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Supersonic film cooling with and without induced shock waves was investigated numerically. The computations used a freestream Mach number of 3.13 and tangential-slot injection-cooling-stream Mach numbers of 1.3 and 1.8. Parameters investigated included shock wave intensity and coolant properties (nitrogen, methane, and helium). The results indicate that the adiabatic effectiveness is reduced by the impinging oblique shock wave because the shock wave reduces the local Mach number in the coolant layer, which increases the adiabatic-wall temperature. In addition, the oblique shock wave enhances the mixing of the cooling stream with the freestream, especially for stronger shock waves or lighter-gas coolants, which also increases the adiabatic-wall temperature. For the same injection-cooling-stream Mach number, the effects of the oblique shock wave are stronger in helium; however, helium had a higher adiabatic effectiveness than methane and nitrogen, even though the helium mass flow rate was lower than that of either methane or nitrogen. The adiabatic effectiveness is reduced much more by the 10 deg shock waves than by either the 7 or 4 deg shock waves. [ABSTRACT FROM AUTHOR]

The three-dimensional modal linear stability analysis is performed for the near-wall flow of a full-scale generic hypersonic vehicle forebody at flight Mach numbers 4, 6, and 8, at different angles of attack. The mean flow is computed with a Navier-Stokes commercial code. A physically sound, computationally efficient original method is proposed to define the integration path in the eN method. It has a significant impact on the computed N factors. The entropy-layer effect on the flow instability is analyzed in the framework of Lees and Lin's asymptotic theory. The entropy layer introduces an additional unstable mode, for which N ≈ 4. Results show that crossflow instability is dominant at Mach 6 and 8, whereas Mack's oblique first mode prevails at Mach 4. This mode is stabilized by a radiating wall, compared with an adiabatic one. Mack's second mode is also present at Mach 8. In any case, none of the instability modes that have been found is strong enough to provoke a natural transition in flight by itself: computed N factors do not exceed 10. Attempts to correlate the results with the ReθMe = Const. criterion are discussed. [ABSTRACT FROM AUTHOR]

A laminar flame propagating toward the open end of a narrow channel filled with a gaseous combustible mixture can accelerate or oscillate, depending on the wall temperature and the channel width. The accelerating flame is able to produce a high-speed flow that has the potential to provide significant thrust. We study these phenomena using multidimensional reactive Navier-Stokes numerical simulations, and show that for adiabatic walls, the maximum flame acceleration occurs when the the channel is about five times larger than the reaction zone of a laminar flame. In three-dimensional square channels, the flame speed increases roughly two times faster than in two-dimensional channels. The accelerating flame generates weak compression waves that propagate with a local sound speed and can accelerate the unreacted material ahead of the flame to velocities close to the speed of sound without creating strong shocks. This combustion regime is of particular interest for micropropulsion because it allows efficient use of fuel and gradual development of thrust. [ABSTRACT FROM AUTHOR]