Hypersonic boundary layer stability on a flared-cone model at angle of attack

An experimental investigation of the effects of angle of attack on hypersonic boundary-layer stability on a flared-cone model was conducted in the low-disturbance Mach-6 Nozzle-Test-Chamber Facility at NASA Langley Research Center. This unique facility provided a "quiet" flow test environment which is well suited for stability experiments because the low levels of freestream "noise" minimize artificial stimulation of flow-disturbance growth. Surface pressure and temperature measurements documented the adverse-pressure gradient and transition-onset location on the model. Hot-wire anemometry diagnostics were applied to identify the boundary-layer instability mechanisms which lead to transition. In support of the experimental measurements the mean flow over the flared-cone geometry was modeled by laminar Navier-Stokes computations.; Results show that the boundary layer becomes more stable on the windward ray and less stable on the leeward ray relative to the zero-degree angle-of-attack case, consistent with previous straight-walled cone experiments. The second-mode instability dominates the transition process at a zero-degree angle of attack, however, on the windward ray at an angle of attack this instability was completely stabilized. The less-dominant first-mode instability was slightly destabilized on the windward ray, contrary to previous straight-walled cone experiments. Discrepancies between measured integrated-growth rates and those of previously reported Linear Stability Theory (LST) predictions are observed when the dominant disturbance saturates and harmonics are generated. These non-linear mechanisms, which limit the applicability of LST, are identified from the magnitude of the flow-disturbance bispectra.