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

Published online by Cambridge University Press:  08 September 2021

HC. Yuan ,
JS. Zhang Open the ORCID record for JS. Zhang [Opens in a new window] ,
YF. Wang  and
GP. Huang
HC. Yuan*
Affiliation:
College of Energy and Power Engineering Nanjing University of Aeronautics and Astronautics, Nanjing210016, China
JS. Zhang
Affiliation:
College of Energy and Power Engineering Nanjing University of Aeronautics and Astronautics, Nanjing210016, China
YF. Wang
Affiliation:
College of Energy and Power Engineering Nanjing University of Aeronautics and Astronautics, Nanjing210016, China
GP. Huang
Affiliation:
College of Energy and Power Engineering Nanjing University of Aeronautics and Astronautics, Nanjing210016, China
*
E-mail: yuan_hch7@sina.com
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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.

Keywords

two-dimensional variable inletlarge bleed windowterminal shock wavenumerical simulationwind tunnel experiment
Type
Research Article
Information
The Aeronautical Journal , Volume 126 , Issue 1297 , March 2022 , pp. 425 - 449
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Copyright
© The Author(s), 2021. Published by Cambridge University Press on behalf of Royal Aeronautical Society

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References

Walker, B, Kennedy, K and Mikkelson, C. US army hypersonic scramjet propelled missile technology program, AIAA-2006-7927, 2006.CrossRefGoogle Scholar
Heiser, WH, Pratt, DT, Daley, DH and Mehta, UB. Hypersonic airbreathing propulsion, AIAA Paper Education Series, Washington, DC, 1994.CrossRefGoogle Scholar
Ding, F, Liu, J, Shen, CB, Huang, W, Liu, Z and Chen, SH. An overview of waverider design concept in airframe/inlet integration methodology for air-breathing hypersonic vehicles, Acta Astronautica, 2018, 152, pp 639–656.CrossRefGoogle Scholar
Liu, JB, Fan, XQ, Tao, Y and Liu, WD. Experimental and numerical study on the local unstart mechanism of hypersonic inlet, Acta Astronautica, 2019, 160, pp 216–221.CrossRefGoogle Scholar
Wang, WX and Guo, RW. Numerical study of unsteady starting characteristics of a hypersonic inlet, Chinese Journal of Aeronautics, 2013, 26, (3), pp 563–571.CrossRefGoogle Scholar
David, WW, Lawrence, DH and Carl, AT. Propulsion airframe integration test techniques for hypersonic airbreathing configurations at NASA Langley Research Center, AIAA-2003-4406, 2003.Google Scholar
Gamble, E, Haid, D, D’Alessandro, S and Defrancesco, R. Dual-Mode Scramjet Performance Model for TBCC Simulation, AIAA-2009-5298, 2009.CrossRefGoogle Scholar
Albertson, CW, Saied, E and Trexler, CA. Mach 4 test results of a dual-flowpath, turbine-based combined cycle inlet, AIAA-2006-8138, 2006.CrossRefGoogle Scholar
Liu, J, Yuan, HC, Hua, ZX, Chen, WF and Ge, N. Experiment and numerical investigation of smooth turbine-based combined-cycle inlet mode transition, Aerospace Science and Technology, 2017, 60, pp 124–130.CrossRefGoogle Scholar
Liu, Y, Wang, L and Qian, ZS. Numerical investigation on the assistant restarting method of variable geometry for high Mach number inlet, Aerospace Science and Technology, 2018, 79, pp 647–657.CrossRefGoogle Scholar
Shi, W, Chang, JT, Zhang, JL, Wang, YY, Hou, WX and Bao, W. Numerical investigation on behaviors of shock train in a hypersonic inlet with translating cowl, Acta Astronautica, 2018, 152, pp 682–691.CrossRefGoogle Scholar
Cai, JH, Zhou, J, Liu, SJ and Lin, ZY. Effects of dynamic backpressure on shock train motions in straight isolator, Acta Astronautica, 2017, 141, pp 237–247.CrossRefGoogle Scholar
Zhang, KY. Research progress of hypersonic inlet reverse design based on curved shock compression system, Acta Aeronautica et Astronatica Sinica, 2015, 36, (1), pp 274–288.CrossRefGoogle Scholar
Yu, ZH, Huang, GP and Xia, C. 3D inverse method of characteristics for hypersonic bump-inlet integration, Acta Astronautica, 2020, 166, pp 11–22.CrossRefGoogle Scholar
Qiao, WY, Yu, AY, Gao, W and Wang, W. Design method with controllable velocity direction at throat for inward-turning inlets, Chinese Journal of Aeronautics, 2019, 32, (6), pp 1403–1415.CrossRefGoogle Scholar
Zucrow, MJ and Hoffman, JD. Gas Dynamics, New York, 1976.Google Scholar
Zhang, KY. Research progress of hypersonic inlet reverse design based on curved shock compression system, Acta Aeronautica et Astronautica Sinica, 2015, 36,(1), pp 274–288.CrossRefGoogle Scholar
Zuo, FY and Huang, GP. A preliminary overview analysis on the internal waverider inlets for ramjet, AIAA-2017-2420.Google Scholar
Lee, CC and Boedicke, C. Subsonic diffuser design and performance foe advanced fighter aircraft, AIAA-85-3073, 1985.CrossRefGoogle Scholar
Sun, S. Design and Investigation for Unconventional Inlets, Nanjing University of Aeronautics and Astronautics, 2006.Google Scholar
Haberle, J and Gulhn, A. Investigation of the performance of a Scramjet Inlet at Mach 6 with Boundary Layer Bleed, AIAA-2006-28139, 2006.CrossRefGoogle Scholar
Chang, JT, Bao, W, Cui, T and Yu, DR. Effect of suctions on maximum backpressure ratios of hypersonic inlets, Journal of Aerospace Power, 2008, 23, (3), pp 505–509.Google Scholar
Yuan, HC and Liang, DW. Effect of suction on starting of hypersonic inlet, Journal of Propulsion Technology, 2006, 27, (6), pp 525–528.Google Scholar
John, D and Anderson, JR. Computational Fluid Dynamics: The Basics with Applications, McGraw-Hill Companies, Inc, NewYork, 1995, pp 82–87.Google Scholar
Roe, PL. Approximate Riemann solvers, Parameter Vectors, and difference scheme, Journal of Computational Physics, 1981, 43, (2), pp 357–372.CrossRefGoogle Scholar
Van, LB. Towards the ultimate conservative difference scheme. V.A second order sequel to Godunov’s method, Journal of Computational Physics, 1979, 32, (1), pp 101–136.Google Scholar
Spalart, PR and Allmaras, SR. A one-equation turbulence model for aerodynamic flows, AIAA-92-0439.Google Scholar
Tamaki, T, Tomita, Y and Yamane, R. A study of pseudo-shock: 1st report, $\lambda$-type pseudo-shock, Bulletin of JSME, 1970, 13, (55), pp 51–58.CrossRefGoogle Scholar
Tamaki, T, Tomita, Y and Yamane, R. A study of pseudo-shock: 2nd report, X-type pseudo-shock, Bulletin of JSME, 1971, 14, (74), pp 807–817.CrossRefGoogle Scholar
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