Abstract
The turbine-based combined cycle (TBCC) engine, as a hypersonic air-breathing propulsion system, involves the transition between multiple bypasses during flight. During the transition process, the flow field of the combustor is affected. In this paper, the flow field of a TBCC multibypass combustor with typical structures is investigated experimentally. The effect of the combined structure on the mixing performance, total pressure loss and flow field distribution downstream of the TBCC multibypass combustor is investigated by changing the inlet aerodynamic parameters and structural parameters of the combustor. The results show that the diffusion ratio increased from 1.0 to 1.4, the thermal mixing efficiency decreased by 24.7%, and the total pressure recovery coefficient increased by 1.86%. Meanwhile, the recirculation zone is wider, and the velocity distribution behind the stabilizer is more uniform. The thermal mixing efficiency increases by 55.8% with the increase in the down inclination (35°∼45°) of the lobe mixer. The mixing characteristic of the triple-bypass combustor is greater than that of the double-bypass combustor. There is an obvious “phase difference” of the vortex shedding in the double-bypass combustor at different spanwise sections, while it does not exist in the triple-bypass combustor. The turbulence disturbance degree in the recirculation zone behind the stabilizer in the triple-bypass combustor is slightly higher than that in the double-bypass combustor, but the velocity distribution is more uniform. The increase in inlet Mach number increases the flow losses and decreases the mixing performance. However, the thermal mixing efficiency is positively impacted with respect to the inlet temperature. In addition, the turbulent kinetic energy of the shear layer behind the stabilizer is large, while it is relatively small in the recirculation zone. There is not only spanwise vortices but also radial vortices behind the stabilizer.
Published Version
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