NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

Technical Memorandum 33-579

A Brief Survey of Carbon/Carbon RefractoryComposites at the Jet Propulsion Laboratory

E. Y. Robinson

(NASA-CR-131745) A BRIEF SURVEY OF N73-533

CARBON/CARBON REFRACTORY COMPOSITES AT THEJET PROPULSION LABORATORY (Jet PropulsionLab.). 19 p HC $3.00 CSCL 11G G.G3 /18 .69456/

JET PROPULSION LABORATORY

CALIFORNIA INSTITUTE OF TECHNOLOGY

PASADENA, CALIFORNIA

December 1, 1972

Reproduced by

NATIONAL TECHNICALINFORMATION SERVICE

US Department of Commerce

Springfield. VA. 22151 Lo?£.

https://ntrs.nasa.gov/search.jsp?R=19730013806 2020-07-01T06:09:15+00:00Z

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

Technical Memorandum 33-579

A Brief Survey of Carbon/Carbon RefractoryComposites at the Jet Propulsion Laboratory

E. Y. Robinson

JET PROPULSION LABORATORY

CALIFORNIA INSTITUTE OF TECHNOLOGY

PASADENA, CALIFORNIA

December 1, 1972

(

Prepared Under Contract No. NAS 7-100National Aeronautics and Space Administration

PRlEFACE

The wark descriihed iun ti report w/as perorffmed by he Afppeix

Mechanics Divisionim off the Jeit Propsim Iabotra iory Tnis report was pre-

selted at thet 2a e ireafrtolr Cn#iaposites Workin Grop MWeettiin, held in

P nb elbmd,, CMnuie Ocrt 19-20.

JPL Technical Memmoramluf 33-579 in

PRECEDING PAGE BLANK NOT FILMED

CONTENTS

I. Refractory Radiation-Cooled. Rocket Nozzles ...............

II. Integrated Propulsion Structure ........................

III. Atmospheric Entry Shells ............................

References ..........................................

TABLE

1. Static-firing test conditions and results ................

FIGURES

1. Configuration of all-carbon nozzle for feasibilityte sting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

2. Layup step in nozzle transition and body fabrication........

3. C/C nozzle during firing ..........................

4. Measured nozzle surface temperatures during andafter static firing of flight-weight motor ...............

5. Evolution of conesphere concept .....................

6. Conesphere motor ..............................

7. C/C sample (5-gr silver 0. 16-cm (1/16-in. ) standoff,low-density) after shaped-charge cutting: (a) sample 1,(b) sample 2 ..................................

8. C/C sample (15-gr lead sheath 0. 32-cm (1/8 in. )standoff high-density) after shaped-charge cutting .........

9. C/C sample (15-gr lead sheath, 0 standoff high-density) after shaped-charge cutting ..................

10. Typical entry shell configuration .

11. Schematic of small doubly curved developmentalaeroshells . . . . . . . . . . . . . . . . . .

JPL Technical Memorandum 33-579

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ABSTLRACT

Refractory composites specificaUlly of carbon-carbon iat¢erial, are

being consideered for application in: ((a)) rocket motor nomzales and s$kirtts,

(b)) a naique integrated prOpnlsion s$truc:tre, and (c) planetary atmospheric

enltry sheIlls.

The first appllication is inttended ffor radiattion-coo led rnozales and slkirits

whBich, with presently aail ble materials,, can meet operat.io.al reqnire-

ments off dee slpace missions at subs$tantiall lower weights.

The seconcd alppllication reqnires very high structural performance as

well as refractory capability, and i$s easible only if high-strength graphite

il]menmts are efficiently incorporated into a carbon-matrix composite.

The third application is comparable in many respects to earth-entry

aeroshells and heat shields; however, planetary atmospheres may pose new

gas-~djac amnd corrosion problems. Furthermmore,, te entry shenn is

likely to be an integral structraln element which i$s nott jettisoned, and to

which critical hard point attachment must be made.

Technical devewlopments and pa]]ns in each off tlhese areas are desacribed

J1PL Technical Memorandum 33-579

I REFRACTORY RAD ATION -COOL1EDROCKET NOZZLES

The feasibility of using all-carbon composite nozzles was demnonstrated.

at JPL in static firing of 27-kg solid propellant moQtors (Ref. ). The con-

figuration of these nozzles is shown in Fig. 1. The significant advancement

in long-buring, solid. propelalnt motor technology is reflected im:

(n1 ) Nozzle weights 0. 4 to 0. 6 that of equivalent flight-weight alblative

nozznles,

(2) Reusabililty ith minor, i any, refurblishment (demonstrated by

repeeated firing of the first nozzle), and

((3) Firing ties representative of space mission requirements.

The current nozzle design appears to be acceptable for projected 355-kg

motors, base. on thermal analysis. The nozzles shown in Fig. 1 were fab-

ricated, by the rosette-lLayup methi d, illustrated i~n Fig, Z. This process

appeared.3 at that timae, to be preferable lto filament-winding techniques and.

to offer improved accommodation of shrinkage during graphitization and.

reimpregnation.

During lthe first firing in air, the nozzle appeared. white hot for approx-

iiately 35 s of the total 47-s cycle. The nozzle is shown under these comndi-

tions in ig.. 3. alculated. thermal protiles, shown in Fig. 4, are consistent

with the properties of C//C andi. are similar to analytical estimates oDf larger

structres.

A suaa ry of tests p.erfried. with the sulbscale nozzl'es is presented.

in Table 1.

2JPL 'Technical Memorandum 33-579 I

Large-scale C/C nozzles are currently under evaluation at JPL. Con-

figurations are similar to the subscale specimens shown, with exit bell

diameters of 61 cm (24 in. ), overall height of 81 cm (32 in. ), an 18-deg angle,

and a throat diameter of approximately 7. 6 cm (3 in. ). A total of three such

large nozzles will be test-fired under various conditions during the first half

of calendar year 1973.

A lower level of interest exists for such nozzles with liquid-propellant

motors. Small-scale tests utilizing pyrocarbon nozzles were performed with

several propellant combinations, including fluorine-hydrazine, Flox-MMH,

and oxygen-fluorine-diborane. At present, only a small C/C insulating insert

is used in connection with Flox-MMH motor development, and long-term plans

for improved nozzle-bell materials are still in the formative stage.

II. INTEGRATED PROPULSION STRUCTURE

High-energy missions place a strong incentive on the weight-efficiency

(mass fraction) of upper propulsion stages (Ref. 2). A four-stage system,

shown schematically in Fig. 5a, contains weight contributions from inter-

stage thrust structure, surrounding fairings, as well as from motor cases

and accessories. Using the nozzles as thrust structure reduces the envelope

(and. firing weight) and eliminates preexisting thrust structure as shown in

Fig. 5b. Reshaping the motor case allows each stage motor case to be

nested within the nozzles of the next stage, further reducing envelope dimen-

sion as shown in Fig. 5c. The double wall of nozzle and motor case may be

replaced by a single dual-purpose shell, shown in Fig. 5d. This configura-

tion has been termed the "conesphere" concept, and is shown in Fig. 6. The

principle of operation is to separate, by a shaped explosive charge, the spent

stage at the point indicated on Fig. 6. The remaining portion of the motor

case becomes the nozzle skirt of the next stage, and so on.

The structural material must provide very high efficiency as a rela-

tively cool pressure-case and, subsequently, must resist high thermal stress

as well as temperatures on the order of 3238 K (5000°F). The most obvious

candidate material is a C/C composite. The pressure-vessel mode, as well

as the requirement to transmit thrust loading and resist launch forces, de-

mands good biaxial strength of the material. Preliminary review indicates

JPL Technical Memorandum 33-5792

that a 0/90 laminate is preferable to +0 orientations; however, a radically

different fabrication approach becomes necessary. In addition, the residual

stress states of 0/90 systems are more extreme.

Industry sources of C/C technology have been, and are being, can-

vassed to determine the prospects of achieving reliable high-performance

C/C structure for this application, using various material/process alterna-

tives and. orientation patterns. This survey activity is expected. to continue

at a moderate pace.

One of the critical feasibility questions associated with the concept is

that of reliable explosive separation. Some very preliminary evaluations of

shaped-charge separation were made at JPL. Effective cutting by a shaped

charge depends on explosive loading, sheath material, and standoff from the

target surface. These parameters must be optimized for each material and

plate thickness. While metal cutting by shaped charge depends to a great

extent on melting and vaporization processes, neither of these phenomena is

likely to be present with C/C composites. It is necessary to develop methods

which make effective use of shock and. abrasive action to achieve a reasonably

clean cut of C/C composites. Results of preliminary trials of shaped.-charge

cutting of low-density (-1. 3 g/cm 3 ) C/C are shown in Fig. 7. A slight

restraint, characteristic of a real structure, appears to be sufficient to pre-

vent extended delamination. Higher density material (~1. 7 g/cm 3 ) is shown

in Figs. 8 and 9. Higher charge densities were used. at two standoff locations.

These preliminary trials suggest that shaped-charge cutting of C/C is feasible,

and that fairly good control of the cut edge could be achieved, for low-density

material. Further experimental development is needed, but prospects for

achieving useful separation techniques appear good.

III. ATMOSPHERIC ENTRY SHELLS

Planetary probes involve high-velocity atmospheric entry and the

requirement for thermal protection of the payload, as well as maintenance

of structural integrity and predictability of flight path. The type of entry

shell considered at JPL is depicted. schematically in Fig. 10. General struc-

tural features shown are the optionally separate nose piece, skirt-stiffener

at the outer periphery, and. an intermediate ring for hard.-point attachment.

JPL Technical Memorandum 33-579 3

igniure I shows structural detilt o doe velometal comcepts Iased on light-

core samdliWich sctrlu re to ptimii e shelli flexural stiffes san. buand lig

resistanmice. Figure 11 also iimdicates a heat shieldi. mmnaterial hich may he up

to l. 25 cm ((. 5-imn. )) thLick . There is interest in the p otential of C//C integral

sandwich ffor a mutipurpose structure hich comn bines adequate strength ami.

thermmal stress resistance witfh payload insulatiom capaility amid resistance

to atmospheric corrosion.

lThere are, at presemt,a nom specific structure development or data gem-

eratiom programs ffor this applicationm at .PILk The state-of-the-art in conm-

plex C/c structulre, anml maofidied C//C afor corrosion protectilonl, is being

mDmomitbored.

REF~ERENCES

1. Bailey,, . L.,, amndl Shaffier,, J.. l"m 1An AM-Carhom Radiating N[ozl;e fforLang-Burnmiag Solid Propellant Motorsr,," JPL Quarterly TechnicalReviev,, VoaL. ,, No, 2,, p. 36, Jet Propnnlsiom Lahoratory, Pasadena,Califf,, JUlY l971

2. Naaura,, Y..o am d Shaffera, J. L,. Solid Propulsiom Advanced Comicepts,Tech}niical l2MemmlDorandumn 33-534,, Jet Propulsion Laboratory, Pasadena,Caiff. Mayl, 1l972.,

JTPL Technical Memorandum 33-5794p

All.:r-caDlEDo lDZ;Zmlce8

Albla'tiv'eTest I Test 2 Test 3 MZZ

EPrqeifallatitt oweight31 Bkg 6 265 542 2i6., (65 27* 5'l

SiaMulated1 altitbuide 11rak Sea level Sea level 15,.9 r0 15i,. '9(0

]MEolbor tbuDrmig tme,, s 47 45 4f8 2(0

Nozzle epansion. ratio CD 53. 5 4( 3 4(0,. 4 35

.(As Ime6 c:faml46&ea r3pesr, 9. 157,.4 1712. 5

I'ni;til 1,. 1 ,)&9 (0,. '9(09 (0. 987 1,. '9:81

Fiiinal 0,926, (0,. 89 R02924 i1 7 177(6

Thiroait diaeter, cmn

riatial 2,. 533 2. '992!0 2.'91(6 /4. 44'5

Final 2. (63!5 2,. '1998 2,. '954 4,. 4'65

JPL TeciianicaL Memo randum 3 3-!5 ;7'9

'T~abllfie A. SStatirc-tfir-imng t (esit c 8oafdtf(ltois a-nad.s:rris~ulfs

:5

Fig. 1. Configuration of all-carbon nozzle for feasibility testing

6 JPL Technical Memorandum 33-579

-FINAL CONFIGURATIONOF NOZZLE TRANSITION(NOT TO SCALE)

Fig. 2. Layup step in nozzle transition and body fabrication

JPL Technical Memorandum 33-579

-FLAT-PLATELAMINATEDBLOCK

FINAL

7

Fig. 3. C/C nozzle during firing

8 J P L Technical Memorandum 33-579

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JIPL Techlmsal t emraml.ho 33-57 9)

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(a) CONVENTIONAL DESIGN

(b) INTERSTAGE STRUCTURE ELIMINATED BY USING NOZZLES

B«3» • M H H | n n m M n B H |

(c) CHAMBER RESHAPED TO SHORTEN AND UTILIZE VOLUME

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(d) NOZZLE AND CHAMBER CONSOLIDATED TO ELIMINATE REDUNDANT STRUCTURE

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F i g . 7 . C / C sample- ((S-gjr s i l v e r ©. 1 6 - c m ((l/lfe-im. )) stamdkafif,, law-af te r s h a p e d ! - c k a r g e ctittiiag; ((a)) s a m p l e 1„ ((b)) s a m p l e 2.

12 J PL, TechiEtical Memararaidhnnnni 33-57/9

Fig. 8. C/C sample (15-gr lead sheath 0. 32-cm ( l / 8 in. ) standoff high-density) after shaped-charge cutting

Fig. 9. C/C sample (15-gr lead sheath, 0 standoff high-density) after shaped-charge cutting

J P L Technical Memorandum 33-579 13

Fig. 10. Typical entry shell configuration

HEAT SHIELD0.25 TO 1 cm(0.1 TO 0.5 in.)THICK

60 deg

FILL TO 2 .5 cm (1.0 in. )OR NEAREST COMPLETE CELL

(18 in.) diam

I cm (24 in.) d

Fig. 11. Schematic of small doublyaero shells

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10 deg

1.9 cm (0.75 in. )

Jam

r curved developmental

JPL Technical Memorandum 33-579NASA - JPL - Coml.. L.A., Calif.