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

Abstract 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.

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Experimental and numerical investigation of a supersonic inlet with double S-bend diffuser

Proceedings of the Institution of Mechanical Engineers Part G Journal of Aerospace Engineering ◽

10.1177/0954410020983981 ◽

2021 ◽

pp. 095441002098398

Author(s):

Jinsheng Zhang ◽

Huacheng Yuan ◽

Yunfei Wang ◽

Guoping Huang

Keyword(s):

Mach Number ◽

Total Pressure ◽

Aerodynamic Performance ◽

Shock Train ◽

Recovery Coefficient ◽

Low Mach Number ◽

Supersonic Inlet ◽

Total Pressure Recovery ◽

High Mach ◽

Pressure Recovery Coefficient

Design of a supersonic inlet with double S-bend diffuser was developed. Numerical simulation and wind tunnel experiment were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the variable geometry scheme adopted solves the contradiction between starting performance at low Mach number and aerodynamic performance at high Mach number. The inlet works normally and stably over a wide speed range. At design point, the total pressure recovery coefficient reaches 0.47. In addition, two different kinds of inlets with double S-bend diffuser and single S-bend diffuser were studied. Compared with the double S-bend diffuser, the total pressure recovery coefficient of the single S-bend diffuser is higher at low Mach number (Ma0 < 3) and lower at high Mach number (Ma0 > 3). With the increase of backpressure, shock train mainly moves upstream along the low-energy flow region in the diffuser. For the double S-bend diffuser, shock train will first move along the lower corner and then along the upper corner. For the single S-bend diffuser, it will only move along the upper corner. The strong secondary flow of the double S-bend is the main reason for the above phenomenon.

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

Proceedings of the Institution of Mechanical Engineers Part G Journal of Aerospace Engineering ◽

10.1177/09544100211011346 ◽

2021 ◽

pp. 095441002110113

Author(s):

Chao Huo ◽

Zhenhua Yang ◽

Zhengze Zhang ◽

Peijin Liu

Keyword(s):

Mach Number ◽

Total Pressure ◽

Back Pressure ◽

Pressure Recovery ◽

Shock Train ◽

Recovery Coefficient ◽

Air Breathing ◽

Supersonic Inlet ◽

Total Pressure Recovery ◽

Pressure Recovery Coefficient

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.

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Three-Dimensional Passive and Active Control Methods of Shock Wave Train Physics in a Duct

International Journal of Applied Mechanics ◽

10.1142/s1758825116500472 ◽

2016 ◽

Vol 08 (04) ◽

pp. 1650047 ◽

Cited By ~ 15

Author(s):

Reza Kamali ◽

Seyed Mahmood Mousavi ◽

Danial Khojasteh

Keyword(s):

Shock Wave ◽

Mach Number ◽

High Performance ◽

Wave Train ◽

Three Dimensional ◽

Maximum Flow ◽

Control Methods ◽

Shock Train ◽

Ansys Fluent ◽

Starting Point

In the present work, the physics of a three-dimensional shock train in a convergent-divergent nozzle is numerically investigated. In this regards, the Ansys-Fluent Software with Algebraic Wall-Modeled Large-Eddy Simulation (WMLES) is used. To estimate precision and errors accumulation we used the Smirinov’s method; fine flow structures are obtained via Laplacian of density called shadowgraph and the shock parameter is defined as multiplication of flow Mach number by the normalized pressure gradient, in which shock wave structures are visible distinctly. The results are compared with the experimental data of Weiss et al. [Experiments in Fluids 49(2) (2010) 355–365], in the same conditions including geometry, boundary conditions, etc. The results show that there is good agreement with experimental trends concerning wall pressure and centerline Mach number profiles. Therefore, the focus of the present study is an assessment of various flow control methods to change the shock structures. Consequently, we investigated the effects of passive (bump and cavity) and active (suction and blowing) control methods on the starting point of shock, shock strength, minimum pressure, maximum flow Mach number, etc. All CFD investigations are carried out by High Performance Computing Center (HPCC).

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