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.

Recent results from several projects in the BAM6QT are presented. An infrared camera system was used to image a circular cone at an angle of attack, and the results are compared to previous TSP measurements. The IR images show clear streaks and demonstrate repeatability and low noise levels compared to TSP. Oil flow and surface pressure sensor measurements are presented for a cone with a slice and ramp. Separation and reattachment are discussed, along with the amplification and dampening of instabilities at various locations on the model. The temperature distribution along the BAM6QT nozzle wall was varied to study the relationship between heating and the percentage of a run which was quiet. No apparent correlation was observed. Pitot-probe measurements were taken at various locations on the nozzle centerline to investigate an increase in noise levels that occurs roughly two seconds into runs. The magnitude of the increase and the time at which it started depended on the Reynolds number. Development of higher-Reynolds number hypersonic quiet tunnel facilities may require the use of suction on the nozzle wall. Initial computations are presented for the design of a flared inlet centerbody that can be tested in the Boeing AFOSR/Mach-6 Quiet tunnel to determine the feasibility of creating sufficiently uniform suction. A stability analysis is performed to determine the most unstable second-mode frequencies and to compute the Görtler numbers on the flared aft-body portion. Finally, the 3 inch shock tube used for PCB calibration has been upgraded with high accuracy sensors and an automated pressure control system.

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While 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 article 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. New Direct Numerical Simulation (DNS) datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the

This paper presents results from four different research projects currently ongoing at Purdue University. (1) Different criteria for detecting the edge of the boundary layer were investigated on the flared cone geometry. It was determined that a method based on the total enthalpy profile would be used for future edge-detection computations on the flared cone geometry. A Rod Insertion Method (RIM) roughness insert was measured using a Zygo ZeGage optical profiler. Experimental results with a single RIM insert are presented. Maximum second-mode magnitudes of nearly 27% were measured 2.5 cm upstream from where spectral filling and intermittency algorithms compute that transition has begun. (2) Preliminary data from a new model shows that high-frequency secondary instabilities of stationary crossflow waves are localized under the troughs of the stationary vortices. Measurements of the growth of the secondary instabilities are reported for two different vortices. (3) In order to better understand the effects that probe geometry has on measured pressure fluctuations, pitot measurements were taken using various sleeves which alter the forward-facing diameter of the probe. The results indicate a clear effect of probe size on the measured power spectral densities. Furthermore, it was found that the geometry effects are Reynolds number dependent. (4) Experiments on a cone with a slice and ramp were completed to determine if transitional shock wave-boundary layers interactions can be measured within the Boeing/AFOSR Mach-6 Quiet Tunnel. Initial experiments showed that with a newly designed model it is possible to measure transitional interactions.