Search Results (6)

The conversion of legacy missile engines into space propulsion systems represents a strategic opportunity to accelerate Europe’s access to orbit while advancing sustainability and circular-economy goals. Rather than discarding decommissioned hardware, repurposing missile propulsion can reduce development timelines, retain valuable materials, and leverage proven architectures for new applications. This perspective outlines the potential of the Soviet-era Isayev S2.720 engine as a representative case, drawing on historical precedents of missile-to-launcher conversions worldwide. A three-pillar methodology is proposed to frame such efforts: (i) the adoption of cleaner propellants such as LOX–LCH₄ in place of toxic hypergolics; (ii) remanufacturing and upgrading of key subsystems through additive manufacturing, AI-assisted inspection, and digital twin modelling; and (iii) validation supported by dedicated testing, life-cycle assessment (LCA), and life-cycle costing (LCC). Beyond the technical aspects, the paper discusses retrofit applicability, cost considerations, and the role of standardization in enabling future certification. By positioning the S2.720 as a model, this study highlights the broader strategic value of adapting decommissioned propulsion systems for modern orbital use, providing insight into how Europe might integrate legacy assets into a more sustainable and resilient space transportation framework. Full article

This investigation presents the comprehensive thermo-structural analysis of the propellant tanks utilized in the CROME propulsion system, focused on examining the structural effects caused by the storage of liquid nitrogen, liquid oxygen, and liquid methane. These fluids operate at extremely low temperatures and generate large thermal stress gradients in the tanks, significantly influencing their structural properties. For this reason, it is of vital importance to inspect the generation of mechanical and thermal stresses in the tanks to assess the risk of structural failure. To accomplish this effort, this analysis evaluates the tanks containing liquid nitrogen, liquid oxygen, and liquid methane at pressures of 200.0 psi and 400.0 psi. A coupled finite element analysis was performed in Star-CCM+ to compute the resulting Von Mises stresses under steady-state conditions. These stress results were used to determine the factor of safety in each case, enabling a quantitative assessment of structural integrity in the tanks while operating with cryogenic fluids under different pressure loadings. Full article

This study examines the thermal management of the Centennial Restartable Oxygen Methane Engine (CROME) feed system under two propellant tank pressure conditions: 33 psi (227.5 kPa) and 100 psi (689.5 kPa), at a constant liquid methane flow rate of 0.9 lbm/s (0.4 kg/s). Using the Eulerian Single-Phase (ESP) model, the initial test validated experimental data, showing close agreement in total pressure (experimental: 31 psi; CFD: 33 psi) and temperature measurements (experimental: −287.3 °F and −300 °F; CFD: −299 °F and −294 °F) with deviations of 6.4% and ≤4.1%, respectively. For the second test, a simplified Volume of Fluid (VOF) model was used, adjusted for varying liquid-to-gas volume fractions. The best agreement with experimental data was found with 100% GN2, showing a 3.1 psi pressure rise and a 3.3% error. These findings show the importance of improving thermal management and precision control in cryogenic LOX-LCH4 feedline systems for optimal engine performance. Future research will focus on exploring pressures up to the propellant tank’s maximum rated limit of 400 psi. Full article

The need to refuel future missions to Mars and the Moon via in situ resource utilization (ISRU) requires the development of LOX/LCH4 engines, which are complex and expensive to develop and improve. This paper discusses how the use of digital engineering—specifically physics-based modeling (PBM)—can aid in developing, testing, and validating a LOX/LCH4 engine. The model, which focuses on propulsion performance and heat transfer through the engine walls, was created using Siemens’ STAR-CCM+ CFD tool. Key features of the model include Eulerian multiphase physics (EMP), complex chemistry (CC) using the eddy dissipation concept (EDC), and segregated solid energy (SSE) for heat transfer. A comparison between the complete GRI 3.0 and Lu’s reduced combustion mechanisms was performed, with Lu’s mechanism being chosen for its cost-effectiveness and similar output to the GRI mechanism. The model’s geometry represents 1/8th of the engine’s volume, with a symmetric rotational boundary. The performance of this engine was investigated using NASA’s chemical equilibrium analysis (CEA) and STAR-CCM+ simulations, focusing on thrust levels of 125 lbf and 500 lbf. Discrepancies between theoretical predictions and simulations ranged from 1.4% to 28.5%, largely due to differences in modeling assumptions. While NASA CEA has a zero-dimensional, steady-state approach based on idealized conditions, STAR-CCM+ accounts for real-world factors such as multiphase flow, turbulence, and heat loss. For the 125 lbf case, a 9.2% deviation in combustion chamber temperature and a 15.0% difference in thrust were noted, with simulations yielding 113.48 lbf compared to the CEA’s 133.52 lbf. In the 500 lbf case, thrust reached 488 lbf, showing a 2.4% deviation from the design target and an 8.6% increase over CEA predictions. Temperature and pressure deviations were also observed, with the highest engine wall temperature at the nozzle throat. Monte Carlo simulations revealed that substituting LNG for LCH4 affects combustion dynamics. The findings emphasize the need for advanced modeling approaches to enhance the prediction accuracy of rocket engine performance, aiding in the development of digital twins for the CROME. Full article

Due to the extreme temperatures inside the combustion chambers of liquid propellant rocket engines, the walls of the combustion chamber and the nozzle are cooled by either the fuel or the oxidizer in what is known as regenerative cooling. This study presents an indigenous computational tool developed for the analysis of heat transfer in regenerative cooling of such rocket engines. The developed tool incorporates a one-dimensional (1-D) combustion analysis to calculate the thermophysical properties of the combustion gas. Basic engine properties were calculated and used to generate a thrust chamber profile based on a bell-shaped nozzle. The hot gas side was analyzed using 1-D isentropic flow assumptions, along with heat transfer correlations. The coolant side was evaluated using the hydraulic analysis in the axial direction and the heat transfer analysis in the radial direction. Thermophysical properties and the phase of the coolant were determined using the given property tables and the instantaneous state of the coolant. This flexible and computationally less demanding tool was used to analyze two small-scale engines utilizing liquid hydrocarbon fuels, which are used in modern rocket propulsion. The wall cooling analyses of a liquid oxygen (LOX)/liquid methane (LCH₄) engine and a liquid oxygen (LOX)/liquid propane (LC₃H₈) engine are presented. Fuel and oxidizer were used separately as coolants for both engines, and both of them experienced phase change. Results reveal the advantage of the high mass flow rate of the oxidizer in cooling performance. In addition, the results of this study show that the cooling of the LOX/LC₃H₈ engine is somewhat more challenging compared to the LOX/LCH₄ engine. Full article

The cooling jackets of liquid rocket engines are composed of narrow passages surrounding the thrust chambers and ensure the reliable operation of the engine. Critical conditions may also be encountered, since the cooling jackets of cryogenic engines, such as those using LO_X/LCH₄ propellants, are based on a regenerative strategy, where the fuel is used as a refrigerant. Consequently, deterioration modes near where pseudocritical conditions are reached or low heat transfer coefficients where the fuel becomes a vapour and must therefore be managed. The verification of the cooling jacket behaviour to consolidate the design solutions in all the extreme points of the operating box represents a very important phase. The present paper discusses the full characterization of the HYPROB (HYdrocarbon PROpulsion test Bench Program) first unit of the final demonstrator, (DEMO-0A), by considering the working points within the limits of the operating box and comparisons with the nominal conditions are given. In this way, a full understanding of the cooling system behaviour, affecting the working of the entire thrust chamber, is accomplished. Moreover, the design strategy and choices have been confirmed, since the verifications also include potentially even more extreme conditions with respect to the nominal ones. The investigation has been numerically performed and supported the thermo-structural analyses accomplished before the final firing campaign, completed in December 2022. Since little information is available in the literature on LO_X/LCH₄ engines, suggestions are given as to the organization of the numerical simulations, which support the design of such rocket engine cooling systems. Full article