Research on Computational Method of Supersonic Inlet/Isolator Internal Flow

In this research, a CFD solver is developed for solving the 2D/3D compressible flow problem: the finite volume method based on multi-block structural grids is used to solve the compressible Reynolds averaged Navier–Stokes equations (RANS). Included in the methodology are multiple high-order reconstruction schemes, such as the 3rd-order MUSCL (Monotone Upstreamcentered Schemes for Conservation Laws), 5th-order WENO (Weight Essentially Non-Oscillatory), and 5th-order MP (Monotonicity-Preserving) schemes. Of the variety of turbulence models that are embedded, this solver is mainly based on the shear stress transport model (SST), which is compatible with OpenMP/MPI parallel algorithms. This research uses the CFD solver to conduct steady-state flow simulation for a two-dimensional supersonic inlet/isolator, incorporating these high-precision reconstruction schemes to accurately capture the shock wave/expansion wave interaction and shock wave/turbulent boundary layer interaction (SWTBLI), among other effects. By comparing the 2D/3D computation results of the same inlet configuration, it is found that the 3D effects of the side wall cannot be ignored due to the existing strong lateral flow near the corner. To obtain a more refined turbulence simulation, the commercial software ANSYS Fluent 18.0 is used to carry out the detached eddy simulation (DES) and the large eddy simulation (LES) of the same supersonic inlet, so as to reveal the flow details near the separation area and boundary layers.

Experimental and numerical investigation on a supersonic inlet with large bleed window

The design of a two-dimensional supersonic inlet with large bleed window at low Mach number was developed. Numerical simulation and wind tunnel experiments were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the single-degree-of-freedom variable geometry scheme adopted in this paper guarantees the steady work of the inlet over a wide speed range. The large bleed window caused by rotation of the compression ramp appears near the throat at low Mach number. Low-pressure airflow near the bleed window neutralises the original high-pressure airflow behind the shock train, which decreases the overall pressure of the downstream region of the internal contraction section. To match the lower pressure, the structure of the shock train changes from strong $\lambda$ -type to weak $\lambda$ -type, and finally to a normal shock wave as backpressure increases at Mach number 2.5. Herein, the total pressure recovery coefficient of the inlet near the critical condition improves by 8.5% as the backpressure ratio (Pe/P0) adds from 13 to 14.6 at Mach number 2.5. It proves that the scheme is effective on terminal shock wave control and inlet performance improvement. In addition, due to the background wave and the bleed window, two kinds of shock wave oscillation occur when the backpressure ratio is 13.1.

Numerical analysis on the conceptual design of an air-breathing engine inlet working in mach number 3∼5.5

Based on the equal-intensity shock theory, this article designed a supersonic inlet working in Mach number 3.0∼5.5 with the background of an air-breathing engine. The inlet applied the four-shock train mixed compression configuration and inserted a sidewall compression at the beginning of the isolator. Through developing effective 3D RANS computations validated by current experiments, the performance of the designed inlet was identified. The designed inlet self-starts at freestream Mach number Ma∞ = 3.0 under which the total pressure recovery coefficient has dramatic increment, and the aerodynamic choking at the inlet throat no longer presents; the inlet keeps working at all studied flight states with zero angle of attack (AoA) and achieves shock-on-lip at the design point Ma∞ = 5.0. Both positive and negative AoAs can disrupt the equal-intensity shock allocations, which degrade the inlet performance. The inlet obtains maximum total pressure recovery coefficient at zero AoA. The maximum back pressure at Ma∞ = 3.0∼5.5 obtained by the inlet surpasses the requirements and keeps a certain margin. The inlet performance basically meets all the goals proposed by the engine.