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

Shigeru Aso; Arif Nur Hakim; Shingo Miyamoto; Kei Inoue; Yasuhiro Tani

doi:10.1016/j.actaastro.2005.03.055

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

Shigeru Aso, Arif Nur Hakim, Shingo Miyamoto, Kei Inoue, Yasuhiro Tani

Space Systems Engineering

Research output: Contribution to journal › Conference article › peer-review

33 Citations (Scopus)

Abstract

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 μs. 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.

Original language	English
Pages (from-to)	384-389
Number of pages	6
Journal	Acta Astronautica
Volume	57
Issue number	2-8
DOIs	https://doi.org/10.1016/j.actaastro.2005.03.055
Publication status	Published - Jul 2005
Event	Infinite Possibilities Global Realities, Selectred Proceedings of the 55th International Astronautical Federation Congress, Vancouver - Duration: Oct 4 2004 → Oct 8 2004

All Science Journal Classification (ASJC) codes

Aerospace Engineering

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10.1016/j.actaastro.2005.03.055

Cite this

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title = "Fundamental study of supersonic combustion in pure air flow with use of shock tunnel",

abstract = "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 μs. 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.",

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AU - Aso, Shigeru

AU - Hakim, Arif Nur

AU - Miyamoto, Shingo

AU - Inoue, Kei

AU - Tani, Yasuhiro

PY - 2005/7

Y1 - 2005/7

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AB - 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 μs. 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.

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SN - 0094-5765

VL - 57

SP - 384

EP - 389

JO - Acta Astronautica

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T2 - Infinite Possibilities Global Realities, Selectred Proceedings of the 55th International Astronautical Federation Congress, Vancouver

Y2 - 4 October 2004 through 8 October 2004

ER -

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

Abstract

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