Fundamental study of supersonic combustion in pure air flow with use of shock tunnel

被引：20
|
作者：
Aso, S [1 ]
Hakim, AN [1 ]
Miyamoto, S [1 ]
Inoue, K [1 ]
Tani, Y [1 ]
机构：
[1] Kyushu Univ, Dept Aeronaut & Astronaut, Fukuoka 812, Japan
来源：
ACTA ASTRONAUTICA | 2005年 / 57卷 / 2-8期
关键词：
D O I：
10.1016/j.actaastro.2005.03.055
中图分类号：
V [航空、航天];
学科分类号：
08 ; 0825 ;
摘要：
An experimental study using reflected-type of shock tunnel has been conducted to investigate the phenomena of supersonic combustion. In the experiment, test air is compressed by reflected shock wave up to stagnation temperature of 2800 K and stagnation pressure of 0.35 MPa. Heated air is used as a reservoir gas of supersonic nozzle. Hydrogen is injected transversely through circular hole into freestream of Mach 2. Flow duration is 300 mu s. Schlieren method and CCD UV camera are used to obtain information on the shock structures and the region of combustion. The effects of total pressure of injection gas to the fuel penetration and the region of combustion have been obtained. (c) 2005 Elsevier Ltd. All rights reserved.
引用
收藏
页码：384 / 389
页数：6
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