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This paper presents an experimental boundary layer transition investigation of the windward side of a generic hypersonic forebody performed in April 2015 in theBoeing/AFOSR Mach 6 quiet tunnel facility at Purdue university (BAM6QT). At 0 and 4 degrees of angle of attack, =11x10 6 /m, flow was fully laminar in quiet conditions. Under noisy conditions, an early transition front (Re θ~2 00) 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 > . 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. NomenclatureCFD = Computational Fluid Dynamics BAM6QT = Boeing/AFOSR Mach 6 Quiet Tunnel NT = Natural Transition BLT = Boundary Layer Transition WT = Wind-Tunnel JISC = Jet In Supersonic Crossflow PR = Pressure Ratio LT = Laminar to Turbulent transition TSP = Thermal Sensitive Painting Re u = Unit Reynolds number, /m St = Stanton number δ = Laminar boundary layer thickness, mm μ = Dynamic viscosity, m2/s ρ = Density, kg/m3 u = Streamwise velocity, m/s M e = Edge boundary layer mach number P e = Edge boundary layer pressure T w = Wall temperature, K 2 k = Roughness element height, mm h = Mach disk height Re hh = Mach disk height Reynolds number Re kk = Roughness element height Reynolds number Re θ = Momentum thickness Reynolds number P i0 = Stagnation pressure, Pa T i0 = Stagnation temperature, K P ∞ = Freestream pressure, Pa T ∞ = Freestream temperature, K J = Jet momentum flux ratio

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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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