Thomas J. Juliano

Transition Research and Improved Performance in the Boeing/AFOSR Mach-6 Quiet Tunnel

Laminar-turbulent transition is critical for vehicles which fly at hypersonic speeds for extended periods, yet conventional wind tunnels suffer from turbulent nozzle-wall boundary layers that produce freestream noise levels 10-100 times higher than flight. The Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) has been developed to provide quiet flow at high Reynolds number, with low noise levels comparable to flight. Laminar nozzle-wall boundary layers and the resulting quiet flow have now been achieved to freestream Reynolds numbers of more than 3.5 × 10/ft., after five years of shakedown. The flawed bleed lip of the original electroformed-nickel throat has been modified to eliminate separation bubbles that are predicted by Rutgers-University computations. The maximum quiet stagnation pressure for the electroformed nozzle has improved to as much as 153 psia. The BAM6QT is now the only operational hypersonic quiet tunnel, anywhere. This quiet flow is here shown to have a marked effect on the development of the stationary crossflow instability on a sharp cone at angle of attack.

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

Inlet Measurements and Quiet Flow Improvements in the Boeing/AFOSR Mach-6 Quiet Tunnel

The Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) has been developed to provide quiet ∞ow at high Reynolds number, with low noise levels comparable to ∞ight. Laminar nozzle-wall boundary layers and the resulting quiet ∞ow have now been achieved to moderately high Reynolds numbers of 2:1 £ 10 6 /ft., after four years of shakedown. The BAM6QT is now the only operational hypersonic quiet tunnel, anywhere. Problems with early transition were apparently due to a small ∞aw in the leading edge of the bleed lip of the original electroformed-nickel throat. The bleed lips of the nickel throat and the new aluminum surrogate throat are now being modifled to eliminate separation bubbles that are still predicted by RutgersUniversity computations. The ∞ow in the inlet to the contraction was also examined in order to determine the level of disturbances present there. A new hot-wire calibration technique was developed that is suitable for the incompressible yet varying density ∞ow present in the contraction. The noise levels were asymmetric but typically on the order of 1%. There is also evidence of a highly asymmetric boundary layer in the contraction. These asymmetries are thought to be caused by pre-run free convection in the contraction.

Measurement of HIFiRE-5 Boundary-Layer Transition in a Mach-6 Quiet Tunnel with Infrared Thermography

The principal goal of the Hypersonic International Flight Research Experimentation (HIFiRE) flight Five is to measure hypersonic boundary-layer transition on a three-dimensional body. The HIFiRE flight tests are supported by a ground test campaign; this paper presents measurements of heat flux and boundary-layer transition in the Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT). This facility has been developed to provide quiet flow at high Reynolds number, with low noise levels comparable to flight. Previously, the global heat flux and location of the transition front were measured with TemperatureSensitive Paint (TSP). Two modes of transition were observed: transition starting in close proximity to the centerline and transition roughly halfway between the centerline and leading edges, probably due to the breakdown of crossflow vortices. One drawback of the TSP was the unavoidable step at the edge of the painted area, which has an unknown effect on the subsequent complex transition front. A new 38.1%-scale HIFiRE-5 model was built with a flush PEEK shell, which makes it suitable for infrared-thermographic heatflux measurements. Quiet-flow tests at Reynolds numbers of 8–12·106 /m and zero angle of attack were conducted and indicate a centerline transition location within 5% of the earlier TSP results, on the order of the uncertainty of the two techniques. This good agreement with independent instrumentation on the flush-walled model reinforces confidence in the earlier measurements.

HIFiRE-1 Boundary-Layer Transition: Ground Test Results and Stability Analysis

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 near zero). A test campaign in the H2K hypersonic wind tunnel at DLR Cologne gathered high-frequency pressure fluctuation data and global heat flux via infrared (IR) thermography at the high angles of attack and Reynolds numbers encountered in the as-flown trajectory. This paper presents analysis of data collected at 0° angle of attack at freestream Reynolds numbers from 5.7 to 10.7e6 /m for 1.6- and 2.5-mm-radius nosetips. The transition onset and end locations derived from IR thermography coincide well with the earliest and largest amplification of pressure fluctuations identified by the fast-response surface-mounted pressure transducers. Stability analysis of the astested conditions was done with the Stability and Transition of Boundary Layers (STABL) nsoftware suite. An N-factor of 5.5 correlates well with transition location for the 1.6-mm radius nosetip. For the blunter nosetip, N ≈ 5.2 at transition. The peak pressure-fluctuation frequencies predicted by STABL agree within 8% of those measured.

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.

High-Reynolds-Number Laminar Flow in the Mach-6 Quiet-Flow Ludwieg Tube

The Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) has been developed to provide laminar nozzle-wall boundary layers at high Reynolds numbers, and thus low noise levels comparable to ∞ight. After four years of shakedown, quiet ∞ow has now been achieved to moderately high unit Reynolds numbers of about 2:7 £ 10 6 /ft., and momentumthickness Reynolds numbers of about 2700, although this performance is not yet reliable. Nevertheless, the BAM6QT is the only operational hypersonic quiet tunnel, anywhere. The freestream pitot-pressure ∞uctuations during high Reynolds number quiet ∞ow are less than 0.02%, the lowest value ever reported. Problems with early transition were apparently due to a small ∞aw in the leading edge of the bleed lip of the original electroformed-nickel throat. The bleed lip of the nickel throat was modifled to eliminate separation bubbles that are still predicted by RutgersUniversity computations. This process should further improve the quiet-∞ow performance of the tunnel. The modiflcation is complete and the resulting bleed-lip shape is reported here, although the performance of the recut bleed-lip tip has not yet been determined. Additional measurements of the temperatures in the contraction entrance are also reported.

Plasma-Actuated Flow Control of Hypersonic Crossflow-Induced Boundary-Layer Transition in a Mach-6 Quiet Tunnel

by Harrison B. Yates The purpose of this research was to control crossflow-induced boundary-layer transition on a cone at angle of attack in hypersonic quiet flow. A 7° half-angle cone model with interchangeable nosetips was designed and fabricated from stainless steel, polyether ether ketone (PEEK), and Macor. Transition was characterized using infrared thermography and Kulite pressure transducers in the Boeing/AFOSR Mach-6 Quiet Tunnel at Purdue University. A plasma-based active flow-control system was used to control the transition location of the stationary crossflow waves, which manifested themselves as hot streaks on the cone. The transition location was accelerated by critical forcing (where the actuator wavenumber equals the wavenumber of naturally largest amplitude waves) and delayed by subcritical forcing (where the actuator wavenumber is larger than the natural waves). The disturbance wavenumber input of the plasma actuators was observed downstream on the model for many of the plasma-on runs, demonstrating that the plasma actuators introduced discrete forcing into the flow. The precise locations of the hot streaks varied for different nosetips, presumably due to differences in their microscale roughness. The experimental data were used to inform an improved stability analysis. Stationary crossflow vortex N-factors were calculated over the surface of a yawed circular