Search Results (3)

This study presents a simulation-based damage modeling and fatigue risk assessment of a reusable ceramic matrix composite thruster designed for short-duration, green bipropellant propulsion systems. The thruster is constructed from a fiber-reinforced ultra-high temperature ceramic matrix composite composed of zirconium diboride, silicon carbide, and carbon fibers. Time-resolved thermal and structural simulations are conducted on a validated thruster geometry to characterize the severity of early-stage thermal shock, stress buildup, and potential degradation pathways. Unlike traditional fatigue studies that rely on empirical fatigue constants or Paris-law-based crack-growth models, this work introduces a simulation-derived stress-margin envelope methodology that incorporates ±20% variability in temperature-dependent material strength, offering a physically grounded yet conservative risk estimate. From this, a normalized risk index is derived to evaluate the likelihood of damage initiation in critical regions over the 0–10 s firing window. The results indicate that the convergent throat region experiences a peak thermal gradient rate of approximately 380 K/s, with the normalized thermal shock index exceeding 43. Stress margins in this region collapse by 2.3 s, while margin loss in the flange curvature appears near 8 s. These findings are mapped into green, yellow, and red risk bands to classify operational safety zones. All the results assume no active cooling, representing conservative operating limits. If regenerative or ablative cooling is implemented, these margins would improve significantly. The framework established here enables a transparent, reproducible methodology for evaluating lifetime safety in ceramic propulsion nozzles and serves as a foundational tool for fatigue-resilient component design in green space engines. Full article

The oxidation and recession of carbon-fiber-reinforced ultrahigh-temperature ceramic matrix composites (C/UHTCMCs) fabricated via reactive melt infiltration (RMI) using Zr-Ti alloys with three different compositions are evaluated via an arc-jet tunnel test at temperatures above 2000 °C for 60 s. Thermodynamic evaluations show that the recession of the UHTCMCs is prevented by the formation of a solid solution of ZrTiO₄ on their exposed surface. Because an increase in the Zr content increases the melting temperature of ZrTiO₄, the recession of the composites increases as the Zr content in the infiltrated alloys decreases. UHTCMCs fabricated with Zr-20at%Ti showed the least recession (<5%). Full article

Ultra high-temperature ceramic matrix composites (UHTCMCs) based on carbon fibre (Cf) have been shown to offer excellent temperature stability exceeding 2000 °C in highly corrosive environments, which are prime requirements for various aerospace applications. In C³Harme, a recent European Union-funded Horizon 2020 project, an experimental campaign has been carried out to assess and screen a range of UHTCMC materials for near-zero ablation rocket nozzle and thermal protection systems. Samples with ZrB₂-impregnated pyrolytic carbon matrices and 2.5D woven continuous carbon fibre preforms, produced by slurry impregnation and radio frequency aided chemical vapour infiltration (RF-CVI), were tested using the vertical free jet facility at DLR, Cologne using solid propellants. When compared to standard CVI, RFCVI accelerates pyrolytic carbon densification, resulting in a much shorter manufacturing time. The samples survived the initial thermal shock and subsequent surface temperatures of >2000 °C with a minimal ablation rate. Post-test characterisation revealed a correlation between surface temperature and an accelerated catalytic activity, which lead to an understanding of the crucial role of preserving the bulk of the sample. Full article