Technology Status of Fuel Cooled Ceramic Matrix Composites for Dual-Mode Ramjet (DMR) and Liquid Rocket Engine (LRE) Applications

Beyer, Steffen; Schmidt-Wimmer, Stephan; Quering, Katharina; Wilhelmi, Christian; Steinhilber, M.

doi:10.2514/6.2012-5877

Cited by 11 publications

(6 citation statements)

References 6 publications

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“…As little flow stagnation occurs in flight, flight simulation species mass fractions were set based on a 3. premature combustion of intake injected fuel [33]. Downstream of the cowl closure, wall temperature is ramped in proportion to the inlet area contraction to reach 1800 K by the combustor entrance, holding these combustor and nozzle walls at the upper operational limit of ceramic matrix composites such as carbon fibre reinforced silicon carbide (C/C-SiC) [34].…”

Section: Tunnel Vs Flight -Flow and Model Conditionsmentioning

confidence: 99%

Performance of high mach number scramjets - Tunnel vs flight

et al. 2018

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While typically analysed through ground-based impulse facilities, scramjets experience significant heating loads in flight, raising engine wall temperatures and the fuel used to cool them beyond standard laboratory conditions. Hence, the present work numerically compares an access-to-space scramjet's performance at both these conditions. The Mach 12 Rectangular-to-Elliptical Shape-Transitioning scramjet flow path is examined via three-dimensional and chemically reacting Reynolds-averaged Navier-Stokes solutions. Flight operation is modelled through 800 K and 1800 K inlet and combustor walls respectively, while fuel is injected at both inlet-and combustor-based stations at 1000 K stagnation temperature. Room temperature walls and fuel plena model shock tunnel conditions. Mixing and combustion performance indicates that while flight conditions promote rapid mixing, high combustor temperatures inhibit the completion of reaction pathways, with reactant dissociation reducing chemical heat release by 16%. However, the heated walls in flight ensured 28% less energy was absorbed by the walls. While inlet fuel injection promotes robust burning of combustor-injected fuel, premature ignition upon the inlet in flight suggests these injectors should be moved further downstream. Coupled with counteracting differences in heat release and loss to the walls, the optimal engine design for flight may differ considerably from that which gives the best performance in the tunnel.

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Section: Tunnel Vs Flight -Flow and Model Conditionsmentioning

confidence: 99%

Performance of high mach number scramjets - Tunnel vs flight

et al. 2018

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“…Carbon fiber‐reinforced silicon carbide matrix composites are one of the most promising material alternatives to the heavy superalloys for applications in rocket propulsion, such as rocket nozzle extensions. Due to their excellent thermal and mechanical properties at high temperatures and due to their low material density, they will increase the payload capacity and improve the performance of future upper‐stage and orbital rocket engines …”

Section: Introductionmentioning

confidence: 99%

“…Due to their excellent thermal and mechanical properties at high temperatures and due to their low material density, they will increase the payload capacity and improve the performance of future upper-stage and orbital rocket engines. [1][2][3][4][5] A variety of different ceramic matrix composite (CMC) processing routes have been studied intensively during the last 10 years, including chemical vapor infiltration (CVI), [6][7][8][9] liquid polymer infiltration and pyrolysis (PIP), 10,11 and reactive melt infiltration, [12][13][14][15][16] and combinations of these. Processing costs and durations are significantly high using CVI or PIP fabrication routes.…”

Section: Introductionmentioning

confidence: 99%

Design, Manufacture, and Characterization of a Carbon Fiber‐Reinforced Silicon Carbide Nozzle Extension

Breede

Hofmann

Jain

et al. 2015

Int J Applied Ceramic Tech

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Nozzle extensions made of ceramic matrix composites (CMCs) have shown the potential to replace heavy superalloy nozzles and improve the performance of future upper‐stage and orbital rocket engines. Gas permeability has been reported to be a critical issue during the manufacture for CMC nozzles. This work shows the manufacture of a dense radiation‐cooled C/C‐SiC nozzle demonstrator. A multi‐angle fiber architecture was applied using filament winding technique to reduce the incidence of delaminations during the manufacturing process under high temperatures. Additional efforts were made to improve the final gas tightness and reduce the amount of residual silicon by means of an adapted liquid silicon infiltration process. The manufacture, the material, and structural characterization as well as a finite element analysis of a performed internal pressure test are presented.

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“…Within the combustor in particular, static temperatures exceeding 2000 K combine with static pressures of approximately 0.5 atm and magnified localised heating at shock impingement locations to foster an extremely hostile working environment (Barth, 2014). This heat must be transported away from the combustor walls, with even high temperature ceramic matrix composites such as carbon fibre reinforced silicon carbide (C/C-SiC) only remaining operable to approximately 1800 K (Beyer et al, 2012). This management, however, is more routine than for leading edges.…”

Section: Heat Managementmentioning

confidence: 99%

“…With this noted, flight forebody and inlet wall temperatures were set to be isothermal at 800 K, which has been demonstrated to not induce premature combustion of intake injected fuel (Gardner et al, 2002). Downstream of the cowl closure, wall temperature is ramped in proportion to the inlet area contraction to reach 1800 K by the combustor entrance, holding these combustor and nozzle walls at the upper operational limit of ceramic matrix composites such as carbon fibre reinforced silicon carbide (C/C-SiC) (Beyer et al, 2012).…”

Section: Tunnel Vs Flight -Flow and Model Conditionsmentioning

confidence: 99%

Improving performance of high Mach number scramjets: fuelling strategies and combustor design

Landsberg¹

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With prospective flight speeds exceeding Mach 12, airbreathing scramjets, incorporated in hybrid satellite launch vehicles, offer efficiency and operability benefits when compared to rocket-based systems. By capturing atmospheric air, the scramjet second stage negates the requirement for onboard oxidisers. Addressing the small satellite niche market, these hybrid systems provide a cost-effective solution to launching such payloads to orbit. At hypersonic speeds, however, conventional wing-mounted engine nacelles are not viable; the scramjet must be integrated to the airframe. Hence, the flight-candidate, Mach 12 Rectangular-to-Elliptical Shape-Transitioning (M12REST) scramjet blends a rectangular capture area (permitting parallel mounting of engine modules to the planar vehicle underside), with a structurally efficient elliptical combustor. As access-to-space scramjets suffer immense heating loads, and engine airflow residence times approach air-fuel reaction timescales, hydrogen fuel is utilised for its exceptional cooling capacity and rapid ignition characteristics. However, efficient fuelling techniques are required to avoid crippling the engine's performance through conservative combustor lengths. This thesis numerically and experimentally investigated whether the performance of such an airframe-integrated scramjet can be sufficiently enhanced through customisation of fuel injection and combustor geometry to achieve net thrust at Mach 12. Three-dimensional, chemically reacting Reynolds-averaged Navier-Stokes solutions enabled numerical analysis, while experimental validation was performed within The University of Queensland's, T4 Stalker Tube.Preliminary work investigated cascaded fuel injectors, targeting mixing and penetration improvements. Two streamwise-aligned jets were employed, with the upstream injector half the diameter of the rear, while the distance between each was varied. The upstream jet induced a low dynamic pressure region in its wake, shielding the downstream jet from the hypersonic crossflow. The leeward jet benefits from an increased local jet-to-freestream momentum ratio, while its larger diameter increases absolute penetration. Examining performance at M12REST mean combustor entrance conditions, unique optimal injector spacings existed for each performance metric, displaying improved penetration, spread, and mixing across the range of flight Mach numbers examined (6 ≤ M ≤ 12). However, jet-to-jet spacings of 4-6 total jet diameters displayed universal performance enhancements over single fuel jets. With penetration improvements of 30-40%, and mixing improvements of 40-70%, the simple fixed geometry, passive technique is ideal for use within an accelerating access-to-space scramjet.While cascaded injectors improved performance within uniform flows, the flows ingested by airframe-integrated scramjets are far less homogeneous. Non-uniform compression fields combine with thick boundary layers developed over the vehicle forebody to deliver density stratified flow to the combustor. This thesis ...

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Technology Status of Fuel Cooled Ceramic Matrix Composites for Dual-Mode Ramjet (DMR) and Liquid Rocket Engine (LRE) Applications

Cited by 11 publications

References 6 publications

Performance of high mach number scramjets - Tunnel vs flight

Performance of high mach number scramjets - Tunnel vs flight

Design, Manufacture, and Characterization of a Carbon Fiber‐Reinforced Silicon Carbide Nozzle Extension

Improving performance of high Mach number scramjets: fuelling strategies and combustor design

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