高超声速弯曲激波压缩侧压式进气道数值研究

张林; 张堃元; 王磊; 王渊

doi:10.13224/j.cnki.jasp.2015.02.004

高超声速弯曲激波压缩侧压式进气道数值研究

doi: 10.13224/j.cnki.jasp.2015.02.004

南京航空航天大学能源与动力学院江苏省航空动力系统重点实验室, 南京 210016

基金项目:

国家自然科学基金(90916029,91116001)

详细信息

作者简介:
张林(1981-),男,四川资阳人,工程师,博士生,主要从事内流气体动力学和航空发动机整机试验技术方面的研究.

中图分类号: V231.3
计量
- 文章访问数: 1361
- HTML浏览量: 0
- PDF量: 1051
- 被引次数: 0
出版历程
- 收稿日期: 2013-10-22
- 刊出日期: 2015-02-28

Numerical investigation of hypersonic curved shock sidewall compression inlet

Jiangsu Province Key Laboratory of Aerospace Power System, College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China

摘要

摘要: 采用壁面马赫数呈线性分布的曲面压缩系统改进参考侧压式进气道的顶板,得到弯曲激波压缩侧压式进气道,并与参考侧压式进气道进行了比较.数值研究结果表明:设计状态无黏时曲面压缩顶板壁面马赫数分布与给定的马赫数分布基本一致,并且有黏时其壁面压力分布也与二维曲面的基本相同;同参考侧压式进气道相比,顶板采用曲面压缩能够一定程度地改善壁面压力分布,使其末端压力梯度变化平缓;并且非设计状态下的性能也得到有效地改善,特别是来流马赫数为4时,其流量系数提高6.0%、达到0.799,喉道截面总压恢复系数提高1.9%;来流马赫数为5时,其流量系数提高5.2%、达到0.909,喉道截面总压恢复系数提高3.2%.随着攻角增大,该进气道流量捕获能力增强、隔离段出口截面流场畸变减小,但喉道截面总压恢复系数下降剧烈.
- 超燃冲压发动机 /
- 高超声速侧压式进气道 /
- 曲面压缩 /
- 马赫数线性分布 /
- 弯曲激波
Abstract: Curved surface compression system with wall Mach number linear distribution was used to improve the topwall of the reference sidewall compression inlet, then curved shock sidewall compression inlet was obtained and Compared with the reference sidewall compression inlet. The numerical result shows that: the wall Mach number distribution of topwall with curved surface is consistent with the given Mach number distribution at the design and inviscid condition, and its wall pressure distribution is basically the same as that of two-dimensional curved surface at viscous condition. Compared with the reference sidewall compression inlet, when the topwall used curved surface compression, its wall pressure distribution was improved to a certain extent, and the pressure gradient at the end of surface changed smoothly. And the performance of off-design condition is improved effectively; especially the mass flow capture ratio increases 6% and reached 0.799 at inflow Mach number of 4, and the total pressure recovery of throat section increases 1.9%. At inflow Mach number of 5, its mass flow capture ratio increases 5.2% and reached 0.909, and the total pressure recovery of throat section increases 3.2%. With the increase of the angle of attack, the inlet's ability of capturing mass flow improved and the flow field distortion of isolator exit section decreased. But the total pressure recovery of throat section decreased sharply.
- scramjet /
- hypersonic sidewall compression inlet /
- curved surface compression /
- Mach number linear distribution /
- curved shock

HTML

参考文献(15)

[1]	Trexler C A.Performance of an inlet for an integrated scramjet concept[J].Journal of Aircraft,1974,11(9):589-591.
[2]	Trexler C A.Inlet performance of the integrated langley scramjet module (Mach 2.3 to 7.6)[R].AIAA 75-1212,1975.
[3]	Holland S D,Murphy K J.An experimental parametric study of geometric,reynolds number and ratio of specific heats effects in three-dimensional sidewall compression scramjet inlets at Mach 6[R].AIAA 93-0740,1993.
[4]	Holland S D,Perkins J N.Contraction ratio effects in a generic three-dimensional sidewall compression scramjet inlet:a computational and experimental investigation[R].AIAA 91-1708,1991.
[5]	Vinogradov V,Stepanov V,Goldfeld M.Experimental and numerical investigation of two concepts of the hypersonic inlet[R].AIAA 95-2721,1995.
[6]	Goldfeld M A,Starov A V,Vinogradov V V.Experimental study of scramjet module[J].Journal of Propulsion and Power,2001,17(6):1222-1226.
[7]	Kanda T,Wakamatsu Y,Ono F,et al.Mach 8 testing of a scramjet engine model[R].AIAA 99-0617,1999.
[8]	Kanda T,Wakamatsu Y,Sakuranaka N,et al.Mach 8 testing of a scramjet engine with ramp compression[R].AIAA 2000-0616,2000.
[9]	杨进军,张堃元,徐辉,等.双模态冲压发动机高超进气道的实验研究[J].推进技术,2001,22(6):４73-475. YANG Jinjun,ZHANG Kunyuan,XU Hui,et al.Experiment study on supersonic inlet in dual-model scramjet[J].Journal of Propulsion Technology,2001,22(6):473-475.(in Chinese)
[10]	金志光.超燃冲压发动机高超侧压式进气道设计方法研究[D].南京:南京航空航天大学,2006. JIN Zhiguang.Design methodology investigation of three-dimensional sidewall compression scramjet inlet[D].Nanjing:Nanjing University of Aeronautics and Astronautics,2006.(in Chinese)
[11]	丁猛,梁剑寒,刘卫东,等.碳氢燃料超燃冲压发动机进气道与燃烧室匹配性能试验研究[J].航空学报,2005,26(1):27-31. DING Meng,LIANG Jianhan,LIU Weidong,et al.Experiment study on the interaction between inlet and combustor of hydrocarbon fueled scramjet[J].Acta Aeronautica et Astronautica Sinica,2005,26(1):27-31.(in Chinese)
[12]	金志光,张堃元.高超侧压式进气道高焓脉冲风洞实验[J].推进技术,2005,26(4):319-323. JIN Zhiguang,ZHANG Kunyuan.Test of a generic sidewall compression scramjet inlet in the high-enthalpy impulse wind tunnel[J].Journal of Propulsion Technology,2005,26(4):319-323.(in Chinese)
[13]	张林,张堃元,王磊,等.新型高超弯曲激波二维进气道初步研究[J].航空动力学报,2011,26(11):2556-2562. ZHANG Lin,ZHANG Kunyuan,WANG Lei,et al.Preliminary study on hypersonic curved shock two-dimensional inlet[J].Journal of Aerospace Power,2011,26(11):2556-2562.(in Chinese)
[14]	张林,张堃元,王磊,等.基于壁面马赫数梯度的高超声速弯曲激波二维进气道数值研究[J].航空动力学报,2013,28(4):752-758. ZHANG Lin,ZHANG Kunyuan,WANG Lei,et al.Computational investigation of hypersonic curved shock two-dimensional inlet with compression surface constant Mach number gradient[J].Journal of Aerospace Power,2013,28(4):752-758.(in Chinese)
[15]	金志光,张堃元,刘媛.马赫数4~7的高超侧压式进气道气动设计与性能[J].航空动力学报,2011,26(6):1201-1208. JIN Zhiguang,ZHANG Kunyuan,LIU Yuan.Design and performance investigation of sidewall compression scramjet inlets operating from Mach 4 to Mach 7[J].Journal of Aerospace Power,2011,26(6):1201-1208.(in Chinese)