TY - GEN
T1 - Numerical experiments of scramjet combustion with boundary-layer bleeding
AU - Kouchi, Toshinori
AU - Mitani, Tohru
AU - Kodera, Masatoshi
AU - Masuya, Goro
PY - 2003
Y1 - 2003
N2 - Airframe-integrated scramjet engines swallow the boundary-layer which develops on the airframe of space planes. The scramjet engine easily falls into engine stall (engine unstart) due to the boundary-layer separation resulting from combustion. In this study, to investigate the unstart characteristics, numerical simulations of a scramjet engine with boundary-layer bleeding were performed using a reacting flow code which includes a one-equation turbulence model. This computation is calibrated by the experimental wall pressure distributions, heat fluxes and thrusts. The computation reproduces the prevention of engine unstart in the combustion tests with bleeding. Bleeding of 0.6% in a captured flow rate suppresses the flow separation and extends the start limit from the fuel equivalence ratio (φ) of 0.5 to 1.0. The computation at φ=1.0 shows that small-scale circular diffusion flames are anchored around individual fuel jets near the injectors. These structures disappear to form a large-scale envelope diffusion flame downstream of the combustor. The circular flames near the injectors account for 80% of the combustion efficiency and control the thrust performance.
AB - Airframe-integrated scramjet engines swallow the boundary-layer which develops on the airframe of space planes. The scramjet engine easily falls into engine stall (engine unstart) due to the boundary-layer separation resulting from combustion. In this study, to investigate the unstart characteristics, numerical simulations of a scramjet engine with boundary-layer bleeding were performed using a reacting flow code which includes a one-equation turbulence model. This computation is calibrated by the experimental wall pressure distributions, heat fluxes and thrusts. The computation reproduces the prevention of engine unstart in the combustion tests with bleeding. Bleeding of 0.6% in a captured flow rate suppresses the flow separation and extends the start limit from the fuel equivalence ratio (φ) of 0.5 to 1.0. The computation at φ=1.0 shows that small-scale circular diffusion flames are anchored around individual fuel jets near the injectors. These structures disappear to form a large-scale envelope diffusion flame downstream of the combustor. The circular flames near the injectors account for 80% of the combustion efficiency and control the thrust performance.
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U2 - 10.2514/6.2003-7038
DO - 10.2514/6.2003-7038
M3 - Conference contribution
AN - SCOPUS:85086488934
SN - 9781624100857
T3 - 12th AIAA International Space Planes and Hypersonic Systems and Technologies
BT - 12th AIAA International Space Planes and Hypersonic Systems and Technologies
PB - American Institute of Aeronautics and Astronautics Inc.
T2 - 12th AIAA International Space Planes and Hypersonic Systems and Technologies 2003
Y2 - 15 December 2003 through 19 December 2003
ER -