Análise aero-estrutural de um demonstrador da combustão supersônica
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 airbreathing propulsion systems (in general, including scramje...
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| Formato: | Dissertação |
| Idioma: | pt_BR |
| Publicado em: |
Universidade Federal do Rio Grande do Norte
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| Endereço do item: | https://repositorio.ufrn.br/handle/123456789/51270 |
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| Resumo: | 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 airbreathing propulsion systems (in
general, including scramjet technology) that use atmospheric air as an oxidant. During
hypersonic flight, aerospace vehicles with airbreathing hypersonic propulsion are subject to
high aerodynamic and thermal loads. In this context, in the present work, the main objective
is to carry out an aero-structural analysis of a generic supersonic combustion demonstrator,
under flight conditions at an altitude of 23 km and a speed of 1723 m/s, corresponding to
Mach number 5.8. In the structural analysis, cases with different plate thicknesses (6 mm, 4
mm, 3 mm and 2.5 mm) were considered and the elements that make up the scramjet are
stainless steel 304 (beams and ribs), aluminum 7075 (ramps and panel side), inconel 718 or
tungsten (leading edges and combustion chamber entrance). For the execution of the
structural analysis, an aerodynamic and dimensional design of a generic scramjet was
carried out, idealized for coupling to the Brazilian rocket engines S30 and S31.
Optimization criteria were applied to the compression section, aiming to reach the
temperature and Mach number conditions required at the entrance of the combustion
chamber to spontaneously burn hydrogen. In the expansion section, the optimization
criterion is based on verifying the point at which the pressure condition is equivalent to the
freestream, defining the region where the coupling to the accelerating vehicle should be
performed. The aerodynamic load was defined from analytical and numerical analysis,
considering the air as a calorically perfect gas and neglecting viscous effects. In the
aerodynamic design and analysis, the cases without fuel injection were evaluated, therefore
no fuel burning (power-off) and with fuel injection and burning (power-on), but in the
structural analysis only power-on was considered. The numerical analysis of the flow and
the 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 at a speed of 1723 m/s, the scramjet with three compression ramps,
with deflection angles of 7.48°, 8.93° and 10.77° was able to generate, at the entrance to the
combustion chamber, a speed corresponding to Mach number 1.71 and a static temperature
of 1071.25 K, greater than 845.15 K, demonstrating the possibility of burning hydrogen. On
the trailing edge, the flow velocity was 1688.96 m/s without injection and no fuel burning
and 1806.98 m/s with fuel injection and burning, greater than 1723 m/s, demonstrating that
the scramjet is only capable of generating thrust upon ignition of the fuel. For the numerical
analysis of the flow, the unstructured mesh with triangular elements was more adequate to
capture the flow conditions after the oblique shock waves established in the compression
section of the scramjet, considering the atmospheric air as calorically perfect gas and
without viscous effects. In the aerodynamic analysis, the numerical results showed good
agreement with the analytical results. In the structural analysis, it was verified that the
maximum value of von Mises equivalent stress is lower than the yield stress of the
materials used for cases with plates with a thickness of 3 mm or greater. Under these
conditions, the structure works in an elastic regime, so that the strains are recoverable if the
loads are removed. Only with 2.5 mm plates, flow was verified from the internal structure
stringers in contact with the combustor surfaces, the region in which the structure is most requested by static pressure loads due to the addition of heat, which simulates the burning
of fuel. In addition, inconel 718 is more suitable than tungsten for application on the
demonstrator's leading edges, providing better mechanical capacity and lower weight, and
for this reason being more advantageous for the scramjet's aero-structural design. |
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