Ablator Recession Speeds

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(NASA-CR-17C801) S6 B T H t F t J A L PROTECTIOHS f S T B R S HATt'BIALS TEST RESULTS IN A NAFC-HEATED Y I T R O G Y N P I V I i i O N H E B T ( L o c k h e e dmissi les a n d S p a c e Co.) 6 2 p HCTAC4/PlF A C 1 UcclasCSCL 1lD S3/24 16C46


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FOREWORD

This report documents the r e su l t s of a materials test of Solid RocketBooster thermal protection s ys tem s conducted at Acurex Corporat ion in the ir1 Megawatt Arc Pksnia Generator Facil i ty .verify the thermal protect ion s yste ms mate r ials performance in a high heat-ing and high enthalpy environm ent similar to Space Shuttle Solid RocketBoost er staging environment. Acu rex personnel conducted the tests, andLockheed-Huntsville provided a tes t monitor .

The purpose of the test was to

Lockheed-Huntsville support for the tests is provided under ContractNASS-32982, 5 o l i d Rocket Booster Thermal Pro tect ion System Mater ia lDevelopment." The NASA -MSFC Contracting Officer' s Representat ive forthis contract is Mr. Williarn Baker, EP44.the Acurex test support contra ct . The Acurex test engineers were M r . L.Arnold and Mr. E. Fretter; the Lockheed-Huntsville test engineer wa s Mr.C. . Wojciechowski.

Mr. Baker was also the COR on

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CONTENTSSection

I2

FOREWORDNOMENCLATUREINTRODUCTION A N D SUMMARYTECHNICAL DISCU S ON2.1 Te st Desc ription2.2 Data Analysis

3 TEST RESULTS4 CONCLUSIONS

REFERENCESAppendix

Acurex Report

LIST OF TABLESTable

123456

Lis t of TPS Tes t SpecimensPan el Local-to-Stagnation Poin t Heating Rate RatioPane l Local Aerothermodynamic Relat ionships for theThre e Test LocationsTPS Tes t Sample s Weights and Thickness M easurem entsAero therm Probe TPS Te st Resu l tsAero therm Panel TPS T est Resu l ts

Pag eiiV

1

192728

A - 1

68

10202122

LIST O F FIGURESFig u re

1 Heating Rate-Shear Stress Relationship for Tes t and Fl igh t 1123

Constant Plu me Impingement Shea r Stress Contours att = 5 sec 12Constant Plum e impingement Shear Stress Contours a tt = 6 sec 13

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Figure Page45678

Constant P!urne Impingement Heating Rate Contoursat t = 5 sec 14Constant Plume Impingement Heating Rate Contoursat t = 6 sec 15Temperature-Enthalpy Relationship for Nitrogen and SSMEPlume Wash 16Wall Temperature Effect8 on Convective Heating Rates forNitrogen and SSME Plume Wash 17Recession Rate v s Heating Rate Design Curves and TPSTest Results 23

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EnglishHhMP;1RT

Greek7-Subs cr i pts

cweL0

OLr

N O M E N C L A T U R EDesc r piion

total enthalpy, Btu/lbmstatic enthalpy, BtuilbmMach numberpressure, lb/inheating rate, Btu/ft -sec

22

rec ess ion rate, rnils/sectemperature, R

2shear stress, lb/in

cold wa ll defined at 460 Rboundary layer edge conditionlocal conditionstagnation point conditionslocal stagnation condition

reco very value

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1. INTRODUCTION AND S U M M A R YThe external sur face of the Solid Rocket Bo oster (SRB) wil l experience

imposed thermal and sh ear environments due to aerodynamic heating andradiation heating during launch, staging and reen try.sys tem (TPS) is a n insulat ion sys tem applied to the external surf aces of theSRB for maintaining the s tr uc tu ra l and component tempe ratu res within theirdesign l imits.TFS mate rials during the staging maneuver .th e Space Shuttle Main Engine (SSME) exhaus t plume s impose sev e re , sh o r tduration, thermal environments on the SRB.w e r e tested in the 1 MW A r c Plasma Genera tor ( A m ) facility of Acurex/Aerotherm.and aerodynamic she ar environments over most of the SRBlocal hot spo ts on the SRB with pre dicted SSME plume wash heating ratesspikes of 360 Btu/ft -se c wer e not simul ated.ing rate obtained i n th e APG faci l i ty was 248 Btu/ft -sec, however, the testduration was such that the to tal heat was m or e than simulated.some loca l high she ar stress levels of 0.04 p s i a we re not simulated. Mostof the SSME plume impingement area on the SRB experiences shear s t ressleve ls of 0.02 psia and lower.were between 0.021 and 0.008 psia.(in the SRB mpin gemen t regio n) of 5260 R tempera ture , 6000 Btu/lbm en-thalpy and 3 psia pres sure w ere simulated using arc heated nitrogen withstagna tion conditions of 9700 R temperature, 4800 Btu/lbm enthalpy and 2.7psia stagnation pressure.

The thermal protectiov

This repor t is concerne d with the performa nce of the var iousDuring staging, the wash f rom

Five d i f fe ren t SRB TPS materials

Th is facility allowed simulation of the SSME aerodynamic heatingarface. Some

2 The maximum simulated heat-2Similar ly ,

Th e sh ea r stress levels on the tes t spec imensThe SSME plume sta gnation conditions

The TPS mate r ial s amples held up as expected o r bet t er than expectedin terms of mater ia l recess ion rates under the simulated SSME plume washenvironments. In t e rms of virgin material r ecess io n rates, the five TPSmater ia ls ranking from highest to lowest are: M A - 2 , MTA-2, P50 and

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phenolic glass (both ranked same), and finally B-Stage cork.of the TPS materials was a nominal 0.30 in. The test data indicates that thisthickness i s more than sufficient to protect against the SSME plume washthermal environments as simulated.

The thickness

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2. TECHNICAL DISCUSSION

Discussed in the first part of this section a re the features of th e TPStest facility, calibration methods, T FS specimen descripti ons, data mea sur e-ments and data reduction.cr i te r ia and the test data analysis.was to obtain SRB TPS material ablat ion character ist ics (virgin materialrecess ion ra te s and sur face tempera tures) in a sho rt duration, high heatingand enthalpy environment representative of the SSME plume wash conditions.On the SRB, the highe st heating rates occur on the SRB s t ruc tura l p ro tuber -an ces which use phenolic g las s TPS.la te the heating rates on the acreage areas where other TPS materials a r eused. To simulat e this range of heating rates two test configurations wereemployed.simulation and a panel configuration was used to simulate the a c r ea g e a r e aheating.phenolic glas s manufactured by Edler Industrie s, Inc., MTA-2 and M A - 2both of which were developed by NASA-MSFC. A l l the ma ter ia ls were tes tedon both model configurations in order to obtain a good variation of virgin ma-terial recess ion r;te as a function of cold wall heating rate and shear s t resslevel,

The second part d i scu sses the flight simulationThe main objective of the test program

In addition, it was a lso des i red to s imu-

The probe configuration was used for the higher heating rate

The mater ials that were tes ted were P50 sheet cork, B-Stage cork,

2.1 TES T DESCRIPTIONThe tes ts we re conducted a t Acure x Corporation in thei r 1 Megawatt

A rc P lasm- Genera tor (APG) facility. A complete description of this facility,a s w ell a s rile Acurex f inal data report is included in this re po rt in the Appendix.The tes t gas used was nitrogen.oxygen-free high enthalpy environment similar to the SSME plume wash.SSME plume wash consis ts mainly of 75% wa ter vap or and 25v0 hydroge n g as.The reaction of the water vapor with the carbon char layer was not simulated.

Nitrogen was selected because it provided anThe

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Hclwever, e stim ates of th e reaction r at es for this reaction under the lowpres sure environment indicates that this reaction would not be dominant.Because of differences in the specific heat8 of the SSME plume wash and thea r c heated nitrogen gas , the temperature-e nthalpy relations hip could not besimulated, Le., stagnation tempe ratu res of 9950 R at an enthalpy of 4 7 8 1 Btu/lbm for the a r c heated nitrogen as compared with SSME plume wash skgna-t ion temperatures of 5260 R a t a n enthalpy of 6000 Btu/lbm. Sincz anthalpypotential is the main driving for ce in convective heat t rans fer , i t was d es ir t dto simulate as close as possible the enthalpy potential as this would bettersimulate the hot wall convective heating rate s. The nozzle siz e was selectedto yield a Mach number of 3.53 ap pro ac h flow which would simu late the lo calSSME plume and Mach number.

With the test gas selected, the next phase wa s to run a s er ie s of cal i-bration tests to dete rmin e the the rm al environment about the models. Sincea 2 in. exit diameter nozzle was used the TPS models were small in o rderto have a s u niform a flow field as possible over the model surface.probe model was a 1 in. dia me ter fla t disc and the panel mod el was 1.25 in.by 3 in. a s shown in F i g s . 2 and 3 on page 4 of the Appendix.brat icn run s, the standard Acurex f lat face slug calor ime ter cal ibrat ion probeand separate pitot probe were used for the probe models.models, a flat panel calibration model w a 8 built by the Lockheed-Huntsvillemodel shop.Appendix. The calibration m odel featured 3 thin skin (0.030 in. nominal) heattransfer sensing are as , one Gardon gage calor ime ter , and three local pres -sur e m easurem ent locations.measured using a micrometer, prior to placing the 30 gage wire chrome1alum el thermocouple junctions.on page 1 1 of the Appendix.

The

For the cal i-

For the fla t panelThe panel calibrati on model is shown in Fig. 5, page 8 , of the

The thin skin arca th icknesses wer e accurately

The calibration test procedures are given

Th e TPS est spec imens were a l l a nominal 0.30 in, thick mounted onindividual 0,125 in . thick aluminum backup plates.c Tuples wer e mounted on the backside of the alumin um s ub st ra te plate.

The backface thermo-

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The probe and panel test specimens a re shown in Fig s. 2 and 3 in theAppendix.mounted in their respective holders for testing.testing the TPS specime ns i s listed on page 14 of t he Appendix. A l is t of a l lthe TPS specimens which were tested a re given in Table 1.prepared by NASA - M S F C Materia ls Laboratory with help from Lockheed-Huntsville.test th icknesses and weights we re m ade by NASA-MSFC M ater ials Lab a s wella s placement of the thermocouples. Po st- tes t virgi n material th icknesseswer e made a t Lockheed-Huntsville. All the models we re f ir st tested at thelower expos ure time and then inspected by the Lockheed-Huntsville onsitetest monitor . If the models looked good with plenty of virgin material re -maining and the backface tem pera ture ri s e was low, the next similar TPSspecimen was tested a t the longer exposure time. A complete descr ipt ionof the te st instrumentation is given in Section 3, pages 6 thro ugh 11 of th eAppendix. All of the Vi sic ord er data reduction and analysis was done on siteby the Lockheed-Huntsville monitor, af ter instruction fro m Acurex personnel .Thi s included both the cali bration runs and the TPS specimen runs.way th ere we re no delays in setup time and communication and the next APGrun could be prepa red by the Acu rex te s? engineer while data fr om the pre-vious run wer e being reduced and analyzed,3 f the reduced Visic order and surface temperature data we re made for ver i-fication and co mpariso n with the Vidor DDAS data for inclusion in the Acurexfinal data report .TPS te st configuration, and Fig. 7 on page 16 (Appendix) show s thc; TPS panelmodel t est configuration.the quartz windows,

F ig u res 1 and 4 in the Appendix show how the te st specim ens we reThe proce dure used when

The models wer e

The models we re photographed pri or to the test at Acurex. Pre -

In this

Upon test completion, copies

Figure 6 on page 15 in the Appendix shows the pr obe modelDuring testing, the models we re viewed (hro ugh

2.2 DATA ANALYSISA detailed listing of the test instrumentation and data reduction methods

is given on pages 6 throug h 11 of the Appendix.cern ed with determining the A P G facility flow field, the model flow field,extrapolation to flight conditions, and TPS spec imen recess ion measurements .

The d iscuss ion he re is con-

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Run No.I23456789

10I 112131 41 51617181920212223

Table 1LIST O F TPS TEST SPECIMENS

Configuration Model No.c-1P C - 1P C -2P A - 3P A -4P A - 5P A - 6P B - 1P B - 2P D - 1P E - 1P E - 2P D - 2c - 2A - 1A - 2B -1E - 1D - 1E - 2D - 2c - 3B-2

TPS Mater ialP50 Sheet CorkP 5 0 Sheet C o r kP50 Sheet C o r kEdler S-Glass PhenolicEdler S-Glass PhenolicEdler S-Glass PhenolicEdler S-Glass PhenolicB-Stage Sheet C o r kB-Stage Sheet CorkM SA - 2M T A -2M T A - 2M S A - 2P 5 0 Sheet C o r kEdl er S-c:lass PhenolicEdle r S -Glass ''henolicB-Stage Sheet CorkM T A - 2M S A - 2M T A - 2M S A - 2P50 She& C o r kB-Stage Sheet C o r k

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2.2.1 APG Facility Flow FieldDuring the calibration phase of the program , the ar c chambe r p res su re

and model pitot pre ssu re and heating ra te s we re mea sure d. Using these dataand the therm al properties of high temp erat ure nitrogen, the Mach number ofthe plasma jet centerline was de termined to be approximately ~ . 5 3 using aneffective gamma (r at io of specifi c heats) of 1.30. The nitrogen gas gammavaries from 1.17 in the chamber to 1.38 in the highly expanded regions of th eflow field.2.2.2 Model Aerothermodynamic Environ ments

During the model TPS tests only the a rc chamber p ress ure and s tagna-tion heating rat e were m easured . This presented no problem for the probeTPS tes ts, but for the panel TPS est s, the local heating r at es had to be de-rived fr o m the stagnation point heating rate. During the cali brat ion phase,both the probe and the panel calibration model were immersed sequentia'a t the sa me stabil ized ar c condition.stagnation poihit heating rate were established for the three instrumentedlocations on the panel as shown in Table 2.

F r o m his data rat ios of local-to.

The model local shear s tr es s calculation was calculated using the sam egene ral fo rm of equation that w a s used in the preflight predictions for theSSME plume wash, (Refs. and 3). The equation used was

0.008372 4M L J T(H r h w p i a ,- 1=

where 2local heating rate, Btu/ft -s eclocal Mach numberboundary layer edge temperature, Rrecovery enthalpy, Btu/lbmwall enthalpy a t 460 R , Btu/lbm

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_I _.LTest N o. Location 1.'."3 1 5 0 - 0 2 . 3 2 03 1 5 0 - 0 3 . 3 5 43 1 5 1 - 0 1 . 3 9 2Test Averages 1 . 3 5 5

Tablc 2HEA TING FZA TE RA TIO:::P A NEL LOCAL - TO - 3 T A G N A TION POINT

L o c a t i o n Z Location 3. 2 5 3 .115,238 . l o 4.262 , 1 0 6.Z51 . l o 8

-I-

c w kefined a s qS C C Fig . 5 of Appendix for o c a t i o n s ..a JI,. ,.

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rLMSC-HREC TM D697786 i'iFor he panel test s the local Mach number was obtained fro m the ra t io ;

of the local pressure to the pitot pre ssu re. Then using the local bdach number,the boundary laye r edge tempe ratur e and enthalpy we re determ ined using idealgas relationships and a gamma of 1.3.tions a re shown in Table 3. A boundary l a y e r recovery fac tor of 0.9 was used.Figure 1 presents the heating ra te-sh ear st re ss var iat ion for both the probeand panel configurat ions. The shear st re ss level OD the probe configurationwas c alculated z t the junction between the TPS specimen and the graphite collar .The flowfield pro pert ies a t this junction we re evaluated with the ass ist anc e ofthe data presented in R e f . 4.s t r e ss l ev el s a t two time points from Ref. 2.corresponding S S M E plume wash heating rate levels f r o m R e f . 2.shown in Fig s. 2 through 5 a r e t 'cle antt body v alues.presented in R e f . 3 but no corresponding shear stress levels a r e presented.Comparison of Fig. 1 with Fig s. 2 through 5 indicate that the heating rat e andshear st re ss levels were well simulated in this tes t for mo st of the SRB im-pingement area .

The panel resu lts f or the three loca-

F ig u res 2 and 3 show the S S M E plume wash shearShown in Fig s. 4 and 5 a re the

The valuesProtubera nce heating is

2.2.3 Data Extrapolation to FlightDue to differences in the heat capacities of the SSME plume wash and

the A P G nitrogen, the relationships between thc ratio of cold wal l to hot wallconvective heating a r e different for the tw o gases .temperature-enthalpy relationships for the two gases.ra ti o of cold wall to hot wall convective heating rate versu s wal l tempera turefor the SSME plume wash and one recovery enthalpy value for the APG. Alsoshown is the r ati o of flight cold wall to tes t cold wall heating ra te as a functionof wall temp erat ure for thes e two particular recove ry enthalpy values, Basic-a l l y for each TPS est , the flight cold wall simulated heating rate was calcu-lated using the following procedure:

F i g u r e 6 depicts theFigure 7 shaws the

1. The local cold wall heating r at e was calculated fr om the mea sure dstagnation point heating rate and the appropriate local factor f orthe panel fro m Table 3.the recov ery enthalpy was determ ined.Using the appropriat e factor f r om Table 3,

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PanelL o c a t i o n123

l ablc 3PA N EL LOCA L AEROTHERMODYNAMIC RELATIONSHIPSF O R THE T HR E E T E S T L OCA T ION S

pe/poL7V H 0 V T 0 M~ (ps i ),,/qo.355 .962 .66 1 1.85 .021 .473.25 1 .94 . 478 2.70 .018 .116.lo8 .930 .376 3.33 .008 .027,

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L , Using the mea sure d wall temperature and Fig.6 , the test hotwal l heatir-g rate was calcula ted using:)1 .(w test - hhwtestt e s t test- cwest (H rhwtest - qcw

Then by definition the test hot wall heating rate was as sumed to beequal to the flight ho t wall heating rate.3. Using the measured wal l tempera ture and Fig . 6, the flight hot wallenthalpy was de termined .5814 Btu/lbm.lated using

The flight recovery enthalpy used wasThe flight cold wall heating rate was then calcu-

!H - h 1rflt C W f l tL (H- - k... 1 -a- Xf l t =hwf l r . f It 'lWf It=cw

where the hcw is defined at 460 R.

2.2.4 Model Reces sion Me asur eme ntsEach model was weighed im mediately aft er test. Post-test photographs

were taken a t MSFC. The amount of virgin material remain ing was measu redaf ter the ch ar layer was carefully machined away unti l the virg in mater ia l waserposed .models and at the three measurem ent loca t ions on the flat panels.

Thickness measurements were taken at the cen ter of the probe

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3. TEST RESULTS

The prel iminary TPS mater ia ls test res ults a r e presented in Table 3of the Appendix.for each mater ials test .and thicknesses with and without the post-test char a re given i n Table 4. Thepost-test weights shown in Table 4 a re somet imes g rea te r than the p re tes tweights. The probable rea so n fo r this is that the pretest weights were madea t MSFC and the post-tes t weights made a t Acurex by different personnel anda different scale.cha r retained, it is evident that the only mate ria ls th at exhibited any charremoval were the MTA-2 material and to a lesser extent the MSA-2 material.The cork mate rial s exhibited swelling during the test. The only correla tion sthat were made in th is report w a s with the pretest thickness and post-testthickness measured with the char removed.

Table 4 in the Appendix presents the APC run conditionsThe TPS tes t sample pr etes t and post-test weights

In observing the post-test ma ter ial thicknesses with the

The probe TPS test r es ult s ar e presented ir. Table 5. The recessionrates presented in Table 5 a r e based on the post-test cha r removed thick-nesses and the exposure time.Table 6.test s and ar e given for the three instrumented locations on the panel.posite plot of the TPS reces sion rat e vers us cold wall heating ra te i s pre-sented in Fig. 8. Also presented in Fig. 8 ar e the cur ren t TPS mater ia lrecession rate design curves for the var ious TFS mate r ials .

The panel TPS test res ults a r e presented inThe recession ra tes we re calculated the same a s for the probe

A com-

A l l of the TPS mat erial s sample s tested held up as expected o r muchbetter than expected under the simulated SSME plume impingement environ-ment. The var ious mater ials tested and their resu lts a re discussed next onan individual basis.

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0 . 0 1 1.o 10 100 10002Cold Wall H e a t i n g R a t e ( B t u / f t - s e c )F i g . 8 - R e c e s s i o n R a t e vs H e a t i n g R a t e D e s i g n Curves a n d TPS T e s t R e s u l ts

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Edle r S-Glass Laminated PhenolicThis mate r ial was tested in the sim ulated flight cold wall heat r at e2range between 20 and 248 Btu/ft -s ec .

matched the flight design curve of R = 0.01377 (4cw)1.257131. A t the higherheat r at e levels, the glas s rein forceme nt me lted and flowed, and tiny bubblesappeared under the outer plies. The ma ter ial form ed a very s tab le charwith no visible evide-ce of char recession. A l l of the test s amp les wer ebonded with epoxy t o a 0.125 in. aluminum substrate using EA 934 adhesive.Some of the m odels showed evidence of bondline fai lu re and on one prob emodel the phenclic spec ime n fel l off after t es t completion. On the flightvehicle, the phenolic will be mechanically attached so this should not be aproblem.pending on the test conditions.predicted flight values.

The virg in mater ia l recess ion ra te

Measured su rfac e tem per atu res varied fr om 2230 to 3000 F de-These tempera tu res a re represen ta t ive of

0 P50 Sheet CorkThis mater i al was tested i n the sim ulated flight cold wall heat r at e range2between 23 and 142 Btu/ft -sec.

below the 2000 R design values by about a factor .sf 4.ra te was s imi l ar to the phenolic recession rate.which describes the data fairing shown in Fig. 8 is R = 0.05279 qcwinater ial formed a ver y stable craze d char with no visible char recession, infact, the material swelled a bit.reason for the low recess ion ia tes .f r o m 2273 to 3100 F depending on the heat ra te level.

The virgin mate r ial rece ssion rat e was wellThe mater ia l recess ion

The recession rat e equation* 0 . 9 9 895 . The

This stable char was probably the mainMeasured sur face temp era tures var ied

0 B-Stage CorkThis mater ia l was tested in the simulated flight cold wall heat r ate range

2between 21 and 166 Btu/ft -s ec .lowest of al l the ma teria ls tested.in Fig. 8 were obtained fro m Ref. 5.

The virgin m ater ial rec ession rate was theThe recession rate design values shown

In this test, the virgin material recession

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rate was approxixately 62% that of the P50 sheet cork. The B-stage c 3 r krecessio n rat e equation which best des cribe s the data fairing in Fig. 8 isR = 0.0447 4,,0.928. The B-Stage cork rcscted to the thermal environmentvery s imil ar to the P50 sheet c ork with possibly slightly mor e swellingoccurr ing.depending onthe heat ra te level.

Measured su rface te mp era tur es varied fr om 2338 to 3047 F

0 MTA - 2 Marshall Trowelable Ablator- 2This mater ia l w as tested in the simulate d flight cold wall heat ra te2rang e between 21 and 138 Btu/ft -s ec . This mater ial w a s developed a s a

c loseout mater ia l fo r ei ther M A - 1 o r cork . A s such, its recess ion ra teis expected to be similar f o r a good closeout mater ial . The virgin mate r ialrecession rate was w e l l below the MSA- 1 and MTA-2 (Ref. 6) design values,and about one-half the P50 co rk design values.twice a s fas t as the P50 cork.the data fairing in Fig. 8 is R = 0.3037 qcw-76455. This material did not ex-hibit a stable char formati on during the test. The cha r continually spalledoff as evidenced by a pulsating surface te mp era tur e history and hot spark scoming off the model.ft -s ec the surfa ce tem per atu re varied between 1940 and 2350 F, and at 136Btu/ft -s ec the surfac e tem pe ra tu re vari ed between 1885 and 2900 F, v e r -aging abou t 2260 F under the last condition.

In this test, it receeded aboutThe reces sion ra te equation which desc ribes

A t a simulate d flight cold wall heating r at e of 77 Btu/22

0 M S A - 2 - Marsha ll Spray-On Ablator-2This m at er ia l is one of s ev er al types being developed by MSFC to re -

place MSA- 1 and cor k TPS ma ter ial s.a sprayable ma ter ial which would eliminate t laborious task of bonding cor kin the areas that MSA-1 will not stand up to the exposed thermal environments.I n these test s , the mater ial de monstrate d that i t could be a di rect r -pla cementf o r P50 cork and MSA-1.rates slightly lower than the P50 cork design values.

This m ate ria l could be developed into

The m ater i al exhibited virgin m ater i al recessionThe recession rate

25

LOCKHEED - HUNTSVILLE RESEARCH b ENGINEERING CEIiTER


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LMSC-HRECTM D39771-6

* L c2927,equation which de scribes the data fairing shown in Fig.during tes t. The material formed a stable char. Surface temperatures vrrizdfrom 2374 F at a heating rate of 74 Btu/ft -s ec to 3041 at a heating rate C C 1272Btu/f t - sec.

i s R = 1.3931 a m * -"tv2

26

LOCKHEED .HUNTSVILLE RESEARCH & ENGINEERING CENTER.I------&ll-llllll -I_ _ -III - , -_IL_ILdIILLI___.L.-l . ~ - "


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4. CONCLUSIONS

For the range cf tes t conditions investigated in this study, the 1P: rnd-te r ia l samples performed as expected o r much bette r than expected.range of heating rates investigated were from 20 to 240 Stu/ft -bet and shears t r e ss l ev el s f ro m 0 . 0 0 8 to 0.021 psia. The aerL*inerm P G facility providedvaluable data for this investigation, and extended the range of applicability ofthe design recession cu rves for the various T PS materials. However, highers h e a r stresses and heating rates a r e s t i l l requ i red to cover the completeplume wash range on the SRB vehicle.to cover this range a s well as simulate the SSME plume wash che mic al species.The Acurex APG acility is a likely candidate facility in which such a futureexperimental program can be conducted.

The-

Future investigations should attempt

27

COCKHEED .HUNTSVILLE RESEARCH L ENGINEERING CENTER


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REFERENCES

1. Arnold, L, "Testing of SRB-TPS Mater ia ls in an Ar c Heated NitrogenEnvironment," Acurex Final Data R e p o r t 79-355, Ac ure x Corpora tion/Ae rot her m, Mountain View, Calif., May 1979.2. Youngblood, W. ., SSME Plume Impingement Environments to SRBsDuring Design Ca se Separation," Northrop Servi ces Inc., MemorandumM-9230-76-51 to Dr . T. F. Greenwood an d Dave Sey mo ur, NASA-MSFC/ED33, 22 July 1976.3. Youngblood, W. ., Design Environments -SSME Plu me Impingement toSRBs During Separation,I1 Nor throp Se rv ic es , Inc., Memo randum M/9230-76-60 to David Sey mou r and Dr. T. Greenw ood, NASA-MSFC/ED33, dat ed2r) September 1976.4 . Boison, J. C., and H. A. Cu rti ss, "An Expe rime ntal Investigation of BluntBody Stagnation Point V e l o c i t y Gradient," ARS J . , February 1959, pp. 130-135.5. Karu, Z.S. , "Space Shuttle SRB 3-Stage Cork TPS T est and Evaluationin AEDC Tun ne l C," LMSC-HREC T N D697584, Lockheed Missiles &Space Company, Huntsville, Ala., 22 Jun e 1979.6. Karu, Z. S ., "SRB TP S Closeout Mat eria ls Characterizat ion,@'LMSC-HREC TN D697757, Lockheed Missiles & Space Company, Huntsville,A l a . , November 1979.

28



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1 APPENDIX 1SRB THERMAL PROTECTIONSYSTEMS MATERIALS TESTNITROGEW ENVIRONMENT

RESULTS IN AN ARC-HEATED

LOCKHEED HUNTSVILLE R E S E A R C H L ENGINEERING CENTER


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Acurex Project 6945

TESTING OF SRB-TPS MATERIALS INAN ARC HEATED NITROGEN

ENVIRONMENTL . Arnold

Acurex Corporation/AerothermAerospace Systems Division

485 Clyde AvenueMountain View, California 94042

May 1979

ACUREX FINAL DATA REPORT 79-355

Prepared forGeorge C. Marshall Space Flight CenterCode: EP44Marshall Space Flight Center, Alabama 35812

NASA Contract NAS8-3401


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Acurex P ro je ct 6945

TEST ING OF SRB-TPS MATERIALS I NAN ARC HEATED NITROGENENVIRONMENT

L. Arnold

Acurex Corporat ion/AerothennAerospace Systems Division485 Clyde AvenueMounta in V iew, Ca l i fo rn ia 94042

May 1979

Acurex Fina l Data Report 79-355

Prepared forGeorge C. Mdrshal l Space F l i g h t CenterCode: EP44Mars hal 1 Space F1 g h t Cen ter, Alabama 3581 2

NASA Contract NAS8-33401


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TABLE OF CONTENTS

S ec t ipn. . . .1 INTRODUCTION . . . . . . . . . . . . . . . . . . . . . 1

. . . . . . . . . . . . . . . . . . . . 1.1 Ob jec t i v e2

3

4

5

. . . . . . . . . . . . . . . . . . . 1EST DESCRIPTION2.1 F a c i l i t y D e s c ri p ti o ns . . . . . . . . . . . . . .2.2 Test Models . . . . . . . . . . . . . . . . . . . 12INSTRUMENTATION 6. . . . . . . . . . . . . . . . . .3.1 Data Ac qu is i t io n and A na ly s i s . . . . . . . . . .3.2 Ar c Chamber Pr es su re . . . . . . . . . . . . . .3.3 Hea t ing Rate . . . . . . . . . . . . . . . . . .3.4 Back wall Temperature . . . . . . . . . . . . . .3.5 Model Surface Temperature . . . . . . . . . . . .3.6 Bu lk Entha lpy . . . . . . . . . . . . . . . . . .3.7 Model Sur face Pressure . . . . . . . . . . . . .3.8 Model Stagnat ion Pressure . . . . . . . . . . . .3.9 Center1 ne Entha lpy . . . . . . . . . . . . . . .3.10 Camera . . . . . . . . . . . . . . . . . . . . .TEST ING . . . . . . . . . . . . . . . . . . . . . . .

6667791010101111

4.1 Tes t Ma t r ix 1111.2 Te st Procedures . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . .1 7ESULTS . . . . . . . . . . . . . . . . . . . . . . .

PRECEDING PAGE BLANK NOT FILMED

V


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1. I TROPUCTIONT his r e p o r t pr e se n ts t h e r e s u l t s o f t e s t i n g 23 SRB-TP5 material

specinens f o r Marsha l l Space F l ig h t Center under NASA cont rac t numberNAS8-334ul and Acurex/Aerothenn contract number 6945. The contractnmn i t o r was Mr. W i l l i a m Baker o f NASA, a n d t h e o n s i t e t e c h n i c a l m o n i t o r wasM r . C a r l Wojciechowski o f Lockheed Mi ss i le s & Space Company, In c . Thetes t s w ere c onduc ted i n t he 1 MW Arc P lasma Generator (APG) f a c i l i t y o fAcurex /Aerotherm f rom 16 A p r i l 1979 t o 27 Apr i l 1979 .1.1 Objec t i ve

The o b je c t i v e o f t he p rogram was t o t e s t t h e SRB-TPS m a t e r i a lspecimens i n a hi gh h eat ing and h ig h entha lpy env i ronmen t under twoc on f i gu ra t i ons .upper f orw ard c o rne r l i p o f t h e TPS where the TPS i n t e r f a c e s w i t h t h e t o pof t he a t ta c h r i n g o r k i c k r i n g o f t h e SRB.s im u la te d t h e f l i g h t h e a t i n g e f f e c t s on t h e s e l f - s u ? p o r t i n g TPS areas onthe forward web o f th e SRB k i c k r i n g and a t t a c h r i n g .2. TEST DESCRIPT ION

The p robe c on f i gu ra t i o n s imu la ted t h e h ea t i n g on t he

The panel conf igurat ion

The ma ter i a ls th a t were tes te d i n t h i s program were P50 cork, gla sspheno l i c , "6 " c ork , MTA-2, and MSA-2. A l l o f t h e m a t e r i a l s w e r e t e s t e dunder both model con f ig ur at i on s .2.1 --F a c i l i t y O e s c r i p t i o n s

Th is t e s t program was condu cted i n the vacuum chamber of theAcurex/Aerotherm 1 MW Arc P l a s m a Genera tor (APG) f a c i l i t y l oc a ted i nMountain V i e w , C a l i f o r n i a . B r i e f l y , t h e VAC-APG produces a h i g h e n t h a lp y ,low pressure s t ream us ing a s uba tmos pher i c p res s ure t es t s ec t i on .vacuum i s prov ided by a f i v e - s t a g e s t e a m e j e c t o r .

TheThe APG i n pu t power i s


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suppl ied by a 600 kW c on ti nuous ra ted , s a tu rab le c o re reac to r , DC r e c t i f i e rpower supply.forms 460 VAC, 60 H Z i n p u t v o lt a g e i n t o a usable DC ou tpu t v o l t age . Thepower supply can prov ide 1.25 MW f o r s h o r t p e r i o d s of t ime.

This power supply uses a r e c t i f i e r t r a n s f o r m e r w h i c h t r a n s -

A l - i n c h d i a m et er c o n s t r i c t e d a r c h e a te r , c o n s i s t i n g of tw osegmented co n s t r i c t or packs 13.5 inches long, was used f o r th i s t e s tprogram. The te s t nozz le had a 0.75- inch th ro a t diameter w i t h a 8.5"h a l f a n gle l ea d in g t o t h e e x i t d i am e te r o f 2 inches.p robe, and c a lo r ime te rs were moved i n and ou t o f t h e t e s t s t ream us ingone o f t he t h ree w a te r-c ooled , pneumat i c a l l y c o n t ro l l e d s t i ngs .2. 2 leest Models

The models , p i t o t

A t o t a l o f 23 spec imens w ere t e s ted i n tw o d i f f e re n t modelcon f igu rat i on s . The probe t e s t specimens su pp l ied by NASA were mountedi n t o a graphi te model ho lde r , as shown i n F igure 1 , and at tached t o thes t i n g .t h e t e s t s t r e a m w i t h a s tando f f d i s t anc e 1 i n c h f r o m t h e n o z z l e e x i t .

The s t i n g was ad ju s ted pe rpend icu lar t o t h e c e n t e r l i n e f l o w o f

The probe specimen i n F igur e 2 shows the shape and size of the specimensbei ng tes ted.

The pane l s pecimen s hape and s i z e a re i l l u s t r a te d i n F igu re 3.The panel te s t specimens sup pl ied by NASA were mounted i n t o a coppern o de l h o l d e r w i t h t h e l e a d i n g a nd t r a i l i n g edges p r o t e c t e d by g r a p h i t ese ct io ns as shown 'n Fig ur e 4.c e n te r l i n e t e s t s t ream f l ow came i n t o c o n tac t w i t h t he s pecimen 5 /8 i nc hback o f t he l ead in g edge w i t h a s t a n d o f f d i s t a n c e 1 i nc h f rom the noz z lee x i t .

The st i n g was po s i t io ne d so t h a t t h e

The specimen/holder was incl ined 30" t o t h e f lo w c e n t e r l i n e .

2


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Graph i teu r o t e c t i v e

".ct\, .... . 1-OF POOH Q JA L I l Y

F igure 1 . The SRB-TPS model ho ld er assembly f o r th e probe cnnf iqurat ion.

3


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ORIGINAL PRGZ EOF POOR QUALIW

0. !O i i ia;er ia: th icknessa 1m i numback p ! a t e

T h e m m o t p le s

Figure 2. The probe specimen dimensions.

/0.187 + i a -4 lo les

Fiqure 3. The panel specimen dimensions.


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\ I /\ i lL0cehvnEalIn

5


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3. INSTRUMENfAT IONThe f o l l ow in g i ns t ru me n ta t i on was used t o c o l l e c t the da ta

r e fe r r e d t o l a t e r i n t h i s r e po r t.3.1 D ata Ac qu i s i t i on and Ana l y s i s

A l l data was co l le c t e d by the Y idar h igh-speed 80-channel d i g i t a ld a t a a c q u i s i t i o n s ys te m with a magnet ic tape record ing.i n c l u d e s a r c c u r r e n t an d v o lt a ge , a r c h e a t e r c o o l i n g w a te r mass f l o w andta i rperature r i s e , ar c chamber pressure, p i o t pressure, py rometer outpu t ,ca lo r im et er values, and thermocouple s ig na ls . The tape was processedthrough an Rcurex computer program t o gi ve power outpu ts , ar c losse s,bu lk entha lp ies , pressures , and temperatures .

The data

I n ad d i t i on t o t he magnet ic t ape , a H oney we ll 1858 V i s i c o r de rwas used t o r e co r d c e r t a i n t e s t d 3 ta f o r i m n e di at e a n a l y si s .?arameters recorded on the v i s i co rd er were pressure, thern ocou ple responses,arc cur re nt , and py rometer and ca lor im et er outputs .3 . ~ Arc Chamber Pressure

Some o f t he

A B e l l and Howell 0-25 ps ia pressu re t ransducer, type 4-326-0003,was used t o measure th e nozz le s ta gn at io n pressure i n th e p lenum ups treamo f t he 0 .75 - inc h d iamete r t h r oa t . The p res s ure t rans ducer ou tp u t s i gna lrJas am p l i f i e d by a B e l l a nd Ho we ll 8 -1 14 s i g n a l c o n d i t i o n i n g u n i t b e f o r ei t w en t t o t h e V id a r f o r r e c o r d i n g .3.3 Heat ing Rate

A s l ug cal or im et er , s i m ul at in g th e probe specimen shape, was usedThe ca lor imetero m easure t h e h e a t i n g r a t e o f t h e p r o b e c o c f i g u r a t i o n .

was a 1.25- inch f l a t f ac ed s l u g w i t h a 0 .06 - i nc h c o rne r rad ius .c a l o r i m e t e r was i ns e r t ed i n t o t he a rc s tream and wi thd raw n a f t e r 1 t o 2seconds had elapsed a t t h e c e r l t e r l i n e o f t he a rc f l ow .

The

6


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A t h i n - s k i n c a l o r i m e t e r , p r o v id e d by NASA t o s im u la te t he s i z e and s hapeof the panel spec imen, was used t o co l l e c t the he at in g r a t e o f t he pane lc on f i gu ra t i on . The t h i n -s k i n c a lo r im e te r had t h ree t y pe K t h e m z o u p l c s spacedevenly across the panel back face and of fset on both s i c k s from t h e c e n t e r l i n eby 3/16 inch.c a lo r ime te r oppos i t e each o f t he t hermocoup les and o f f s e t from the c en te r l i ne .A Gardon gage c a l o r i m t e r was u sed on t h e t r a i l i n g e dge o f t h e t h i n - s k i ncalorimeter t o provide a secondary measurement o f t h e h e a t i n g ra t e .Model C-1117-GX-60-120, s e r i a l number 44118. The th in - s k in ca lo ri m et er , show0i n F igu re 5, was i n s e r t e d i n t o t h e a r c f lo w an d h e l d f o r 3 seconds a t thec en te r1 ne be fo re be ing w i thd raw n.3.4 -- ackwall Temperature

Three pressure taps were a l so spaced even ly aamss th e t h i n -s k i n

This was a

Both the probe and th e panel specimens were ins t rumen ted w i t h ? d- m i l l ,t ype K (chrome l-alumel) thermocouples. The probe co nf ig ur at io n had one thermo-c oup le a t t ac hed a t t he c en te r o f t h e s pecimen and ano the r o f f s e t abou t 1 / 4 i n c ht o one s ide as shown i n Fig ure 2 . The panel co nf ig ur at io n had thr ee thenno-c oup les ev en l y spaced from the l ead ing eds e o f t he s pec imen t o t he t r a i l i n gedge on the speiiiwn's c e n t e r l i n e as shown i n F ig ure 3 .3.5 Model Surface Temperature

Ff i r t he p robe c on f i gu ra t i o n , a f i b e r op t i c py romete r was us ed t o r e c o r dThe pyrometer was a Vanzetti Model 1317-1 185-8-0H2,he sur face temperature.

se r i a l number 101719, w i th a 3- inch fo ca l le ng th and an e f f e c t i v e spotd iamete r o f 0.035 I nc h . The sens i t i ve range i s 0.7 t o 0.97 microns .

inches f rom thewas 1400F t o 4500F.

TD-9F was use d t o

The

an e f f e c t i v e s p o t

pyrometer was mounted on the n o z z l e and pos i t i nned ?specimen. The tem pera ture range

c o n f i g u r a t i o n , a Therrnodot Modetemperature. Th is py ron eter has

7

c en te r G f t h e probeFor the pane

measure the surface


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00c?

ulNLN T-n

.--IrnN7 -

-1'

8


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diameter o f 0.076 i n c h w i t h a s e n s i t i v e r a n g e o f 0.75 t o 0.90 microns.The pyrometer was p os i t i on ed on th e ou ts i de o f the vacuum chamber lookingth rough a qua r t z window , w i t h t he e f f e c t i v e s po t p l ac ed w here t he t e s ts t ream c en te r l i ne f l o w h i t t he pane l s pecimen .was :750F t o 2600F.3.6 B u j k Entha lpy

The range of the py rometer

The enth alp y o f th e gas was determ ined by an energy balan ce o f theAPG system, in c l u d in g th e ar c column f rom th e cathode t o th e anode, plenum,and nozzle. B u l k e n t h a lp y i s d e f i n e d a s :

whereI = Arc current (amps)V = Vol tage dropped f rom cathode t o anode (v o l t s )

= Mass f l o w ra te o f t he c oo l i ng w a te r t h rough t hemHZ0 arc, plenum, and noz z le ( lb /s ec )= D i f f e r e n c e i n t h e t e m p e ra tu r e be tw ee n th e i n l e tATH'20

m

and th e o u t l e t c o o l i n g w a t er f o r t h e a r c ,plenum, and no zz le ( O F )= Total mass f l ow o f a rc hea ted gas ( l b / s ec )tgas

Water f lowrates were measured wi th an ASME s harp edge o r i f i c e and 9d i f f e r e n t ia l pressure t ransducer . Temperatures were measured w i t h ad i f f e r e n t i a l t h er m o p il e . Gas f lo w was measured on two flowmetersc a l i b r a t e d f o r a f i xe d pressure, temperature, and f lo w range.

9


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3.7 Model Surface PressureThree Statham 0-5.15 p s i a press ure t ransduce rs, typ e PA 732 TC-5.15-350,

were used t o measure th e model surface pressu re. The 1/16- inch diameterp res s ure pov ts w ere l o c a ted on t he t h i n - s k i n c a lo r im e te r , o f f s e t 3 / i 6 i n chf ro n i t h e c e n t e r l i n e a nd pl a ce d a l t e r n a t e l y on each s i d e o f t h e c e n t e r l i n e .A smal l tube was ru n f rom the pressure p or ts t o the pressure t ransducers .The s u r f ac e p res s ures were t hen t ak en du r i n g t he c a l i b ra t i on runs f o r t hep a n e l c o n f i g u r a t i o n .3.8 Model Sta gna tion Pres sure

A p i t o t tube was used t o measure the s ta gn at io n pressure on thec e n t e r l i n e . The 3/8- inch d iameter p i t o t probe had a 1 /16- inch opening.The p i t o t probe was connected t o a Statham 0-15 ps ia pressure t ransd ucer,type P68-15A-300, which obta ined th e pressure re ading . V i s i c o r d e r r e c o r d so f t he p res s ure respons e w ere used t o ens ure t h a t a s t eady -s ta te c on d i t i o nhad been reached before removal from the gas stream.3.9 Ce nte r l in e Entha lpy

The cen ter1 ne enth a lpy was ca lcu la t ed us ing the measured quant i t ie so f model s t3gna t i on p ress ure, t he c o ldw a l l he a t i ng ra te , and t he f o l l o w in gZ3b.y equa t i on f o r N2 (Reference 1) :

4cwJGh ( B t u / l b j = -E 0.0431 dp,2

where4,, = C oldw a l l hea t f l u x f rom the 1 .25- inc h f l a t f aced

Pc a l o r i m e t e r ( B t u / f t 2 s e c )

= Model s ta gn at io n pressure (a tm)t2JReff = 0.421 f o r t h e c a l o r i m e t e r c o n f i g u r a t i o n ( f t 1/21

10


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3.10 CaineraThe camera used t o r ec or d the models under te s t co nd i i io n s was a

Locam camera w i t h a 50 mm l e n s .The f s t o p was e i t h e r 16 o r 22, depending on t h e t y p e o f f i l m be ing used.

Th re e d i f f e r e n t k i n ds o f f i l m were used. The f i r s t was th e Kodak

The speed was s e t a t 10 0 frames a second.

Plus-X Reversal F i l m 7276 w i t h an AS A o f 50. The second f i l m was EastmanEktachrorne Video News F i l m 7240-Tungs ten w i t h an ASA o f 125.of f i l m use d was Eastman Ekta chrom e Vid eo News T i l m , Hi gh Speed, 7250-T ungstenw i t h an ASA 400.4. TEST1NG4.1 ---est M a t r i x

The l a s t t yp e

T a bl e 1 l i s t s t h e t e s t se que nce f o r t h e m od els , t e s t d u r a t i o n , t e s thea t ing c on d i t i on s , and m odel c on f i g u r a t i on .

The c o l d w a l l h e a t i n g r a t e d i d n o t s i m u l a t e t h e s u r f ac e t em p e ra t ur eand t h e h ea t l o a d as c l o s e l y t o a c t u a l f l i g h t c o n d i t i o n s as had beenexpec ted. The ho twa l l he at in g ra t e , however , was determ ined t o c l o se lys i m ul a te ac t ua l f l i g h t c o n d i ti o n s . By t a k i n g t h e h o t w a l l h e a t i n g r a t e ,b o t h t h e f l i g h t s u r fa c e t em p er at ur e a nd t h e f l i g h t h e a t l o a d c o u l d bes im u la ted us ing t he a r c hea te r . The s im u la ted c on d i t i on s were ob ta ine dw i t h an N2 t e s t g as i n a s b p e rs o n i c s tr ea m h a v i n g a minimum Mach numbero f 2.5.4.2 Test Procedur?;_ _ _ _ ~

Cal ib ra t ion runs were made a t each t e s t c o n f i g u r a t i o n b e f o r e t h emodel runs were made. The c a l i b r a t i o n sequence was as fo l l ows :

1 . Hook up a l l i n s tr u m e nt a ti o n t o t h e s l u g o r t h i n - s k i nca 1o r i meter .

11


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TABLE 1 . TEST MATRIX OF SRB-TPS MATERIALS

PC-2PA - 3PA- 4PA- 5PA- 6PB-1FB-2PD-1PE- 1PE -2

Heat ing TimeRun No. C onf i gu ra t i on Mode l No. C o n d i t i o n ( s e c )L

1234567891011121314151617181920212223

H iLoLoH iH iLoH iLoLoLoLo1.0Lo:4 iH iH iH iH iH iH iH iH iH i

4444885485588151 52515151520252525

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2.

3 .4 .

5.

6.7.0 .

9.

1 0 .

11.

1 2 .

1 3 .

Calibrate al l the modules on the 1858 Visicorder with aknown voltage.Pump down t he a r c vaciiuni chamber.When chamber i s pumped down t o desi red pressure , zero thet ransducers .Check vacuum chamber and al l inst rumenta t ion l ines f o r1ea kage .Set gas l ine pressure.Cold flow gases t o ensur e proper mass f low rate .S t a r t arc on a r g o n , turn magnetic tape on , and switch nitroqznon when the arc current is 100 amperes above the desiredt e s t p o i n t .After switching t o ni t rogen, lower the arc current t o thet e s t p o i n t .Inse r t t he p i to t p robe ( Pa t t he cen te r l ine O F the flow f o r 2 seconds : , i t h 1858 Visicord err u n n i n g a t 1 i p s , and then h i t h d r a w i t .With the 1858 Visicorder r u n n i n g a t II J i p s , i n se r t t h et h i n - s k i n calor imeter into the arc flow and leave i t a t the flowcenter1 ine for 3 seconds before withdrawal.W i t h the 1858 Visiccrder r u n n i n g a t 10 ips , i nse r t t he s lugcalor imeter i n t o the arc stream f o r 1 t o 2 seconds beforewithdrawal .Using the information on the magnetic tape, run a computerprogram to yie ld the d a t a fo r the arc condi 'ions; analyze thed a t a from the 1858 Visicorder, and record the re su l t s .

) i n t o the arc f low, s ta t io; iaryt 2

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14. R epeat t h i s p roc ess s ev era l t imes t o ens ure opera t i o n a t t hed e s i r e d t e s t p o i n t i s obta ined.

The procedu re used when t e s t in g th e models was as fo l l o w s:1.2 .3 .4.5.

6.

7.8.9.

10.11.

1 2 .

Take p re te s t photograph.Record pretes t weight .R ecord p re te s t t h i c k nes s .Connect model thermocouples t o the record ing sys tem.Sec ure l y moun t model t o t he s t i n g w i t h t he p robe c en te ron t he c e r i t e r l i ne o f t he a rc , 1 i nc h away f rom the noz z lee x i t , o r t h e p an el i n c l i n e d 30" t o t h e a r c f lo w , w i t h t h e a r cf lo w c e n t e r l i n e h i t t i n g t h e model 5 / 8 i n c h f ro m t h e l e a d i n gedge o f the spec imen; the panel s t ag na t ion po in t had a l - i n c hs t a n d sf f d i s t a n c e f ro m t h e n o zz l e e x i t .Check a l l i n s t r u m e n t a t i o n t o e n su re t h a t i t i s w or kin gproper1y .Lower the vacuum chamber pressure to the des i red level .Cold f l ow the gases t o ensure proper mass f lo wra te .S t a r t the arc and sw i tch over to n i t rogen at 100 amperesov er t h e des i red c u r re n t .L o w e r t h e a r c c u r r e n t t o t h e t e s t s e t t i n g .k i t h th e 1 858 V is i c o rd e r t u rn e d on a t 1 0 i p s , i n s e r t t h e s l u gc a l o r i m e t e r i n t o t h e a r c s tr ea m f o r 1 t o 2 seconds and thenwi thdraw i t .With the Locam camera and the 1858 V i s ic or de r turned on, in se r tt he p robe o r pane l moae l i n t o t h e a rc s t ream w i t h t he du ra t i o no f t he t es t s t a r t i n g when t he model reaches t he a rc f l o wc e n te r l i n e as shown i n F igu res 6 and 7 .

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EOVWnnec,W00E

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E0U.r

LaJ&IrnaJcUmc

It-4

mU

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13. A ft e r ru nnin g the test . , b r i n g the vacuuiit chamber up t oa tmosphe r ic co nd i t i o n dnd rwovc ! the rade l f ro ii i t he s t i n g .

1 4 . Give t he iiiode l t o the ons itc . t ech n i ca l r ep re sen ta t i ve f o ro b s er v a ti o ns o f t h e c h a r r i n g and spa11 n g ( t h e p o s t - t e s tphotog raphs , p os t - t es t t h i ckness , and po s t - te s t we i gh t w i l lbe taken a t NASA) .Repeat steps 1-14 f o r each te s t model .5.

5. RESULTSTable 2 p r e se n t s t h e t e n c a l i b r a t i o n r u ns t a k e n f o r b o t h t h e p a ne l

and the F rsbe con f i gu ra t i ons .p ro be c o n f i g u r a t i o n t o de t e rm i n t h e s t a g n a t i o n p re s su r e, c o l d w a l l h e a t i n gra te , a rc c ur r en t , and chamber p ressure .c e n t e r l i n e e nt h a lp y , c e n t e r l i n e t em p e ra tc r e, an d h o t w a l l hea t inc j ra te wereca l cu l a ted . The ho t wa l l / co l dw a l l co r re c t i on was made us i ng th e equat ion ,

Seven cal ibrat ion runs were made f o r t h e

From t h i s i n f o r m a t i o n , t h e

Based on the ac tua l f l i g h t da ta , a su r face tempe ra tu re o f 3500'F wasassumed fo r t he ho twa l l , co r respondi ng t o an en tha l py of 1330 B t u / l b .

Thre e c a ? i b r a t i o n r u ns were made f o r t h e p a ne l c o n f i y ~ r a t i o n . Thec a l i b r a t i o n r u ns p ro v i d e d i n f o r m a t i o r ! o n t h e c o l d w a l l h e a t i n g r a t e , c e n t e r -1 ne en tha l py , ce n te r l i n e i empe r i tu re , a r c cu r ren t , chamber p re ssw e , andsu r face p ressu re ac ross the pane l . From the th i n - s k i n c a l o r i m e te r , t hec o l d w a l l h e a t i n g r a t e and t h e h o t w a l l h e a t i n g r a t e a t two temperatures,3500F and 2000F, were found by ca l cu l a t i ons .

A t q t a l o f 23 specimens suppl ied by NASA were t es t ed du r i ng theprogram. Twelve Specimens were te s te d i n t h e p ro be c o n f i g u r a t i o n . O f

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TULE 2. CALIBRATION RUNS FOR THE PROBE AN D PANEL CONFIGURATIOPiSOF THE SRB-TPS TEST PROGRAMU Probe

3148-02 321 509 143 w o 9160 . i s 3 0.978 113314801 522 48t 200 4373 lrSl0 . l a 8 1.136 1683150-til 224 514 95 2447 8019 -1438 0.888 65-02 520 402 174 3982 !M9t .lsl 1.133 142-02 520 482 23 3982 9491 - 1.133 -

-03 525 484 190 4335 96% .la 1.142 158-03 525 48) 19 4335 9695 - 1.142 --04 224 514 07 2244 7657 .1453 0.886 583151-01 520 461 180 4069 9509 . I875 1.136 147-01 SM 481 19 &9 9509 - 1.136 -

5567

70

1445

16 n19,s

11 .q

FOLDOUT FRAME118


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U45

43

l a t = 45 51 35 41 13 15 0.08719." 56 63 9- 42 16 18 0.068

ProbeProbeProbe0.005 PanelProbe0.005 PanelProbePiobe

11.Q 58 66 3 9 4 4 10 11 0.0791 0.0213 0.0099 Panel

FOLDOUT FPAMr


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those 12, three soecimep were tes ted a t t he low hea t l oad s imu la t i ng pcond i t i ons .5 i n tu l a ti n g f l i g h t c o n d i ti o n s.

?he remaining n ine specimens were te ste d a t th e h igh heatThe r es u l ts a re shown i n rab ies 3 ana 4.

umeoad

The other 11 specimens were tes ted i n the pane l c o n f i g u r a t i o n . Theywere a l l t es ted a t t he h igh hea t l oad t o s i m u l a t e f l i g h t c o n d i t i o n s .res u l t s f r oi ii the pane l co n f i gu ra t i o n te s t s a re a l so shown i n Tables 3 and 4.

The

In Table 3 , the exposure t im e s ta r te d when th e t e s t mode l en teredinto the test st ream and ended when the model was wi thdrawn f rom the teststream. The ce nt er l in e temperature i s t h e a c t u a l t i m e t h e t e s t model wasa t t he cen te r l i ne o f t he tes t s t r eam.

The ho twa l l hea t i ng ra te i n Tab le 2 a n d t h e t e s t h o t w a l l h e a ti n g r a t ei n T ab le 3 were f rom the probe calor imeter .i n T ab le 2 was from the probe ca lo r imeter .

A l s o t h e c o ld w a l l h e a t i n g r a t e

All o f the backwal l tem?eratures i n Table 3 were taken f rom th eVis icorder t races made dur ing the exposure t i m e o f t h e t e s t m odel.

The po st - tes t photographs o f a l l specimens, ce r t a i n pa st - tes tweights, and a l l t he pos t - t es t t h i cknesses w i l l be taken a t NASA. TheDre tes t photographs and the V is icord er t race s o f the t e s t runs were takent o NASA by t h e t e c h n i c a l on s i te m on i to r t o be ana lyzed.

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llodeln No. No.-52-01t C-1

-02 pc-2-03 PA-3-W PA-4

-OZ+ PB-1-02 pe-2-04 .-kt

pE-I- 0 2 E - 2-0 3 -2c-2

-02t A-1-03t A-2-wt 6-1

-02t D1- 0 3 E- 2-041 D2

TABLE 3 .

Config.

RESULTS FROM TPE TEST!G GF THE SRB-TPS MCITERIALSI N TH E AR C P L A S M GENEiVITOR

#ttri a1PanelPmk

7

Pane 1

v

p50P50P50phtnollcFileno11c

PhenolicPhcnol c-8" cork"8" corkI M - 2UTA-2UTA-2I M - 2

P50PhmollcPhcnolic"6" corkrnA-7I M - 2M A - 2I M - 2

PSO.Ern cork

2345283631012912290222202988297830472991287030412897

2273242422672370232023742350234521002338

TBYl( O F )

93130159187.365216207122283138139307238

-

16325424 1165284166216264273268

TBW(OF)-82i 0165194'36520420511626712 8128270224106123151

X14 1

XXX

XX

Them were no n u k r 3 t h e m m u p l c s on the probe configurationThe thcrrpcouplc ms mt hooked upTo be nrasured o r weighted a t NASAMovies t r k m

3.M4-254.304.208 .28.26. 04. 38.35.55.38. 28.4

15.515.425.515.415.415.420.725.625.525.5-

4.254.45.15.459 .1a.85.755. 08.85.565.919.48.6

15.916.2726.115.716.015.6521.126.026.125.7

bill,(OF)-756869686876737175767474786768777476760384

9497

TBW2i(OF)-75686968697673767676747378666877X76XXX

XX

Dwbe

(Btu/ft*-sec 1648580139142

;I% s t

5714s107a471

100957455505352545358585448

FOLDOUT FRAMEI20


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Probe(Btu/ f t * -sec)%est

64858C13914257

1491078471

100957455505352545358585448

38.019.810.714.714.514.614.71o.m10.93510.21810.04511-06310.23938.42457.01556.437.438.934.838.834.938.237.5

POS! !g !37.6241'3.71014.71614.4011 5 . 7 3 3

14.75614.68510.50410.68910.078

37.07056.09255.14136.10235.97634.06334.17532.40935.14835.691

;re ( i n ;f6'7

. ~ .

.e90.42a-385- 385

.3e5-387- 428.429- 431-431-430-432-426- 390.385.429-430-429.429.426-428-429

Pos t ( i n )t

t

tt

t

t

t

ttt

t

t

t

t

t

t

t

RenarksBarelf scorchedGood c h a r buildup; 1 /32 i n char dep thGood , h 3 r bur ldup; 3/64 n char depthGlass fiou and o u t e r p l y bubbled; 1 /16 i n char dep thGlass floued and bubbles occurred; mdel dcbonded a f t e r t e s t; 3/32-1nchar depthNo glass f l o w o r bubbles; model looke d good; 3 /64 i n char depthk d c l looked good; g lass melted; 1/16 i n char depthModel looked good; gooG char; 1/16 i n char depth; no recessModel looked good; model suel!ed 1/32 in; good cchar depthModel iooked good; 3 /64- in ch ar depth; m e s s 1/32 i r .Char spal led of f ; Ts r facc var ied; rcasura ble wces sSaw hc t char sp al l ; ]Ysuf lace avg. 2255F; s l i g h t sur fa ce dinolenode l debonded a f t e r shutdown; good model (no s p a l l i n g ) ;recess = .M n; char = -0 6 i n6004 char bui ldup; f i n a l - - 32 th i c k; n i n i n u ch ar e ro s io n i nspots; -1 0 i n char depthSl igh t g lass mel t; debonded a f t e r test ; oood model; norecess; -05 i n c h a r d e p thS l i gh t g l ass m el t ; debonded af ter test ; good model; IIOrecess; -0 8 i n c h a r & p thGood char; :surface - 2330 t o 2370'FChar saal led of fGood m d el ; st ab le char; T2 and T3 were e l lm ina ted t o save time ;TsUrfa.1T2 and f 3 deleted to save tine; c h ar s o a l le d o f f ; m j o r su rf ac eGood stable char almost t o alu nin rm back; minor s urfac e e.T;SiOn;charred surface crazedGood char buildu p; surfa ce crazed; 76 = 320F I n 3.5 sec a f t e rShJtdovnAd21 , i lmor t a l l charred; TB = 387F a ft e r shutdown

-; 3/3?- in

237C"F; mino r cha r eros ion; cra zin gerosio7; Ts,rface 2642F

3 OLDOUT FRAME


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ORtGINAL Pi3'3s !s'OF POOR QUALITY

Ia:mmwr --.-

WI+LL0v,z0

Dz0uew-Im+

2u1- 3

-44\ '

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-.- - N ~ U , -NO- ,rum - N ~ U ma - N0 0000

hmC )m u 7m P ) n 3 o o

7 '3000 000 0000 0000 00, 1 1 1 I l l 1 1 1 1 I l l 1 I Imv,n m7zI I I I c- r n a~ . n m c- r- - 7

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RE FE RENCE

1. Zoby, E. V . , "Empirical Stagnation-Point Heat-Transfer Relation inSeveral Gas M i xtures a t High Enthalpy le ve ls I' National Aeronauti csand Space Administration, June 24 , 1968.

Ablator Recession Speeds

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