The growing demand for increasing the engine power of a liquid rocket is driving the development of high-power De-Laval nozzles, which is primarily achieved by increasing the expansion ratio. A high-expansion-ratio for De-Laval nozzles can cause flow separation, resulting in unsteady, asymmetric forces that can limit nozzle life. To enhance nozzle performance, various separation control methods have been proposed, but no methods have been fully implemented thus far due to the uncertainties associated with simulating flow phenomena. A numerical study of a high-area-ratio rocket engine is performed to analyze the aeroelastic performance of its structure under flow separation conditions. Based on numerical methodology, the flow inside a rocket nozzle (the VOLVO S1) is analyzed, and different separation patterns are comprehensively discussed, including both free shock separation (FSS) and restricted shock separation (RSS). Since the location of the flow separation point strongly depends on the turbulence model, both the single transport equation and two-transport-equation turbulence models are simulated, and the findings are compared with the experimental results. Therefore, the Spalart–Allmaras (SA) turbulence model is the ideal choice for this rocket nozzle geometry. A wavelet is used to analyze the amplitude frequencies from 0 to 100 Hz under various pressure fluctuation conditions. Based on a clear understanding of the flow field, an aeroelastic coupling method is carried out with loosely coupled computational fluid dynamics (CFD)/computational structural dynamics (CSD). Some insights into the aeroelasticity of the nozzle under separated flow conditions are obtained. The simulation results show the significant impact of the structural response on the inherent pressure pulsation characteristics resulting from flow separation.

The accuracy and efficiency of reacting-flow computations in supersonic combustors directly influence the performance analysis and optimal design of hypersonic propulsion systems. To address the substantial cost of detailed chemistry integration, this work proposes a chemistry integration surrogate method based on Kolmogorov–Arnold networks (KANs). Through learnable univariate function structures and locally adjustable B-spline basis functions, the method achieves high-accuracy prediction of the complex nonlinear mapping between thermochemical states and species increments with fewer parameters, replacing the solution of the conventional chemical source term. The results show that, at comparable parameter counts, the KAN achieves lower test loss and lower ignition delay prediction error. Even with 10.5 times as many parameters as the KAN, the baseline multilayer perceptron has a higher test loss, indicating the parameter efficiency of the KAN. In the three-dimensional supersonic combustor case, both the KAN and the baseline model reproduce the spatial distribution of the temperature field and the major species fields. Compared with this multilayer perceptron, the KAN yields a lower domain-averaged absolute temperature error at each evaluation instant, together with smaller absolute errors and narrower error distributions for the major species. The KAN delivers a lower test loss, higher three-dimensional flow-field accuracy and satisfactory continuous integration stability with fewer parameters, which makes it suitable for efficient combustion chemistry computation.

The increasing demand for higher power in rocket launchers is driving the development of higher power nozzles, achieved essentially by increasing the expansion ratio. However, this can lead to flow separation and resulting unsteady, asymmetrical forces, called side loads, which can limit the life of both the nozzle itself and other engine components. If this problem can be overcome, engine performance can be significantly increased. Therefore, various separation control methods have been proposed, but none have been fully implemented to date due to the uncertainties associated with modeling and predicting the flow phenomena. A numerical study of high area ratio rocket engine is presented to analyze the aeroelastic performance of the structure under the condition of flow separation. The finite volume method, the 2nd order upwind scheme and the extended wall function are used to analyze the flow inside the rocket nozzle and the transition of the Free Shock Separation (FSS) to Restricted Shock Separation (RSS) separation pattern ) to be discussed in detail. Since the location of the flow separation point depends heavily on the flow.