Steven P. Schneider

Laminar to turbulent transition on the HIFiRE-1 cone at Mach 7 and high angle of attack

During the descent phase of the transition flight experiment HIFiRE-1 the angle of attack was higher than expected, since an anomaly occurred in the exoatmospheric pointing maneuver. All pre-flight ground tests were carried out at angles of attack below 6◦. Therefore several post-flight experiments at high angles of attack were performed in the hypersonic wind tunnel (H2K) of the German Aerospace Center in Cologne. The selected Mach number of 7 and the Reynolds number range cover the flow conditions of the flight phase which are relevant for the transition experiment. The test campaign included highfrequency surface pressure measurements with PCB R

Instability Measurements in the Boeing/AFOSR Mach-6 Quiet Tunnel

A new 7 half-angle cone at 6 angle of attack was used to investigate the growth of secondary instabilities on hypersonic stationary crossflow waves. The new cone allows rotation of the sensor array before each run, independently of the rest of the cone. Good repeatability of the cone flow with respect to rotation was established. Measurements of secondary instabilities show the existence of at least two modes, each maximum in different areas of the stationary vortex. Experiments were also performed on a 30 half-angle cone probe equipped with two different pressure transducers to measure the disturbance level of a Mach 6 Ludweig Tube and compare the disturbance levels measured to another facility. It was found that the low frequency disturbance levels differed greatly between the two facilities. Third, RIM insert with 45 roughness elements was tested on a flared cone geometry. Peak second-mode pressure fluctuations of nearly 35% were measured, similar to previous results with a smooth wall. Temperature sensitive paint measurements showed a modified hot-cold-hot heating pattern. Fourth, preliminary measurements on a 7 half angle cone with a slice and flap were made to study free-shear layer transition in a low noise environment. Boundary-layer transition within the free-shear layer has been observed and is altered by the freestream noise levels.

HIFiRE-1 Surface Pressure Fluctuations from High Reynolds, High Angle Ground Test

The HIFiRE-1 is a 7-degree half-angle circular cone with a 2.5-mm nose radius. A successful HIFiRE-1 flight experiment was carried out in March 2010. Due to an anomaly in the exoatmospheric pointing maneuver, the reentry angle of attack was higher than anticipated (5–15 degrees instead of �0). A test campaign in the H2K hypersonic wind tunnel at DLR Cologne was mounted to gather high-frequency pressure fluctuation data and global heat flux via infrared thermography at the high angles of attack and Reynolds numbers encountered in the as-flown trajectory. This paper describes the analysis and interpretation of the surface pressure fluctuations; thermographic analysis is contained in a companion paper. Pressure-fluctuation power spectra were computed at azimuths from the windward to leeward rays in 45° increments at ten sensor locations along the ray, and the expected laminar, transitional, and turbulent regimes were encountered. The disturbances along the windward and leeward rays are presumed to arise from amplified second-mode waves. At 6° angle of attack, 250-kHz disturbances were detected along the 135°-from-windward ray; it is unclear whether these fluctuations arise from second-mode or crossflow instabilities. At 9° angle of attack, similar pressure fluctuations occur along the 135° ray prior to separation near the leeward ray.

Boundary-Layer Transition Measurements in the Boeing/AFOSR Mach-6 Quiet Tunnel

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.

Mach 6 Quiet Tunnel Laminar to Turbulent Investigation of a Generic Hypersonic Forebody

This paper presents an experimental boundary layer transition investigation of the windward side of a generic hypersonic forebody performed in April 2015 in the Boeing/AFOSR Mach 6 quiet tunnel facility at Purdue university (BAM6QT). At 0 and 4 degrees of angle of attack, Reu=11x10 6 /m, flow was fully laminar in quiet conditions. Under noisy conditions, an early transition front (Reθ~200) was observed, even when dividing the unit freestream Reynolds number by 6. In quiet conditions, several diamond-shaped roughness trips were found to efficiently trip the laminar boundary layer when Rekk > 1000. Temperature-Sensitive Paint (TSP) enabled a global measurement of the heat flux distribution and detection of the transition front. PCB sensors confirmed the state of the boundary layer : laminar, turbulent or transitional. Transition results with a single continuously blowing sonic air jet was also collected at various pressure ratios, giving the laminar to turbulent threshold.

Instability and Transition Experiments in the Boeing/AFOSR Mach 6 Quiet Tunnel

This paper presents results for three projects in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. The second-mode instability was measured on a straight 3 half-angle cone under quiet conditions using pressure sensors and temperature sensitive paint. Natural transition was observed near the aft end of the model at a unit Reynolds of 12.1×10/m. Maximum pressure fluctuation magnitudes prior to the onset of transition were measured to be 27% of the mean surface pressure. Secondly, experiments on a 7 half-angle cone at 6 angle of attack were performed to verify the presence of a secondary instability of a stationary crossflow wave. Using TSP imaging techniques and PCB pressure transducers, a possible secondary instability with a peak frequency near 300 kHz was measured at azimuthal angles between 125 and 140 at Reynolds numbers near 10.7×10/m. Lastly, Roughness elements and an air jet were tested as boundary layer trips on a generic hypersonic forebody.

Secondary Instability of Stationary Crossflow Vortices on an Inclined Cone at Mach 6

Using temperature-sensitive paint (TSP) to measure global surface heat transfer on a cone at angle of attack in Mach-6 flow, stationary crossflow waves were observed as hot streaks. When the stationary waves grew large and appeared to begin to break down to turbulence in the vicinity of a pressure sensor, a high-frequency disturbance was detected. This disturbance may be evidence of the secondary instability required to break down the stationary crossflow instability to turbulence. The frequency of the instability appeared to be sensitive to the local boundary layer thickness and small changes in azimuthal position. The azimuthally concentrated signature and the apparent dependence on a threshold stationary wave amplitude indicate that this instability may be caused by the secondary instability of the primary stationary crossflow wave.

Boundary-Layer Transition Experiments in a Hypersonic Quiet Wind Tunnel

This paper reports the progress of three projects in the Boeing/AFOSR Mach-6 quiet tunnel at Purdue University. The first project used a 7-deg half-angle cone at 6-deg angleof-attack with temperature-sensitive paint applied to the frustum and small roughness dots added near the nosetip. Depending on the spacing of the roughness, the spacing and breakdown of the stationary vortices was altered. The second project looked at modifications to a pulsed jet perturber to reduce the perturbation duration. This was accomplished through both electronic and physical modifications to the perturber system. Shorter duration perturbations were achieved, but further progress is required. The third project measured low-frequency disturbances in the first-mode instability frequency range with Kulite pressure transducers on the surface of a cone-ogive-cylinder model. Initial measurements show that the magnitude of the low-frequency disturbance on the cone-ogive-cylinder is greatest outside the boundary layer, which may indicate an entropy-layer instability.

Measurements in the Boeing/AFOSR Mach-6 Quiet Tunnel on Hypersonic Boundary-Layer Transition

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.