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Direct numerical simulations of the Navier-Stokes equations have been performed to investigate the receptivity and breakdown mechanisms in a Mach 6 flow over a generic forebody geometry with freestream acoustic disturbances. The simulations are based on transition experiments carried out in April 2015 in the Boeing/U.S. Air Force Office of Scientific Research Mach 6 facility at Purdue University. A three-dimensional model for both fast and slow freestream acoustic waves with multiple frequencies and spanwise wave numbers has been adopted in the numerical simulations, for which high-amplitude disturbances have been considered in order to simulate noisy wind-tunnel conditions. The numerical results revealed similarities in comparison to the experimental observations, especially when slow acoustic waves were considered as freestream disturbances. In particular, slow acoustic waves have been found to induce the breakdown process via crossflow instabilities located in the off-centerline region, with formation of streamwise streaks. Fast acoustic waves, in contrast, appeared more efficient in inducing earlier nonlinear growth through destabilization of the boundary layer along the symmetry plane of the body. [ABSTRACT FROM AUTHOR]

With the objective of inserting laminar-turbulent transition prediction inside computational fluid dynamics codes (Navier-Stokes solvers and boundary-layer codes), a physically based transition criterion for laminar-turbulent longitudinal transition caused by the amplification of boundary-layer instabilities is developed and implemented into both the ONERA-The French Aerospace Lab's elsA and the 3C3D codes. This transition criterion is an extended version of the incompressible Arnal-Habiballah-Delcourt criterion, which is here applicable in a range of Mach numbers from zero to four, and takes into account wall temperature effects. Larger Mach numbers, up to 4.5, may even be considered for low-pressure gradient cases. An example of application, at the very upper limit in Mach number, is also presented. This criterion covers the entire range where the first mode of instability may be the most amplified. A second model, to be created, will be necessary to take into account the second-mode contribution for larger Mach numbers. [ABSTRACT FROM AUTHOR]

In hypersonic flows, it is often necessary to be able to trip the transition to turbulence, upstream of air intakes, for example. Direct numerical simulations have been performed to identify the roughness-induced transition mechanisms on a wedgelike forebody at Mach 6 and unit Reynolds number Re=11 million (/m). Good agreement with the experiments performed in the Boeing/Air Force Office of Scientific Research Mach-6 Quiet Tunnel at Purdue University was obtained in terms of wall heat-flux and wall-pressure fluctuations. First, an isolated roughness was considered. The presence of the roughness in the span-inhomogeneous base flow leads to the formation of a crossflowlike vortex. High-frequency secondary instabilities of the stationary crossflow vortex are observed in the wake and are found to be responsible for the breakdown to turbulence. Spatial linear modal instability analysis of this flow has been performed at selected streamwise locations. The linear stability approach is found to give accurate predictions in terms of mode shapes, most-amplified disturbance frequencies, and growth rates, as it only underpredicts the N-factor of the most unstable mode by 10% compared to the direct numerical simulations. Unsteady simulations were then carried out for a trip array configuration and showed that it does not change the transition mechanisms, but the frequencies of the most unstable secondary instabilities were found to be higher. [ABSTRACT FROM AUTHOR]