Matthew P. Borg

Stability Analysis for HIFiRE Experiments

The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13.5. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis. The transition behavior during this re-entry segment of HIFiRE-1 flight shares some common features with the predicted transition front along the elliptic cone shaped HIFiRE-5 flight article, which was designed to provide hypersonic transition data for a fully 3D geometric configuration. To compare and contrast the crossflow dominated transition over the HIFiRE-1 and HIFiRE-5 configurations, this paper also analyzes boundary layer instabilities over a subscale model of the HIFiRE-5 flight configuration that was tested in the Mach 6 quiet tunnel facility at Purdue University.

Quiet Tunnel Measurements of HIFiRE-5 Boundary-Layer Transition

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratory and Australian Defence Science and Technology Organisation. The principal goal of 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. This facility has been developed to provide quiet flow at high Reynolds number, with low noise levels comparable to flight. This tunnel’s good optical access enabled global measurement of the heat flux by means of temperature-sensitive paint. Two modes of transition were observed: transition along the centerline, suspected to arise from the amplification of second-mode waves in the boundary layer, and transition roughly halfway between the centerline and leading edges, probably due to the bre...

Traveling Crossflow Instability for the HIFiRE-5 Elliptic Cone

A 38.1%-scale model of the Hypersonic International Flight Research Experimentation Program’s Flight Five 2:1 elliptic cone flight vehicle was used to investigate the traveling crossflow instability in a Mach 6 quiet wind tunnel. Traveling crossflow waves were detected with pressure sensors mounted flush with the model surface. The crossflow instability phase speed and wave angle were calculated from the cross spectra of the three pressure sensors. Both quantities showed good agreement with linear stability theory. Duplicate runs at the same initial conditions showed excellent repeatability in traveling crossflow wave properties. Traveling crossflow waves in quiet flow showed very low levels of nonlinear interactions. No traveling crossflow waves were observed for any Reynolds number for elevated freestream noise levels, but transition occurred for a much lower Reynolds number than in quiet flow. Due to the lack of nonlinear growth in quiet flow and the absence of traveling crossflow waves in noisy flow, ...

Starting Issues and Forward-Facing Cavity Resonance in a Hypersonic Quiet Tunnel

Blunt 70° sphere-cones with a 2.5-in. diameter were successfully started in the Boeing/AFOSR Mach-6 Quiet Tunnel when the flow was quiet and the nozzle-wall boundary layer was laminar. However, under noisy flow with thicker turbulent boundary layers, similar models with a 2.0-in. diameter barely started. A porous Apollo capsule with a 2.0-in. diameter at 24° angle of attack failed to start. In an effort to start larger models with stronger bow shocks, new sting-support and diffuser sections were installed. Downstream of the nozzle exit, the diameter was increased from 9.5 to 14.1 in. A series of inserts are being installed behind the backward-facing step in attempts to control the shock/boundarylayer interactions induced by models. These efforts have not yet been successful. During this time, the maximum stagnation pressure for quiet flow has typically ranged from 120 to 145 psia. Pressure measurements near the exit of the Mach-6 diffuser indicated that quiet flow may extend far downstream under some circumstances. A laser differential interferometer was rebuilt for the Mach-6 tunnel with an improved signal-to-noise ratio. Quiet-flow experiments were also carried out at Mach 4 with a pressure transducer at the base of a forward-facing cavity. Computational predictions of self-resonance in deep cavities were confirmed experimentally. For moderately deep cavities, small freestream disturbances were amplified even under quiet flow.

Crossflow Instability for HIFiRE-5 in a Quiet Hypersonic Wind Tunnel

Both the stationary crossflow instability and traveling disturbances were experimentally investigated for the HIFiRE-5 2:1 elliptic cone geometry in Purdue University’s quiet, Mach 6 wind tunnel using surfacemounted pressure transducers, temperature-sensitive paint, and oil-flow visualization. With quiet conditions, stationary crossflow vortices and co-located traveling disturbances were observed. With noisy flow, no traveling disturbances were detected, even at the lowest Reynolds number tested. However, near the model’s leading edges, streamwise streaks were seen in the oil flow, suggesting the presence of stationary crossflow vortices. Discrete roughness elements with various spanwise roughness spacings were utilized in an attempt to change the unperturbed stationary crossflow wavelength. For quiet flow, the unperturbed stationary wavelength was found to be damped out with some roughness spacings, which also excited wavelengths that had been stable in the unperturbed boundary layer. These same roughness elements were also found to damp the measured traveling disturbance with greater effectiveness for decreasing roughness spacing. In noisy flow, some roughness spacings led to detectable streamwise vortices that persisted downstream through what was, in the smooth-wall case, a turbulent boundary layer.

Traveling Crossflow Instability for HIFiRE-5 in a Quiet Hypersonic Wind Tunnel

A scale model of the 2:1 elliptic cone HIFiRE-5 flight vehicle was used to investigate the traveling crossflow instability at Mach 6 in Purdue University’s Mach-6 quiet wind tunnel. Traveling crossflow waves were measured with surface-mounted pressure sensors. The crossflow instability phase speed and wave angle were calculated from the cross spectra of three surface-mounted pressure sensors. Both quantities show good agreement with computational values from about 30-50 kHz. Repeated runs at the same initial condition show excellent repeatability in traveling crossflow wave properties, and give an estimate of the experimental uncertainty associated with this technique. Additionally, autobispectral analysis showed the onset and development of moderate nonlinear quadratic phase-locking prior to transition, but not for the peak traveling crossflow wave. The bicoherence achieved only moderate values. No traveling crossflow waves were observed when freestream noise levels were intentionally elevated, but transition occurred for a much lower Reynolds number. It appears that the traveling crossflow instability is not the primary transition mechanism in the noisy flow of Purdue’s Mach 6 wind tunnel.

Transition Research with Temperature-Sensitive Paints in the Boeing/AFOSR Mach-6 Quiet Tunnel

Abstract : The Boeing/AFOSR Mach-6 Quiet Tunnel (BAM6QT) is used for the study of noise effects on transition. A 12-inch ball valve was installed in the BAM6QT in April 2011, replacing a slow gate valve. This ball valve allows the tunnel to run without the use of diaphragms, though experiments are still typically run with the double burst diaphragm system. Four projects in the BAM6QT are also described in this paper. The first project tested a method of calibrating temperature-sensitive paints using Schmidt-Boelter heat transfer gauges. A 7-deg half-angle cone was tested at 0-deg angle of attack and compared to theory. The second project tested two von Karman ogive models. On the 7.62-cm- diameter ogive model at 0-deg angle of attack, the flow remained laminar for a smooth nosetip, a nosetip with a two-dimensional roughness strip, and a nosetip with distributed roughness. Isolated roughness elements larger than 51 m cause transition on this model at higher Reynolds numbers. A smaller 5.08-cm-diameter ogive model was constructed to allow the model to start at 2-deg angle of attack and to prevent a reflected bow shock from impinging on the model. Transition occurred on the lee ray on the smaller model. Forward-facing and aft-facing steps on the model nosetip did not appear to affect transition. Third, a 3-m circular-arc flared cone was run in different axial positions in the tunnel to determine if there was an effect. Sensors were also installed aft of the model to try to measure noise levels with an installed model in an attempt to show that transition occurs on the cone in fully quiet flow. For the last project, roughness dots were added to the same flared cone in an attempt to change vortex spacing. The flared cone remains a subject for future research.

HIFiRE-5 Attachment-Line and Crossflow Instability in a Quiet Hypersonic Wind Tunnel

Experiments were conducted on the HIFiRE-5 geometry in a quiet Mach-6 wind tunnel with both low and conventional freestream noise levels. The effect of freestream noise on natural transition on the windward surface and roughness-induced transition on the attachment line was investigated. Elevated freestream noise seemed to change the primary transition mechanism on the windward surface. For quiet flow, there appears to be significant stationary crossflow waves. Traveling crossflow waves were also possibly observed coincident with the stationary waves. With noisy flow, neither stationary nor traveling crossflow instabilities were observed. The primary windward transition mechanism with noisy flow remains unknown. Both two-and three-dimensional roughness geometries were tested on the attachment line. As expected, elevated levels of freestream noise served to decrease the critical roughness height for all roughness geometries tested. Three-dimensional roughness elements proved to be the most destabilizing for both noisy and quiet flow. Based on these results, the specified flight roughness tolerance appears to be quite conservative and should prevent early roughness-induced transition.

Simultaneous Infrared And Pressure Measurements Of Crossflow Instability Modes For HIFiRE 5 (POSTPRINT)

A 2:1 elliptic cone model was tested in a Mach-6 quiet wind tunnel with both low and elevated freestream noise levels. Simultaneous measurements were made using an infrared camera and 22 pressure sensors mounted flush with the model surface. Infrared measurements confirmed the presence of stationary crossflow vortices in quiet flow. This is the first time infrared imaging has identified stationary crossflow vortices in a quiet wind tunnel. Power spectral densities of the signals from the pressure transducers also revealed the presence of traveling crossflow waves much farther upstream than had previously been measured. Stationary crossflow wavelengths and traveling crossflow wave properties matched those measured on a previous model of the same geometry. Neither stationary nor traveling crossflow vortices were observed in noisy flow, even though the boundary layer was seen to transition from laminar to turbulent as the freestream Reynolds number was increased. The primary transition mechanism in noisy flow is unknown.

Bypass Transition on the Nozzle Wall of the Boeing/AFOSR Mach-6 Quiet Tunnel

Purdue University continues to develop a 9.5-inch Mach-6 Ludwieg tube for quiet-flow operation to high Reynolds number. Although the facility is operational, and stability measurements are underway, quiet flow has so far been achieved only at low Reynolds number. Bypass transition occurs on the nozzle wall, since the noise rises on the nozzle centerline at a total pressure of 8 psia, for all measured locations, beginning halfway down the nozzle. The bleed-slot flow was plumbed directly to the vacuum tank, eliminating jets that previously existing in the diffuser, but this had no effect on the onset of quiet flow. New measurements of the fluctuations in the diffuser also suggest that noise propagated from downstream is unlikely to be the cause of the bypass. Preliminary measurements of the nonuniformities and fluctuations in the contraction entrance leave open the question of whether these are sufficient to trip the nozzle-wall boundary layer. The wake of a probe in the contraction reduces the Reynolds number of quiet-flow onset but not dramatically. Preliminary measurements of condensation are also reported, along with preliminary hot-wire calibrations and hot-wire measurements of the fluctuations on sharp cones, and the successful fabrication of a new sting support. ∗Associate Professor. Associate Fellow, AIAA. †Research Assistant. Student Member, AIAA. ‡Research Assistant. Student Member, AIAA. §Research Assistant. Student Member, AIAA. ¶Research Assistant. Student Member, AIAA. Copyright c ©2004 by Steven P. Schneider. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Introduction Hypersonic Laminar-Turbulent Transition Laminar-turbulent transition in hypersonic boundary layers is important for prediction and control of heat transfer, skin friction, and other boundary layer properties. However, the mechanisms leading to transition are still poorly understood, even in low-noise environments. Applications hindered by this lack of understanding include reusable launch vehicles [1], high-speed interceptor missiles [2], hypersonic cruise vehicles [3], and reentry vehicles [4]. Many transition experiments have been carried out in conventional ground-testing facilities over the past 50 years. However, these experiments are contaminated by the high levels of noise that radiate from the turbulent boundary layers normally present on the wind tunnel walls [5]. These noise levels, typically 0.5-1% of the mean, are an order of magnitude larger than those observed in flight [6, 7]. These high noise levels can cause transition to occur an order of magnitude earlier than in flight [5, 7]. In addition, the mechanisms of transition operational in small-disturbance environments can be changed or bypassed altogether in high-noise environments; these changes in the mechanisms change the parametric trends in transition [6]. Development of Quiet-Flow Wind Tunnels Only in the last two decades have low-noise supersonic wind tunnels been developed [5, 8]. This development has been difficult, since the test-section wall boundary-layers must be kept laminar in order to avoid high levels of eddy-Mach-wave acoustic radiation from the normally-present turbulent boundary layers. A Mach 3.5 tunnel was the first to be successfully developed at NASA Langley [9]. Langley then developed a Mach 6 quiet nozzle, which