials. Ex- tendingthe time durationof entry reduces the 'heating conditions tolevels which permits lightweight ablators over most ofthe vehicle.

Direct bondingof an all-ablator TPS Clow density SLA-561) tothe Orbiter structure yielded the lowest TPS weight of all theheat shield systems evaluated [weight factor (WF) = TPS(i)/ TPS(RS1) = 0.901. An RSI TPS was nextlowest (WF = l.oO),fol- lowedby a series ofdesigns involving mechanically attached sub- panelssupporting SLA-561 (WF = 1.OO to 2 .OO) . Fastenerspacing was influentialin the total weight of the latter designs.

0 A feasiblecost model, involving a weightpenalty of $1610/kg ($731/lb), was derivedbased on an apportionment of program costs tothe first 151 flights (assumedduration of utilization of all- ablator TPS) andthe total payload weight carried in these flights. Thispenalty was employed inevery instance where the total heat shieldweight exceeded a givenparameter.

0 The direct bond ablatorsystem had the lowest program cost ofall the ablator configurations examined (the RSI system was notcosted). No weightpenalty (dollars) was requiredfor this system.

0 The nextbest cost ablative system, magnesium HM-21A subplates, directlyattached, would incur $12 million more thanthe direct bondedarrangement. This was closelyfollowed by the similar systemusing 2024-T81aluminum ($14 million more) .

0 Inthe three candidate ablator designs highlighted above, approximately 3/4 ofthe TPS operationalcost involves the fabri- cationof the-ablator slab. The otherquarter encompasses assembly,installation, removal, tooling, repair, and inspection.

21 A typical TPS operationalcost is approximately 10% of the total program'sestimated operational cost.

0 The use of a nonablator,insulative material inthe gaps be- tweenpanels tended to make the structure along these lines hotter thanthe remainder, as demonstrated in a large scale test.

0 A test toinvestigate the feasibility of experiencing gap closure before high heating was encountered was inconclusivebe- causeof poor heat distribution in the test assembly.

0 A conceptof a fastenerdesign that would provide some degree of movement betweenablator subpanels and the structure was es- tablished.

0 An earlydecision in the design of anablative TPS mustbe made concerningthe incorporation of anchor nuts in the structure ofan Orbiter to accommodate fasteners.

Recommendations

Costreductions with respect to ablative systems should con- centrate on thebasic slab fabrication--materials, processes, inspection, etc.

Additionaleffort should be expended to find an acceptable gap sealer; i.e., caulking, etc.

0 Additionalinvestigations should be made on theconcept of self-sealing of gapsbefore the high heat time period.

The fastenerpresented should be reevaluated for greater tol- erancesand, possibly, Teflon coating.

0 The feasibilityof reuse of siliconeablators installed in lowheat regions should be further examined.

22 NASA-Langley, 1975 CR-2S79 REFERENCES

1. E. S. Love: "Advanced Technology and the Space Shuttle" Tenth von Karman Lecture, Antronautics and Aeronautics, February 1973 2. S. Timoshenko and S. Woinowsky-Krieger: Theory of Plates and Shells, McGraw-Hill Book Co., 1959.

3. s* Woinowski-Krieger: Bending of Plates Supported by Rows of Equidistant Columns, Mc-Graw-Hill Book Co., 1959.