Shann Rufer

Measurements and Computations of Second-Mode Instability Waves in Three Hypersonic Wind Tunnels

High-frequency pressure-fluctuation measurements were made in AEDC Tunnel 9 at Mach 10 and the NASA Langley 15-Inch Mach 6 and 31-Inch Mach 10 tunnels. Measurements were made on a 7{sup o}-half-angle cone model. Pitot measurements of freestream pressure fluctuations were also made in Tunnel 9 and the Langley Mach-6 tunnel. For the first time, second-mode waves were measured in all of these tunnels, using 1-MHz-response pressure sensors. In Tunnel 9, second-mode waves could be seen in power spectra computed from records as short as 80 {micro}s. The second-mode wave amplitudes were observed to saturate and then begin to decrease in the Langley tunnels, indicating wave breakdown. Breakdown was estimated to occur near N {approx} 5 in the Langley Mach-10 tunnel. The unit-Reynolds-number variations in the data from Tunnel 9 were too large to see the same processes.

Hypersonic Transition Research in the Boeing/AFOSR Mach-6 Quiet Tunnel

A Mach-6 Ludwieg tube is being developed for high Reynolds number quiet-flow operation. The modelsupport centerbody had been causing upstream separation when the nozzle-wall boundary layers became laminar at low pressures. The centerbody has now been removed, resulting in attached Mach 5.7 quiet flow below 8 psia total pressure. Pitot measurements show that low-noise flow begins at the same 8 psia, from halfway down the nozzle to near the nozzle exit. Thus, transition in the nozzle-wall boundary layer is apparently bypassing the usual linear instability processes. Measurements of the static pressure on the diffuser walls show large fluctuation when the nozzle-wall boundary layer is laminar, probably due to upstream propagation of bleed-slot jet noise from the diffuser. Finally, initial oil-flow images show the development of crossflow vortices on a sharp cone at angle of attack, and hot-wire measurements show initial evidence of instability waves on a sharp cone at zero angle of attack.

Progress in the Operation 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. The design, fabrication, and initial operation were reported earlier. The present paper reports progress in achieving and characterizing the quiet Mach-6 flow, and in developing instrumentation. A new design for the bleed-slot throat geometry enabled achieving some initial quiet flow, although only at very low Reynolds numbers of about 200,000. A thicker hot wire of 0.0002inch diameter is successfully surviving many tunnel runs. Preliminary measurements were obtained on the Hyper2000 generic scramjet forebody using temperature-sensitive paint. These show the development of streamwise vortices from the leading edge imperfections. These vortices become much more evident following the first compression corner, and can be generated in a controlled fashion using small roughness strips on the leading edge.

Experimental Study of Second-Mode Instabilities on a 7-Degree Cone at Mach 6

Experiments have been carried out in the NASA Langley Research Center 20-Inch Mach 6 Air Tunnel to measure the second-mode boundary-layer instability on a 7° half-angle cone using high-frequency pressure sensors. Data were obtained with both blunt and sharp nosetips installed on the cone. The second-mode wave amplitudes were observed to saturate and then begin to decrease in the Langley tunnels, indicating wave breakdown. Pressure fluctuation measurements and thermocouple data indicated the location of transition along the cone at the different conditions tested. Comparisons between the power density spectra obtained during the current test and previous data from the Langley 15-Inch Mach 6 High Temperature Tunnel and the Boeing/AFOSR Mach 6 Quiet tunnel illustrate the effect of tunnel noise on instability growth and transition.

Shakedown of the Purdue Mach-6 Quiet-Flow Ludwieg Tube

Purdue University continues to develop a 9.5-inch Mach-6 Ludwieg tube for quiet-flow operation to high Reynolds number. The aerodynamic and mechanical design were reported earlier, along with the design and testing of several facility subsystems. Measurements of the coordinates of the completed upstream nozzle sections are here reported. The development and further testing of other facility systems are also described: these include the second round of burst-diaphragm tests, the bleed-slot suction system, and the driver-tube and contraction heating systems. Delivery of the quiet-flow nozzle is presently scheduled for February 2001. Progress in developing repeatable hot-wire calibrations is also reported.

Pressure Fluctuation Measurements in the NASA Langley 20-Inch Mach 6 Wind Tunnel

Experiments have been carried out in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to measure the freestream pressure fluctuations, or tunnel noise, using a pitot rake. These experiments are part of an on-going effort to characterize the freestream disturbances of the Langley hypersonic wind tunnels along with other facilities around the country. Once the freestream disturbances have been characterized, a better understanding of the effect of these disturbances on boundary layer instability and transition measurements can be gained. The current experiments use a multi-probe pitot rake instrumented with both Kulite and PCB pressure transducers. Data were obtained over a range of Reynolds numbers and test section axial and radial positions. In general, noise levels were consistent spatially across the test section and ranged from 1% at the highest Reynolds numbers tested to approximately 1.6% at the lowest Reynolds number tested.

Hypersonic Stability and Transition Experiments on Blunt Cones and a Generic Scramjet Forebody

Purdue University continues to develop a 9.5-inch Mach-6 Ludwieg tube, which presently operates with quiet flow only at low Reynolds number. Efforts towards achieving high quiet Reynolds numbers are reported. Measurements of stability and transition are also being carried out, using the existing conventional-noise Mach-6 flow. Model geometries include blunt round cones at zero and non-zero angle of attack, and the Hyper2000 forebody, which is generic for the Hyper-X class of airbreathing cruise vehicles. The transition literature for these cases is reviewed. Stationary streamwise-vortex instabilities are induced on the Hyper2000 using small roughness elements. Their growth is measured with temperature-sensitive paints. Temperature-paints measurements on a sharp cone at angle of attack show preliminary indications of the stationary crossflow vortices. Finally, a preliminary hot-wire profile was obtained on a blunt cone in a single tunnel run using a newly automated traverse. Hypersonic Transition and Quiet Tunnels 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 lownoise environments. Applications hindered by this ∗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 ©2003 by Steven P. Schneider. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. lack of understanding include reusable launch vehicles [1], high-speed interceptor missiles [2], hypersonic cruise vehicles [3], and ballistic 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]. Only in the last two decades have low-noise supersonic wind tunnels been developed [5, 8]. This development has been difficult, since the test-sectionwall 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 [10]. Unfortunately, this nozzle was removed from service due to a space conflict. No hypersonic quiet tunnels are presently operational anywhere in the world. The general prediction of transition based on simulations of the transition mechanisms is a very complex and difficult problem. There are several known receptivity mechanisms, several different known forms of instability waves, many different parameters that affect the mean flow and therefore modify the stability properties, and many known nonlinear breakdown mechanisms. The parameter space is large. The scramjet-vehicle forebody and

HOT-WIRE MEASUREMENTS OF INSTABILITY WAVES ON A BLUNT CONE AT MACH-6

Hot wire calibrations were completed in the supersonic jet facility at Purdue University. This jet consists of a Mach-4 nozzle and allows for a continuous run time. The hot wires were calibrated under conditions similar to those seen in the Mach-6 facility at Purdue University. These hot wires were then used in the Mach-6 tunnel to study the amplitude and growth of instability waves on 7-deg half-angle sharp and blunt cones.

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

Hot-Wire Measurements of Instabilty Waves on Cones at Mach 6

Hot wires were used in the Mach-6 tunnel to obtain mass flux profiles and to study the amplitude and growth of instability waves on 7° half-angle sharp and blunt (0.020-inch radius) cones. Measurements were taken at several stagnation pressures (45, 70, 90, and 125 psia) on each of the cones. The boundary layer thickness at convention noise was similar for the sharp and blunt cones, though it was noticed that for the same pressure and axial location the boundary layer on the sharp cone was slightly thinner than that on the blunt cone. The measurements taken with the hot wires were then compared to computations using the STABL code. These comparisons show good agreement both in the shape of the mass flux profiles and the value of mass flux at the edge of the boundary layer. Good agreement was also seen in the frequency of the second-mode instabilities at various pressures; the frequencies found experimentally were within 8% of those calculated by STABL for all cases. Hot wire measurements were also obtained on a sharp cone under quiet-flow conditions and though no instabilities were observed at 0° angle of attack, when placed at a 3° angle of attack, peaks in the spectra were observed at 130 kHz which could be attributed to instabilities.