carbon c / sic

Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

Advanced Ceramic Matrix Composites (CMC's) for Space Propulsion Systems

U. Papenburg**, S. Beyer*, H. Laube*, S. Walter**, G. Langel*, M. Selzer.**

*Daimler-Benz Aerospace AG, Space InfrastructureLiquid Rocket Propulsion Development

P.O. Box 80 11 68D-81663 Munich, Germany

**Industrieanlagen-BetriebsgesellschaftmbH(IABG)High Temperature Technology

and Advanced MaterialsD-85521 Ottobrunn, Germany

ABSTRACT

For future rocket engine components advanced materialswith excellent thermo-mechanical properties and chemicalresistance are required to enhance the rocket engineperformance and reduce weight and cost. A technology forthe development and fabrication of high quality/highperformance rocket engine components out of continuousand/or short carbon fibre reinforced silicon carbide(C/SiC) is presented. Application of the C/SiC propulsioncomponent technology is demonstrated with respect todesign, structural and thermo-mechanical performance.The advantages of ceramic matrix composites (e.g. C/C,C/SiC) are: low density, high temperature and thermalshock resistance, high damage tolerance, low thermalexpansion, high and tunable stiffness and strenght as wellas their good thermal and electrical conductivity. An opti-mized application design/upscaling capability can beachieved by the flexibility of possible modifications inmaterial manufacturing. Within a joined development ofDasa and IABG, new types of advanced C/SiC compositeshave been investigated. Low density C/C-structures withcontinuous and/or short carbon fibre reinforcement wereinfiltrated with pyrolytic carbon by CVI-process and withliquid silicon, which is partly reacted to SiC. After a grin-ding process, oxidation and abrasion protection coatings(e.g. SiC, SiC>2) on the component surface can be achievedby PVD- and CVD-techniques. In addition, specificcharacterization aspects for qualification of advancedpropulsion components will be proposed and the furtherrequired technology improvements will be outlined. Thispaper gives a short overview of Dasa's and lABG's activi-ties on ceramic matrix composites with continuous andshort fibre reinforcements for propulsion components andfurther applications [I].

1. INTRODUCTION

For future rocket engine components with high heat loads,advanced materials are requested to master the extremeloads encounterd at high pressure and heat flux operation.Due to high pressures and heat fluxes the limits of conven-tional, i.e. metallic, materials are reached both for cryo-genic and storable propellants.

Copyright© 1997 by DasaPublished by the American Institute of Aeronauticsand Astronautics, Inc. with Permission

Qualified state of the art materials used today for Dasa'sliquid propulsion systems are copper and copper alloys,platin, nickel and nickel-based alloys, stainless steel,cobalt-based alloys and electrodeposited Nickel. However,to reduce the critical factors weight, manufacturing timeand costs, new materials and processes are needed toenhance enorme performance and reduce engine costs.

Ceramic matrix composites are a prime candidate for hightemperature components in space propulsion systems, offe-ring potential application at temperatures up to 1700°C.These materials are already in use as in-space minorsupport structures, brake disks for airplanes and cars, andsolid propellant engine nozzles. Manufacturing and experi-mental experience exist for other propulsion and spacerelated applications, such as e.g. thermal protectionsystems, and thrust deflectors.

In a first step, CMC could be introduced in hot propulsionsystem components with temperatures below about 1600°Cand with relatively small structural loadings. Examples arefilm and radiatively cooled combustion chambers andnozzles for small bipropellant engines, and nozzle exten-sion skirts for large liquid propellant engines.

In the mid 1980s, Dasa (formerly MBB) demonstration-tested C/C nozzle extensions with a cryogenic subscalecombustion chamber, operating at a pressure of 100 bar anda mixture ratio of 5 - 6, yielding a combustion gas tempera-ture of about 3500 K. Figure 1 shows the C/C nozzle testhardware, Figure 2 hot test with this C/C nozzle (withoutoxidation protection coating). The nozzle was mounted tothe combustion chamber by means of a compression joint.Total test duration was 100 seconds, with the C/C testarticle showing no structural damage, demonstrating thefeasibility of a radiation cooled nozzle extension on a highpressure cryogenic engine [1/2/3].

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

Figure 1: C/C-nozzles (test hardware)

Figure 2: Cryogenic subscale test on test facility P59.1at Dasa (MBB) in Ottobrunn, Germany

Since then, new CMC materials and manufacturingprocesses have been developed by various institutionsthroughout Europe [1], and SEP has meanwhile demon-strated the feasibility of their material in hot fire tests witha full-scale radiation cooled HM7-engine nozzle extension,and has developed the radiation cooled C/C-extensiblenozzle for Pratt & Whitney's RL 10B-2 engine [2/3/4].

This paper describes the development and fabrication ofnew high quality/high performance rocket engine compo-nents like 400 N combustion chambers for small bipropel-lant engines and nozzle extensions for present and futureengines made of C/SiC composites. Chapter 2 focuses onthe C/SiC materials and capabilities at Dasa/IABG, andchapters 3 and 4 discusses present and future applications.

2. MATERIAL SELECTION ANDMANUFACTURING PROCESSES

The typical C/SiC manufacturing process used at IABGfor rocket propulsion components is shown in Figure 3.Both long and short fibre materials are included.

Manufacturing Process ||i

Continuous woven C/C-Pregregs orRandom orientated chopped C/C-felt

iCFRP-Moulding or Winding

Techniques

Carbonization in Vacuum, 1000°C

Graphitization in Vacuum, 2100°C

iMachining and Joining of Medium

Size Components

iChemical Vapour Infiltration (CVI) 1

with Pyrolitic Carbon |

1Liquid Silicon Infiltration (LI) and ISiC-Reaktion in Vacuum, 1 800°C j

Pre-Grinding

PVD/CVD Oxidation/ErosionProtection Coating (SiC, SiC>2\

C/SiC Rocket Propulsion 1Components |

11

1Figure 3: Manufacturing of C/SiC-rocket propulsion

components

2.1 RAW MATERIALS, MACHINING ANDJOINING TECHNOLOGY

Depending on the requirements from different propulsionapplications, different types of carbon fibre reinforcement(continuous and short fibre reinforcement) have to beconsidered. Propulsion units with continuous fibre reinfor-cement (2D-C/SiC) can be realized with woven C/C-prep-regs or C-rovings in different winding techniques up todiameters of 2000 mm (see Figure 4).

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

The raw material used for the short carbon fibre reinforce-ment (isotropic C/SiC) is a porous C/C rigid felt, made ofshort and isotropically (random) oriented carbon fibres,which are moulded with phenolic resins or pitch at highpressures to a kind of CFRP blank, which is available invarious sizes. Dependent on the moulding process, theshort carbon fibres are randomly orientated in the blankbody and, hence, an isotropic mechanical behaviour isachieved [5/6/7].

Figure 4: Winding of continuous carbon fibre reinforcedCFRP-tubes up to diameters of 2000 mm

During pyrolisation/carbonization heat treatment up to1000°C, the phenolic matrix reacts to carbon matrix (C/C).The density of these C/C compounds is between 0.7 and1.5 g/cm . A graphitization process in inert atmospheres(by heat treatment at temperatures up to 2100°C) and a CVI(Chemical Vapour Infillration)-process with pyroliticcarbon reduces the chemical reactivity of the carbon fibreswith liquid silicon. This process has a decisive influence onthe physical and mechanical properties of the C/SiCcomposites. These C/C-raw materials are deliverable inblanks of diameters up to 2500 mm. However, the shortfibre reinforced C/C felt is sufficient rigid to mil l it tovirtually any shape, as shown in Figure 5 with a rathersophisticated support rear structure all cut out of a single"base" by standard NC milling. The ribs shown here are notthicker than 1.25 mm with ± 0.1 mm tolerance. This is oneof the most significant advantages of this material since itdrastically reduces the forming costs [8/9/10].

Figure 5: Greenbody machining of short fibre reinforcedC/C-felt

The typical CVD coating process needed to protect theC/SiC from the oxidizing atmosphere in a combustiondevice is shown in Figure 6.

Figure 6: 2D-C/SJC nozzle extension in the plasmasupported CVD-coating process

Medium size C/C components (0 < 1500 mm) can also bejoined at their mechanical interfaces with C/C bolts andscrews (before Si-infiltration). Also, the component mustremain free of oil or similar contamination to ensure ahomogeniuos Si-infil tartion in the next process [11/12].

2.2 CERAMIC INFILTRATION PROCESSING

After careful microscopic inspection of the pre-shapedC/C-body the uni t wi l l be mounted in a dedicated hightemperature furnace under vacuum with silicon supply inthe l iquid phase, i.e. at > 1500°C. The lower end of thestructure are dipped into the l iquid silicon. Because of thecapilary forces inherent in the porous C/C component, themolten silicon is sucked upwards into the structure. Subse-

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

quently the temperature is increased to 1800°C and thesilicon reacts with the carbon matrix and the surface of thecarbon fibres to form silicon carbide matrix (conversionprocess). Dependent on the amount of carbon and infil-trated silicon the resulting ceramic matrix compositeconsist of carbon fibres plus two matrix constituents (SiCand Si) in various concentrations. As mentioned before, theamount of carbon and silicon has to be apportioned exactlyto prevent a chemical reaction with the silicon and the rein-forcing carbon fibres. However, the fraction of metallicsilicon contained in the blank after conversion amounts to 5- 25% by weight. The density of the infiltrated C/SiCcomposite is typically between 2.1 g/cnr (2D-C/SJC) and2.7 g/cnr (isotropic C/SiC). The use of vacuum conditionsis not mandatory for all applications, yet was found manda-tory for pore-free surfaces. The process wil l work atambient pressure also with Argon. The duration of thethermal infi l trat ion process is a direct measure of the extentto which the carbon matrix and carbon fibres react with Sito SiC, i.e. how much of the carbon fibre reinforcementwil l remain in the item. This parameter together with thetemperature control to some extent the ratio of stiffness andstrenght versus ductility of the C/SiC. After controlledcool-down of the item, it wi l l be carefully examined againmicroscopically and by NOT methods (X-ray radiography)with respect to density variations, pores, microcracks etc.[11/12/13]. The main material properties of short (Iso) andcontinuous (2D) fibre C/SiC are summarized below insection 2.4.

2.3 FACILITIES FOR MANUFACTURING OFC/SIC COMPONENTS

All required facilities to perform the development andmanufacturing work are available at IABG and Dasa inOttobrunn. The various development and manufacturingsteps are accompanied by analytical work at the IABGlaboratory, like SEM investigations of materials, mecha-nical tests including fracture toughness analysis and inve-stigations of polished sections to estimate the receivedquality. The required equipment is available, i nc lud ingSEM's, CTE-measurcment system, room and high tempe-rature mechanical test devices and NdE (Non-destructiveEvaluation) devices etc.. The manufacturing of C/SiCpropulsion structures up to diameters of 1500 mm are stateof the art. The fabrication of such propulsion structures canproceed immediatly without upgrading existing facilities.The facilities for propulsion un i t manufacturing/processingup to > 2500 mm are also available (see Figure 7).

Figure 7: High temperature vacuum furnace for C/SiCmanufacturing (0 4 m)

Currently the C/C greenbody joining technology has beendemonstrated on C/SiC burner components with a lenght of3000 mm. No ageing and creeping effects (since 1 year atan application temperature of 1400°C) were determined.Rapid manufacturing and breadboarding is achieved bydirect data l ink of the CATIA design stations for the NC-programming for net-shape C/C-milling. Of all of the highperformance, commercially feasible and advanced mate-rials, only C/SiC offers the freedom to be moulded intointricate "sculpture-like" shapes. It can be easily applied tovery small and very large propulsion structures. The mainfeatures and material properties of C/SiC manufactured bythe above described processes are listed in the next section.

2.4 C/SIC MATERIAL PROPERTIES AND MAINFEATURES

The C/SiC features and advantages can be summarized asfollows:

• Low specific density (2.10-2.70 g/cm )• Low CTE (3.5-6.5 lO^K'1)• Good and tuneable thermal conductivity (5-135 W/mK)• Chemical and erosion resistant• High temperature resitance (> 1700°C)• No detected ageing and creep deformation under stress• No open porosity• Fast and low-cost near net shaping• Short time manufacturing processing• High f lexibi l i ty in structural design, ultra-light

weight and up-scaling capability• High and to some extent tunable stiffness

(90-250 GPa) and strenght (140-350 MPa)

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

Note that for some parameters only the typical values arelisted, e.g. the stiffness and strenght. The C/SiC structure inthis state has to be with suitable diamond tools to achievethe required surface quality [11/12/13].

MaterialProperties

Max. OperatingTemp. (°C)

Density(g/cm-'')Young's

Modulus (GPa)Tensile Strenghl

(MPa)Thermal Expan-

sion (10-6/K)Thermal Condu-

t ivi ty (W/rnK)

Haynes 25

I K K )

9,1

225

980

13

10

Al

300

2,7

70

250

25

170

AISI347

900

7,9

200

620

18

13

Iso-C/SiC(short fibre)

1600

2,7

260

140

3,5

135

2D-C/S1C(cont. fibre)

1600

2.1 - 2.3

90 - 140

250 - 300

4.5 ± - 0.5 ||

10 1 - 30 ||

Table 1: Properties of candidate propulsion materials incomparison

3. C/SIC COMPOSITES FOR SPACEPROPULSION SYSTEMS

3.1 TACTICAL MISSILE PROPULSION

A typical field of application for CMC has been the area ofsolid-propulsion nozzles.

IABG developed and produced CMC-components (nozzles,thruslers, etc.) for solid-propellant propulsion systems fortactical missiles funded by the German Ministry ofDefence. C/SiC composite-based HT-structures offernumerous advantages, obviously led by high temperatureresistance significant weight savings. Figure 8 shows C/SiCthrust nozzles for solid rocket propulsion.

In experiments the thermoshock, temperature and mecha-nical resistance of different nozzles (average chamber pres-sure to 80 bar) was demonstrated. The original matallicnozzles were exchanced by CMC-nozzles without any redi-sign in the nozzle mounting. Because nozzles or nozzlethroat inserts of CMC have the advanlge of a low abrasion,the loss of thrust by nozzle throat expanding during opera-tion is less severe with C/SiC nozzles.

Figure 8: Thrust nozzles for solid rocket propulsion

3.2 COMBUSTION CHAMBERS FOR BI-PROPELLANT SATELLITE THRUSTERS

The small bipropellant engines in operation tools usemetallic combustion chambers, which are more or lessrefractory according to the type of cooling used. An engineusing regenerative cooling may very well be made of stain-less steel, but film or radiation cooled engines requirehigher performance materials. Beryllium may be used up tomaximum temperlures up to 1100°C. Another candidate,niobium, is limited by its susceptibility to oxidation, resul-ting in a demand for additional surface coating rather thanby its melting temperature (2400°C). Typical operatingconditions are chamber pressures of 10 bar and combustiontemperatures around 3000 K. The propellants are MMHand NTO. The chamber walls arc typically film- and radia-tion cooled; some systems requires in addition regenerati-vely cooled throut sections. Increased wall temperatures byreduced active cooling result in higher performanceengines.

While preparing a new generation of small high-perfor-mance bi-propcllanl engines, Dasa and IABG have investi-gated CMC's. The main advantage of these CMC compo-sites is the increased maximum operating temperature up to1700°C with improved resistance to thermal cycles . Figure9 shows a ION C/SiC bipropellant engine combustionchamber/nozzle, made with short carbon fibre reinforcedsilicon carbide. First laboratory hot tests with 10 Ncombustion chambers show no signif icant thermal ageingor erosion effects after 400 thermal cycles (cumulativeoperating time: 50 h, corresponding time 6 h, see Fig. 10)[6/7/8].

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

Figure 9: 10 N C/SiC satellite nozzle with short fibrereinforcement

Figure 1 1 : 400 N Iso-C/SiC (short fibre) combustionChamber

Figure 10: I O N combustion chamber during hot firing(experimental)

Figure 12: 400 N 2D-C/SJC (continuous fibre)combustion chamber

In a further step, Dasa and IABG will test in the near futuredifferent 400 N bipropellant engines on the satellitepropulsion test facility in Lampoldshausen, Germany.These 400 N combustion chambers will be made of diffe-rent C/SiC like short and continues fibres with differentprotection systems.

The 400 N bipropellant test engines made of C/SiC atIABG are shown in Figures 1 1 and 12.

It is also foreseen to test chambers with continuous fibres,made at Dasa-Dornier/Friedrichshafen.

3.3 LARGE LIQUID PROPELLANT ENGINES

Typical applications of CMCs in hot components of largeliquid propellant engines arc today envisionized for hot gasflow components and nozzle extensions, the main advan-tages being life and/or performance increase and possiblyweight and cost savings in scene applications.

The requirement for materials for liquid propellant enginesare extremely high. Materials for an engine with storablepropellants, are exposed to combustion temperatures about3000 K at combustion pressures up to 50 bar (for thecombustion chamber) and have to be chemically anderosion resistant. The nozzle extension of such a thrustchamber is typically exposed to temperatures of about 1300K (metallic version).

To use CMCs like C/SiC for high pressure cryogenicengines the requirements concerning heat fluxes, thermal-schock resistance, engine loads and thermo-mcheanicalbehaviour are even higher in comparison to storableengines (chamber pressures up to 200 bar, combustion gastemperatures around 3500 K).

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

As described in the introduction, the feasibili ty of a CMCnozzle extension has already been demonstrated, and a nextpotential application of C/SiC is seen as film/radiationcooled nozzle extension in storable propellant upper stageengines (low structural loads) [9/10].

One application of C/SiC in a flight engine is the thermalprotection coating of the Vulcain engine igniter How path,see Figure 13. The figure shows cut of the igniter el vowafter 16 ignitions. No delaminations or significant erosionof the C/SiC coating were detected.

Figure 14: Igniter elbow with C/SiC thermal protectionof the Vulcain engine after engine testing

A hot fire feasibility demonstration test of this fl ight quali-fied short fibre reinforced C/SiC coaling of the igniterelbow are shown in Figure 15 (hot fire feasibility demon-stration test) [1 1/12/13]. This C/SiC hot gas pipe withstandin 30 tests gas temperatures up to 2000°C combined withabrasive particles and thermalshocks of 2100K/sec withoutany ageing, abrasion effects or damages.

Figure 15: Hot fire feasibil i ty demonstration lest of shortfibre reinforced C/SiC thermal prolcction forIhe igniler elbow of Ihc Vulcain Engine

3.4 RAMJET ENGINE INLET RAMP FLAP

Polenlial CMC-componenls in ramjels are hoi parls such asshock diffusor, subsonic diffusor, inlel cone combustionchamber insulation, flame holder and thrust nozzle. Thematerial requirements of ihesc componcnls are Iher-malshock , temperature, mechanical and erosion resistance.The level of pressure and slresses depend on ihe operalingallitude of cruise missiles. Gas lemperatures with morethan 2000°C are possible in ramjels depending on Ihe lypeof l iquid and solid propellanl and ihe combuslion chamberpressure.

A C/SiC inlel ramp flap of a hypersonic Sa'nger Iransporla-lion system was designed, bui l l , and Icsled in Ihe frame ofihe German ramjet propulsion technology Sa'nger program.Figure 16 shows ihis inlel ramp Hap [14/15].

Figure 16: C/SiC inlel ramp (lap demonslralor of Ihehypersonic Sa'nger space iransporlation system

4. OTHER C/SIC COMPOSITE APPLICATIONS

Before presenting ihe use of C/SiC for rockel propulsionapplicalion in the next chapter, two oilier areospacc exam-ples of successful! applicalion of C/SiC components wil l bedescribed.

4.1 C/SiC COMPOSITES FOR THRUSTDEFLECTORS

Highly promising results have already been obtained in ihefield of ihrust vector control systems for jels and laclicalmissile propulsion, where tests of new concepts have beenmade possible by the use of C/SiC composites (thruslers,movable nozzles, valving of hoi gases, jcl blasts, etc.).

Figure 16 shows the 2D-C/SJC Ihrust deflector for the X-31experimental aircraft which successfully demonstrated theuse of this material under aclual f l igh t condit ion.

The successful experimental verif ication of thcrmalshockand Icmpcralurc resistance of C/SiC Ihrustcrs give legili-malc hopes for short-lime applications of CMC lo olhcr

Copyright © 1997, American Institute of Aeronautics and Astronautics, Inc.

components of high velocity missiles. All exposed surfacecomponents of high velocity missiles are set out hightemperatures and mechanical stresses by effects of aerody-namic heating and friction [16/17/18].

Figure 16: 2D-C/SiC thrust deflector for the X-31aircraft

4.2 ULTRA-LIGHTWEIGHT C/SIC MIRRORSC/SiC with short fibre reinforcement has high applicationpotential for optomechanical, especially ultra-leightweightapplications (e.g. mirrors, antennas, optical benches,

^telescope structures) with a mass of < 18 kg/m . C/SiC formirror and telescope structure applications actually resultedfrom an extended trade of available materials in context ofthe FIRST reflector (diameter: 3500 mm) and of the MSG(Meteosat Second Generation Satellite) imager scan mirrorwhich revealed that none of the "classical" materials wouldfulfi l l all given requirements, such as mass versus size,stiffness, CTE radiation resistance, thermal conductivityetc.. These mirrors will operate under extreme conditions:in geostationary orbit, exposed to hot and cold space, radia-tion, and worst: rotating with the satellite. One of the mostadvantageous features for experienced space-borne opto-mechanical instrument designers is the combination of highstiffness, low CTE and good thermal and electrical conduc-tivity, particulary in contrast to Zerodur, Aluminium andBeryllium [19/20/21/22/23/24]. Figure 17 shows the ultra-lightweight C/SiC-Scan Mirror for MSG.

Figure 17: Ultra-lightweight C/SiC-Scan Mirror forMSG

5. CONCLUSIONS AND OUTLOOK

Different space propulsion and related components made ofdifferent CMC materials are being introduced into flightsystems. Examples are solid propulsion nozzles, thrustdeflectors, hot gas flow path components, thermal protec-tion systems, and in-space high precision mirror supports.Nozzle extensions are on the verge of being used in largeliquid propulsion upper stage engines.This paper described advanced C/SiC materials developedat IABG and Dasa in Ottobrunn, Germany. These materialsand processes are available for future combustionchamber/nozzle assemblies for small bipropellant thrusters(Dasa 10 N and 400 N class) and for storable upper stagenozzle extensions (film/radiation cooled). Other potentialfuture applications include the use of C/SiC as combustionchamber material for low pressure storable propellantengines with regenerative cooling.

Copyright© 1997, American Institute of Aeronautics and Astronautics, Inc.

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[19] U. Papenburg, B. Kunkel, V. Billig: "C/SiC - aPromising Material for Demanding Applications inOptics, Space and Other Severe Environments", UN-Conference, New York, July 1994.

[20] lABG/Dasa Sheet on "C/SiC - a New Material forOptical and Optomechanical Applications", Juli1996.

[21] G. Appel, K.K.O. Ban'Thermo-mechanical Testingof HERMES Winglet Box", Proc. AIAA/DGLR, 5thInt. Aerospace Planes and Hypersonic TechnologiesConference, IABG, Munich, 1993.