ABSTRACT

In the current scenario of the aerospace sector, there is a great limitation related to the payload that can be launched into orbit or beyond. Rocket engines, propulsive technology in operation, have a low specific impulse compared to systems with airbreathing propulsion (scramjet technology) that use atmospheric air as an oxidant. During hypersonic flight, aerospace vehicles with hypersonic airbreathing propulsion are subject to high aerodynamic and thermal loads. In this context, in the present work the main objective is to perform a structural analysis of a generic supersonic combustion demonstrator, under flight conditions at an altitude of 23 km and speed corresponding to Mach 5,8. To carry out the structural analysis, an aerodynamic and dimensional design of a generic scramjet was carried out, designed for coupling to the national rocket engines S30 and S31. Optimization criteria were applied to the compression section, aiming to achieve the required temperature and Mach number conditions at the entrance of the combustion chamber to spontaneously burn hydrogen. In the expansion section, the optimization criterion is based on checking the point at which the pressure condition is equivalent to that of free flow, defining the region where the coupling to the accelerator vehicle must be carried out. The aerodynamic load was defined from analytical and numerical aerodynamic analysis, considering air as a calorically perfect gas and neglecting viscous effects. The design and aerodynamic analysis evaluated the case without fuel burning (power-off) and with fuel burning (power-on), but in the structural analysis only power-on was considered. Numerical flow analysis and numerical structural analysis were respectively performed in the Fluent and Static Structural modules of the Ansys software. The aerodynamic analysis showed that flying at an altitude of 23 km with a speed of 1723 m/s, the scramjet with three compression ramps with deflection angles of 7,48°, 8,93° and 10,77° is capable of generating, at the entrance to the combustion chamber, speed corresponding to Mach number 1,709 and static temperature of 1071,255 K, demonstrating the possibility of burning hydrogen. At the trailing edge, the flow velocity is 1688,958 m/s without fuel burning and 1806,977 m/s with fuel burning, demonstrating that the scramjet is only capable of generating thrust with fuel ignition. For the numerical analysis of the flow, the unstructured mesh with triangular elements proved to be more suitable to capture the flow conditions after the oblique shock waves in the scramjet, considering atmospheric air as a calorically perfect gas and without viscous effects. In the aerodynamic analysis the numerical results showed good agreement with the analytical results. In the numerical structural analysis, the maximum von-Mises equivalent stress is 122,93 MPa and occurs at the leading edge of the fairing, close to contact with the side panel, due to the thickness of the wedge at the leading edge and the high pressures of the chamber of combustion. However, this value is lower than the yield stress of the materials used, so that strains and displacements in the structure occur in the elastic regime of the materials and are therefore recoverable.