Hermann Hald

Operational limits for reusable space transportation systems due to physical boundaries of C/SiC materials

Abstract The paper concentrates on the surface interaction phenomena of both SiC-coated and uncoated C/C–SiC material of DLRs well known liquid siliconizing process (LSI) while exposing samples and structures to simulated and real space vehicle reentry conditions. During reentry, physical limits for these types of CMC materials are defined by oxidation resistance, passive–active transition conditions and erosion effects by various reasons like ‘mechanical’ fluid–surface interaction. Special emphasis is given to the passive–active transition behaviour of SiC and the related ‘temperature jump effect’ which will be explained by new theoretical research results. In addition a quantitative evaluation of the effect is given based on experimental test sample data. The considerations given in the paper are checked against literature data, observations and experience which have been achieved so far by manifold tests with samples in various facilities, experimental space reentry tests with samples on Russian FOTON capsules and a test with a hot structure during the EXPRESS mission (experiment CETEX). Finally, the understanding of the mechanisms may lead to a revised strategy for optimizing criterions for reentry trajectories of reusable space transportation systems.

Application of Transpiration Cooling for Hot Structures

Specific parts of re-entry vehicles are exposed to severe conditions. Thereby, the material’s capabilities can be exceeded by far and advanced cooling methods become necessary. Within the scope of this work, transpiration cooling was investigated in arc jet heated plasma flows by means of flat plate models. Screening tests pointed out, that transpiration cooling at the conditions tested is working well. Extensive testing at more severe conditions was done using three porous sample materials: Standard C/C with coolant flows parallel and perpendicular to the material’s fibre layers and highly porous C/C. Coolant gases used were air, argon, helium and nitrogen. Minimal optimal coolant mass flows of 0.5 g/s Ar, 0.2 g/s He and 0.4 g/s N2 were determined resulting in sample under surface temperature reductions of 50-60%. Altogether, sample under surface temperature reductions of 64% for He, 65% for Ar, 67% for air and 70% for N2 were detected. These test series verified that transpiration cooling can be applied successfully for hot structures at application relevant re-entry conditions.

Technological Aspects of Transpiration Cooled Composite Structures for Thrust Chamber Applications

The long-term development of transpiration cooled ceramic rocket thrust chambers at the German Aerospace Center (DLR) currently culminates in designs of self-sustaining fibre reinforced rocket engine chamber structures. This paper explains characteristic issues and potential benefits introduced by this new technology, which seem to be achievable in terms of weight and cost reduction, increased reliability and higher lifetime due to no longer existing thermal cycling sensitivity. The paper furthermore describes design and functional aspects of the chamber, the component manufacturing process and shows some experimental results of test campaigns with respect to structure and material relevance. DLR’s development road map is sketched and the technology readiness level achieved so far is discussed. Nomenclature ISP = specific impulse kd = coefficient of Darcyan permeability kf = coefficient of non-Darcyan (Forchheimer) permeability L = flow length pin = inflow pressure pout = outflow pressure T = temperature v = velocity Δp = pressure loss λ = thermal conductivity (CTC) e = porosity η = dynamic viscosity ρ = density

Transpiration-Cooled Ceramic Thrust Chamber Applicability for High-Thrust Rocket Engines

The development of ceramic rocket engine thrust chambers at the German Aerospace Center (DLR) currently concentrates on designs of self-sustaining, transpiration-cooled, fiber-reinforced ceramic rocket engine chamber structures. This paper discusses characteristic issues and potential benefits introduced by this technology. Achievable benefits are the reduction of weight and manufacturing cost, as well as an increased reliability and higher lifetime due to thermal cycle stability. This paper discusses the current status of DLRs ceramic thrust chamber technology and potential applications for high thrust engines. The test results of DLRs ceramic thrust chamber project KSK are used for a rough approximation of the performance of high thrust applications. Based on the KSK test results an extrapolation is performed. c*-efficiency and geometrical scaling effects are taken into consideration. Due to favorable scaling effects, high thrust applications will profit by all benefits of the discussed technology, while avoiding the most significant performance drawbacks.

Porous Versus Porthole Fuel Injection in a Radical Farming Scramjet: Numerical Analysis

Numerically computed engine performance of a nominally two-dimensional radical farming scramjet with porous (permeable C/C ceramic) and porthole fuel injection is presented. Inflow conditions with Mach number, stagnation pressure, and enthalpy of 6.44, 40.2MPa, and 4.31 MJ/kg respectively, and fuel/air equivalence ratio of 0.44 were maintained, along with engine geometry. Hydrogen fuel was injected at an axial location of 92.33mm downstream of the leading edge for each investigated injection method. Results from this study show that porous fuel injection results in enhanced mixing and combustion compared to porthole fuel injection. This is particularly evident within the first half of the combustion chamber where porous fuel injection resulted in mixing and combustion efficiencies of 76% and 63% respectively. At the same location, porthole fuel injection resulted in efficiencies respectively of 58% and 46%. Key mechanisms contributing to the observed improved performance were the formation of an attached oblique fuel injection shock and associated stronger shock-expansion train ingested by the engine, enhanced spreading of the fuel in all directions and a more rapidly growing mixing layer.

Investigations on Fibre Reinforced Combustion Chamber Structures under Effusion Cooled LOX/LH2 Operation

The development of effusion cooled ceramic rocket thrust chambers at the German Aerospace Center (DLR) since more than one decade leads currently to first designs of selfsustaining fibre reinforced chamber structures. Within the German Research Network ‘Propulsion 2010’, a closed co-operation between the DLR and the German Space Propulsion Industry (EADS Astrium Space Transportation), the technology demonstration for a ceramic thrust chamber assembly until the end of 2010 is foreseen. The combination of a porous metal injector, an inner C/C combustion chamber liner, covered by load carrying CFRP, and a fuel cooled CMC nozzle extension is planned to be demonstrated at the European Research and Technology Test Bench P8 at DLR Lampoldshausen, in LOX/LH2 operation within the KSK project (KSK – Keramische Schub-Kammer). The final tests in 2010 will be conducted at 80 mm sub-scale level of inner chamber diameter. First a series of preliminary tests on 50 mm level have been performed in 2008, with the purpose of investigating both functional and structural principles.

H2-O2 porous fuel injection in a radical farming scramjet

This paper reports on the experimental testing of oxygen compatible ceramic matrix composite porous injectors in a nominally two-dimensional hydrogen fuelled and oxygen enriched radical farming scramjet in the T4 shock tunnel facility. All experiments were performed at a dynamic pressure of 146 kPa, an equivalent flight Mach number of 9.7, a stagnation pressure and enthalpy of 40 MPa and 4.3 MJ/kg respectively and at a fuelling condition that resulted in an average equivalence ratio of 0.472. Oxygen was pre-mixed with the fuel prior to injection to achieve enrichment percentages of approximately 13%, 15% and 17%. These levels ensured that the hydrogen-oxidiser mix injected into the engine always remained too fuel rich to sustain a flame without any additional mixing with the captured air. Addition of pre-mixed oxygen with the fuel was found to significantly alter the performance of the engine; enhancing both combustion and ignition and converting a previously observed limited combustion condition into one with sustained and noticeable combustion induced pressure rise. Increases in the enrichment percentage lead to further increases in combustion levels and acted to reduce ignition lengths within the engine. Suppressed combustion runs, where a nitrogen test gas was used, confirmed that the pressure rise observed in these experiments as attributed to the oxygen enrichment and not associated with the increased mass injected.

Structural Investigations on Cryogenically Operated and Transpiration Cooled Fiber Reinforced Rocket Thrust Chambers

This paper will give a subsumption of the structural implementation of an integrated fiber reinforced rocket thrust chamber design, which took in the meanwhile more than one decade of development effort. Of particular interest in this case is a hybrid design approach showing Ceramic Matrix Composits as high temperature inner liner materials and a covering light weight Carbon Fiber Plastic Housing, whereas these principally different material components are joined by metallic flanges at the front edges in an extensively decoupled design philosophy, excluding compulsive loads as far as possible. In 2010 the long and intensive experimental work led into a test campaign at the European Research and Technology Test Facility P8 at DLR-Lampoldshausen using for the first time the fully integrated chamber design under cryogenic high performance conditions, using LOX/LH2. In this campaign the structural concept could be proved completely under all structurally relevant parameters. With respect to the demonstration of the system efficiency under optimized operational flow parameters a further test campaign early in 2012 had been performed at the recently renewed local P6.1-test bench. The paper discusses predominantly structural design aspects.

Reentry Flight Testing of a C/C-SiC Structure with Yttrium Silicate Oxidation Protection

Hot structure design with ceramic matrix composites (CMC) which are based on carbon fiber reinforced materials like C/C-SiC is a key technology for the development of lightweight thermal protection systems (TPS) of advanced space reentry vehicles. In order to select suitable CMC and protective coating materials, ground based hypersonic flight simulation with plasma wind tunnel test facilities represents a valuable method for qualifying several structure components and to achieve a deeper understanding of the thermo-chemical interactions at the material’s surface. However, regarding the technical limitations with ground based test facilities it is indispensable to perform real flight testing with CMC structures in terms of validation of new design concepts which are based on innovative materials. This paper describes a reentry technology experiment – named “KERAMIK” – with its structural components designed and manufactured fully in C/CSiC material, a special type of CMC developed by DLR. It was successfully flown on the Russian Foton-M2 mission in June 2005. One of its C/C-SiC panels was partly protected with different coating. Special attention is paid to an effective two-layer coating system that is based on yttrium silicate and was chosen to improve both oxidation protection and erosion resistance of the C/C-SiC material. Before selected for the flight test, this coating system was extensively tested in plasma wind tunnel facilities. A comparison of the test results with such coated C/C-SiC specimens obtained from test campaigns in different plasma wind tunnel facilities shows two principal approaches for systematic ground based reentry qualification tests.

Empirical Verification of Effusion Cooled CMC Rocket Thrust Chambers

Thrust chambers are one of the most sensitive components in rocket propulsion systems due to safety and efficiency related to reasonable costs. Competitive space transportation systems ask for such low cost and high sophisticated solutions. A very promising approach in this field is the development of an effusion cooled CMC combustion chamber design, which offers a new perspective on the way to reliable future cryogenic rocket engines. The undoubted advantages of transpiration cooling under the critical view of efficiency, damage tolerance and low cost aspects can be accomplished with a relatively simple and low weight concept using carbon fiber composite materials. In the recent years DLR works on the application of effusion technology. The paper illustrates the empirical development steps accompanied by numerical simulations until a break through in form of the latest high performance tests early in 2005.