The Boeing/AFOSR Mach-6 Quiet Tunnel achieved quiet flow to a stagnation pressure of 163 psia in Dec. 2008, the highest value observed so far. It remains quiet at pressures above 160 psia. Under noisy conditions, nozzle-wall boundary-layer separation and the associated tunnel shutdown appear to propagate slowly upstream, whereas under quiet conditions, the propagation is very rapid. A new diffuser insert has been designed, fabricated, and installed in the tunnel in order to start larger blunt models and increase run time. A flared cone with a circular-arc geometry was designed to generate large second-mode N factors under quiet flow conditions. When the computed N factor was 13, large instability waves were measured under quiet flow conditions using fast pressure sensors, but the flow remained laminar. Transition was observed only under noisy conditions. A laminar instability was detected in the wake of an isolated roughness element in the boundary layer on the nozzle wall; this appears to be the first such measurement at hypersonic speeds.

The Boeing/AFOSR Mach-6 Quiet Tunnel has achieved quiet flow to stagnation pressures of 146 psia, and intermittently quiet flow between 146 and 169 psia. In an attempt to measure natural transition under quiet flow, a 3-m-circular-arc compression cone was tested with a nearly sharp nosetip. Using temperature-sensitive paint, hot streaks were observed to develop near the rear of the cone at high pressures under quiet flow. The streaks do not appear under noisy flow. The cause of the hot streaks remains unknown, though they may be instabilities or artifacts of nonlinear breakdown. Under quiet flow, the cone boundary layer remained laminar up to N factors of at least 15 and possibly as high as 19. Transition occurred at N = 9 under noisy flow. It is unknown why laminar flow persisted to such high N factors. As part of an investigation of crossflow vortices, a 7 • half-angle cone was tested at 6 • angle of attack with temperature-sensitive paint finishes of varying roughness. The roughness of the paint finish was observed to have an effect on crossflow vortices, in some cases inducing transition under noisy flow. Heat-transfer measurements were made at the stagnation point of a hemisphere to observe the effect of freestream noise; no effect was evident.

A unique new removable anechoic system and new acoustic treatment for the Virginia Tech Stability Wind Tunnel is described. The new system consists of a Kevlar-walled acoustic test section flanked by two anechoic chambers. In its new configuration the facility is closed aerodynamically and open acoustically, allowing far-field acoustic measurements with a flow quality comparable to that of a hardwalled wind tunnel. An extensive program of experiments has been conducted to examine the performance of this new hardware under a range of conditions, both to examine the effects of acoustic treatment on overall test-section noise levels and to ascertain the aerodynamic characteristics of the new test section. Noise levels in the test section of the anechoic facility are down by as much as 25 dB compared to the original hard-walled configuration. Lift interference corrections (for a baseline NACA 0012 airfoil) are less than half those expected in an open-jet wind tunnel. Acoustic measurements of airfoil trailing edge noise using a microphone phased array are compared to past experiments conducted on similar airfoils in an open-jet facility.

Although low-disturbance ("quiet") hypersonic wind tunnels are believed to provide more reliable extrapolation of boundary-layer transition behavior from ground to flight, the presently available quiet facilities are limited to Mach 6, moderate Reynolds numbers, low freestream enthalpy, and subscale models. As a result, only conventional ("noisy") wind tunnels can reproduce both Reynolds numbers and enthalpies of hypersonic flight configurations and must therefore be used for flight vehicle test and evaluation involving high Mach number, high enthalpy, and larger models. This paper outlines the recent progress and achievements in the characterization of tunnel noise that have resulted from the coordinated effort within the AVT-240 specialists group on hypersonic boundary-layer transition prediction. The new experimental measurements cover a range of conventional wind tunnels with different sizes and Mach numbers from 6 to 14 and extend the database of freestream fluctuations within the spectral range of boundary-layer instability waves over commonly tested models. New direct numerical simulation datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the spatiotemporal structure of the freestream noise, and account for the propagation and transfer of the freestream disturbances to a Pitot-mounted sensor.

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