Search

Experiments were performed within Rutgers University’s Mach 3.4 wind tunnel to visualize the unsteady flow over a wall-mounted hemisphere using high-speed schlieren imaging and stereoscopic particl...

The influence of a finite-width shock generator on the incident shock in the impinging shock-wave/boundary-layer interaction is investigated using different widths of the shock generator. The shock...

A wavelet-based optical flow analysis (wOFA) method for processing background-oriented schlieren (BOS) images is presented and demonstrated on synthetic and experimental data. Optical flow is inher...

A novel configuration of focused laser differential interferometry (FLDI) was developed, simulated, and tested. The new configuration is referred to as cylindrical FLDI or CFLDI. Using cylindrical ...

This paper describes the development and use of background-oriented schlieren imaging of a full-scale supersonic jet in flight. A series of flight tests was performed in April 2011, October 2014, F...

The three-dimensional (3-D) shock-wave/boundary-layer interaction (SBLI) between a fin-generated shock and a turbulent boundary layer is examined using plenoptic particle image velocimetry (PIV): a...

An experimental study has been conducted to investigate unsteady flow phenomena observed within various two-dimensional configurations of shock/boundary layer interactions. Six configurations have been tested in Mach 2 flow: ϕ1 = 14◦ and 20◦ compression ramps, and incident shock reflections from ϕ1 = 7◦ , 8 ◦ , 9 ◦ , and 10◦ shock generators; Reynolds numbers in each case are Reθ ≈ 8350. The flow is assessed using an array of fast-response pressure transducers in conjunction with a high-repetition rate PIV system. Development of the mean flow structures early in each interaction is observed to be consistent with the Free Interaction concept. Unsteady wall-pressure energy content at frequencies above those associated with the characteristic low-frequency shock motion also show significant similarities in the vicinity of the shock foot. Results confirm that this low-frequency peak is not associated with a narrow-band forcing mechanism from either upstream or downstream, but rather a characteristic frequency that varies with interaction strength, which describes the flow’s dynamic response. These findings support various models published in literature that have sought to explain the source of low-frequency unsteady shock motion.

Extensive wind-tunnel tests were performed on a Y-shaped diverterless supersonic inlet (DSI). All tests were conducted at a free stream Mach number of M∞=1.65, the design Mach number for this inlet...

An image-based flow velocimetry technique coupling pulse-burst laser technology and the cross-correlation Doppler global velocimetry (CC-DGV) technique is presented as a faster variant of the CC-DG...

A control method is developed to suppress pressure oscillations induced by supersonic cavity flow using high-speed upstream injection. The injection is generated with a large blowing coefficient th...

The shock wave boundary layer interactions (SWBLIs) at compression ramps (ramp angle α=20o -30o ) are studied at Ma=2.0 and under two Reynolds numbers (Re1=18,600 and Re2=35,600, Re based on boundary layer thickness). High-speed schlieren operating at 20 kHz is used as the flow diagnostics. The flow structures in the compression ramp SWBLIs, including the shock wave, interaction region and induced turbulent region over the ramp surface, are discussed. Their variation under increasing ramp angle and Reynolds number are further examined. The low-frequency shock wave oscillations are also studied through tracking the shock wave motion. A larger ramp angle increases the spectral intensity of the shock wave’s low-frequency unsteadiness, while increasing the Reynolds number results in a lower peak frequency for the separation and reattachment shock waves.

This study experimentally investigated the effects of repetitive laser energy deposition using a supersonic intake model with a central conical compression surface and a 19×19 mm^2 cross-sectional duct in a Mach 1.92 indraft wind tunnel: especially in the subcritical and buzz modes. A single-pulse energy deposition was observed to suppress the flow separation at the compression surface by “sweeping” a shock wave system with the thermal bubble generated by the energy deposition. The duration of the sweeping effect was approximately 160 μs in the subcritical mode and 100–420 μs in the buzz mode. Furthermore, repetitive deposition of laser pulse energies was observed to moderate instabilities in both modes, and it increased the pressure recovery by as much as 8%; also, the occurrence of buzz was delayed, thereby widening the stable, subcritical regime. In both modes, there was a threshold value for the laser pulse repetition frequency fL, thd, which corresponded to the duration of the sweeping effect (e.g., fthd=6 kHz in subcritical mode). Below this frequency, the increase in the pressure was proportional to the repetition frequency; whereas above fthd, the effect per single pulse was reduced., Published Online:26 Oct 2019

Experiments on supersonic flows are typically conducted in wind tunnels with rectangular cross-sections, which use two-dimensional nozzles of two different types. A “full” setup consists of two contoured nozzle surfaces symmetric about the tunnel centre-height. The “half” configuration is also common, with a curved ceiling nozzle surface and a straight horizontal floor.

This paper presents a combined experimental and computational study of a low-sweep transonic natural laminar flow (NLF) wing with shock-control bumps (SCBs). A transonic NLF wing with a relatively low sweep angle of 20 deg was chosen for this study. To avoid the complexity of the flow introduced by perforated/slotted walls commonly used for transonic wind-tunnel tests for reducing the wall interference, both experimental tests and computational simulations were conducted with solid wind-tunnel wall conditions. This allows for like-to-like validation of the computational simulation. Optimization of the shock-control bumps was first conducted to design the wind-tunnel test model with bumps. Two critical parameters of the three-dimensional SCBs for shock control (i.e., bump crest position and bump height) were optimized in terms of total drag reduction at the given design point in the wind tunnel. We show that the strong shock wave on the low-sweep NLF wing can be effective controlled by well-designed SCBs deployed along the wing span. The optimized SCBs result in 18.5% pressure drag reduction with 5% viscous drag penalty, and the SCBs also bring some benefits at off-design conditions. The wind-tunnel tests include pressure measurement, particle image velocimetry, and temperature-sensitive paint to provide detailed insight into the shock-control flowfield and to validate the computational simulations. Comparisons include surface pressure profile, velocity distribution, and transition location.

To greatly improve the supersonic inlet stability at low cost of structural weight and complexity, a novel buzz suppression strategy based on fixed-geometry air bleed is developed. It is designed t...

The base flow of an axisymmetric generic space launcher model is investigated experimentally by means of particle image velocimetry and dynamic pressure measurements at a Mach number of 0.76 and a Reynolds number of 1.5 × 106, based on the main body diameter. The flow separation at the end of the main body forms a highly dynamic recirculation region with strong pressure fluctuations on the reattaching surface. The time-averaged reattachment on the rear sting is at 1.05 main body diameters downstream of the step. This work investigates the application of passive flow control devices for their potential of reducing the loads on the space launcher’s nozzle. It is shown that rectangular or circular grooves at the end of the main body force enhanced mixing in the separated shear layer, leading to a reduction of the reattachment length of 55%. Additionally, the fluctuations of the reattachment are significantly reduced, which results in lower-pressure fluctuations and thus reduced dynamic loads.

A deformable shape-memory alloy (SMA) bump is introduced to control the succeeding cowl-shock/boundary-layer interactions in a supersonic inlet with an operating Mach range of 2.0–3.8. The deformat...

The mechanism of oblique shock wave control by surface arc discharge plasma is explored through a combined numerical and experimental study. The experiments are conducted in a supersonic wind tunne...

Shock-wave/boundary-layer interactions occurring in the inlets cause degradation of inlet efficiency. To better control these interactions and improve inlet performance, shock-wave/boundary-layer i...

Aerodynamic interference can occur between high-speed slender bodies when in close proximity. A complex flowfield develops where shock and expansion waves from a generator body impinge upon the adjacent receiver body and modify its aerodynamic characteristics in comparison to the isolated case. The aim of this research is to quantify and understand the multibody interference effects that arise between a finned slender body and a second disturbance generator body. A parametric wind tunnel study was performed in which the effects of the receiver incidence and axial stagger were considered. Computational fluid dynamic simulations showed good agreement with the measurements, and these were used in the interpretation of the experimental results. The overall interference loads for a given multibody configuration were found to be a complex function of the pressure footprints from the compression and expansion waves emanating from the generator body as well as the flow pitch induced by the generator shockwave. These induced interference loads change sign as the shock impingement location moves aft over the receiver and in some cases cause the receiver body to become statically unstable. Overall, the observed interference effects can modify the subsequent body trajectories and may increase the likelihood of a collision.

A theory based upon linearized governing equations is developed that describes the operation principles and design parameters for low-speed wind tunnels with longitudinal freestream oscillations. Existing measurements made in unsteady wind tunnels are shown to be consistent with the theory and targeted validation experiments performed in a variable-geometry blowdown-type wind tunnel revealed excellent correspondence with the theoretical results. In particular, the tunnel frequency bandwidth is proportional to the mean tunnel freestream velocity and inversely proportional to the test-section length and the square of the exit area to test-section area ratio. The acoustics equations reveal a “Helmholtz damping ratio” that is not only dependent on the tunnel geometry and exit area but also proportional to the freestream Mach number. At appreciable reduced frequencies, dynamic stall on the louver vanes increases pressure losses, thereby reducing the mean flow speeds. Varying the exit area results in louver-van...

The application of subsonic and supersonic vortex generators for flow control in an external compression, axisymmetric low-boom concept inlet was investigated experimentally and computationally. The low-boom inlet design is conceptually based on a previous scale model tested in the NASA 1×1 ft supersonic wind tunnel. It features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. A design of experiments computational methodology was used to select device size and location to determine the optimal choice of device geometry and placement, with the objective of reducing radial flow distortion for on-design conditions. The resulting test matrix of vortex generator geometry and placement was specified for model inlet testing conducted at the 8×6 ft supersonic wind tunnel at NASA Glenn Research Center in the Fall of 2010. Comparisons of Reynolds-averaged Navier–Stokes predictions with experimental data were subsequen...

A computational-fluid-dynamics-based waverider design approach is discussed, which can allow arbitrary generating flowfields (for which no analytical solution exists) to be used in waverider design. To validate this design technique, a waverider with a lift-to-drag ratio L/D of 2.62 was designed and optimized using a Mach 3.5 conical flowfield for a series of wind-tunnel experiments. A payload volume constraint was implemented such that an aerodynamically optimized design could be generated that met wind-tunnel installation and model manufacturability requirements. A variety of experiments at both on- and off-design conditions were performed using a blowdown supersonic wind tunnel. Measurements were obtained using a sting balance, pressure-sensitive paint (with an a priori temperature compensation approach), and schlieren and oil-flow visualization. For comparison with the experimental results, computational-fluid-dynamics simulations of the configuration were also performed; the design L/D and aerodynami...

Spectra of wall-pressure fluctuations caused by separation shock unsteadiness were compared for data obtained from wind-tunnel experiments, the Hypersonic International Flight Research Experimentation flight test 1, and large-eddy simulations. The results were found to be in generally good agreement, despite differences in Mach number and two orders of magnitude difference in Reynolds number. Relatively good agreement was obtained between these spectra and the predictions of a theory developed by Plotkin. The predictions of this theory are also qualitatively consistent with the results of experiments in which the shock motion was synchronized to controlled perturbations. The results presented here support the idea that separation unsteadiness has common features across a broad range of compressible flows and that it behaves as a selective amplifier of large-scale disturbances in the incoming flow.

Previous studies of transonic shock control bumps have often been either numerical or experimental. Comparisons between the two have been hampered by the limitations of either approach. The present work aims to bridge the gap between computational fluid dynamics and experiment by planning a joint approach from the outset. This enables high-quality validation data to be produced and ensures that the conclusions of either aspect of the study are directly relevant to the application. Experiments conducted with bumps mounted on the floor of a blowdown tunnel were modified to include an additional postshock adverse pressure gradient through the use of a diffuser as well as introducing boundary-layer suction ahead of the test section to enable the in-flow boundary layer to be manipulated. This has the advantage of being an inexpensive and highly repeatable method. Computations were performed on a standard airfoil model, with the flight conditions as free parameters. The experimental and computational setups wer...

It is proposed that the adverse effects from secondary diaphragm rupture in an expansion tunnel may be reduced or eliminated by orienting the tunnel vertically, matching the test gas pressure and the accelerator gas pressure, and initially separating the test gas from the accelerator gas by density stratification. This proposed configuration is termed the vertical expansion tunnel. Two benefits are 1) the removal of the diaphragm particulates in the test gas after its rupture, and 2) the elimination of the wave system that is a result of a real secondary diaphragm having a finite mass and thickness. An inviscid perfect-gas analysis and quasi-one-dimensional Euler computations are performed to find the available effective reservoir conditions (pressure and mass specific enthalpy) and useful test time in a vertical expansion tunnel for comparison to a conventional expansion tunnel and a reflected-shock tunnel. The maximum effective reservoir conditions of the vertical expansion tunnel are higher than the reflected-shock tunnel but lower than the expansion tunnel. The useful test time in the vertical expansion tunnel is slightly longer than the expansion tunnel but shorter than the reflected-shock tunnel. If some sacrifice of the effective reservoir conditions can be made, the vertical expansion tunnel could be used in hypervelocity ground testing without the problems associated with secondary diaphragm rupture.

To investigate whether vortex generators can be an effective form of passive flow control an experimental investigation has been conducted in a small-scale wind tunnel. With specific emphasis on supersonic inlet applications flow separation was initiated using a combined terminal shock wave and subsonic diffuser: a configuration that has been developed as a part of a program to produce a more inlet-relevant flowfield in a small-scale wind tunnel than previous studies. When flow control was initially introduced little overall flow improvement was obtained as the losses tended to be redistributed instead of removed. It became apparent that there existed a strong coupling between the center-span flow and the corner flows. As a consequence, only when flow control was applied to both the corner flows and center-span flow was a significant flow improvement obtained. When corner suction and center-span vortex generators were employed in tandem separation was much reduced and wall-pressure and stagnation pressure...

A small-scale Mach 5 blow down wind tunnel, with ample access for optical diagnostics and ability to generate steady-state nonequilibrium flows, has been designed and operated. The wind tunnel uses transverse repetitively pulsed nanosecond discharge, fully overlapped with a transverse DC discharge and operated at high plenum pressures (P0=0.5-1.0 atm) to load internal energy modes of nitrogen and oxygen molecules. The discharge remains stable at energy loadings of up to ~0.1 eV/molecule in nitrogen (discharge power up to 2.5 kW). The wind tunnel generates nonequilibrium nitrogen and air flows with steady-state run time of 5-10 seconds, translational / rotational temperature of T0~300-400 K, and estimated upper bound nitrogen vibrational temperature of Tv0~2,000 K. Internal energy mode disequilibrium in the flow is controlled by injecting nitric oxide, hydrogen, or water vapor into the subsonic flow between the discharge section and the nozzle throat. The effect of energy mode disequilibrium is studied in a flow over a cylinder model placed in the Mach 5 test section. The flow field in the supersonic test section is well predicted by a 3-D compressible Navier-Stokes flow code, indicating good flow quality. The supersonic flow field over the model is visualized by schlieren imaging and NO PLIF imaging, using a burst mode laser operated in the vicinity of 226 nm, at a pulse repetition rate of 10-20 kHz. The laser was operated in the injection-seeded mode, generating narrow linewidth (~0.1 cm -1 ) output for single-line NO excitation in the flow. Nitric oxide was either injected into the flow in the plenum or generated in a repetitively pulsed nanosecond discharge in dry air. Both single-pulse PLIF images and images integrated over 10-50 laser pulses have been obtained. Two single-line NO PLIF images on a NO(X,v˝=0→A,v´=0) transition are used for measurements of 2-D temperature distributions in nitrogen flows in the supersonic test section. Another single-line NO PLIF image on a NO(X,v˝=1→A,v´=1) transition is used to estimate NO vibrational temperature behind the bow shock, TV(NO)=550 ± 100 K. The NO vibrational temperature increases when the energy loading in the discharge is increased. Kinetic modeling calculations indicate that low NO vibrational temperature is due to fairly low vibrational energy loading per nitrogen molecule in the discharge. Schlieren images of a supersonic flow over the cylinder model demonstrate that the shock stand-off distance is reduced by approximately 5% when the discharge in the wind tunnel is in operation and water vapor or hydrogen are injected into the flow between the discharge section and the nozzle throat. This effect is attributed to additional heating of the flow in the plenum during relaxation of vibrationally excited nitrogen in the presence of water vapor or hydrogen.

The disruption of simulated fuel droplets in supersonic flow is examined experimentally in a drawdown supersonic wind tunnel. The droplets are accelerated in the supersonic flow, achieving supersonic velocities relative to the surrounding air with relative Mach numbers as high as 1.8 and Weber numbers as high as 300. Monodisperse 100m-diam fluid droplets are generated using a droplet-on-demand generator upstream of the tunnel entrance. The droplets are imaged by direct close-up singleandmultiple-exposure imaging. Three test liquidswere employed: 2-propanol and tetraethylene glycol dimethyl ether as nonvolatile fluids, and a more volatile 50=50 hexanol-pentane mixture. The decreased static pressure in the supersonic stream had the potential to give rise to superheating of the droplet fluid, as in some cases, the static pressure became significantly lower than the vapor pressure of the droplet liquid. Droplet lifetimes for the hexanol/pentane mixture appear to be shorter due to accelerated vaporization consistent with superheating, although little impact is observed on the droplet velocity and relative Mach number. Droplet-disruption patterns for these supersonic flow conditions can be classified into four different flow regions by considering the changes in the Weber number with downstream distance as the droplets accelerate. The drag coefficients associatedwith the droplet disruption under locally supersonic conditions are generally higher than those expected for solid spheres, largely due to the cross-sectional area change associated with droplet deformation/ breakup.

DOI: 10.2514/1.J051202 This paper presents a numerical simulation analysis of flutter-boundary prediction of a two-dimensional morphing wing, which starts to flutter in the process of structural adaptation, although the flight speed is kept constant.Thestructuraladaptationisexpressedrepresentativelybyacontinuouschangeinthenaturalfrequencyof the heaving or the pitching motion of the wing, which induces the closeness and coalescence of the two frequencies of the aeroelastic modes. The dynamic characteristics of the wing are evaluated by applying a digital estimation techniquetoasequenceofdigitalizednonstationaryresponsesofthewingexcitedbystationaryairturbulencein flow during the ongoing adaptation. The critical boundary of the wing is predicted by the flutter margin for digitalized systems based on Jury’s stability criterion, which is mathematically equivalent to Routh-Hurwitz’s criterion. Numericalresultsshowthatthispredictionapproachbasedonthe fluttermarginfordiscrete-timesystems,whichthe authorsproposedoriginallyforordinarynonmorphingwings,isalsoveryeffectiveinpredictingthe flutterboundary of the morphing wing in the process of adaptation.

Jets in a supersonic crossflow are known to produce a three-dimensional bow-shock structure due to the blockage of the flow. In the present study, streamwise linear arrays of high-momentum microjets are used to generate either single or multiple oblique shocks in a supersonic crossflow. The shocks generated using microjets can be tailored in terms of their strength and be made either parallel or coalescing, depending on the application. Flow visualization using shadowgraph and density field measurements using background-oriented schlieren (BOS) technique were carried out for a range of microjet operating conditions. The results obtained using the two methods are consistent and complementary and show a linear variation of oblique shock angles with a microjet pressure ratio over the range of conditions tested. The density field obtained using BOS clearly shows the oblique shocks generated using these microjet arrays, the jump in density across the shock, the extent of the high-density field, the expansion fan, and the associated decrease in density. The results suggest that microjet arrays can be successfully used to develop techniques for sonic boom mitigation and high-performance supersonic inlets.

The fluorescence ratio technique for processing planar laser-induced fluorescence data was generalized for quantitative imaging of the injectant mole fraction and extended to quantify the density distributions in a nonreacting supersonic mixing flowfield. The original fluorescence ratio approach was first developed by Hartfield et al. (Hartfield, R. J., Jr., Abbitt, J. D., III, and McDaniel, J. C., "Injectant Mole Fraction Imaging in Compressible Mixing Flow Using Planar Laser-Induced Iodine Fluorescence," Optics Letters, Vol. 1, No. 16, Aug. 1989, pp. 850-852.) for tests in a special closed-loop wind tunnel to eliminate the effects of thermodynamic property variations on planar laser-induced fluorescence signals in compressible flowfields. This approach provided us a quantitative means of planar mole-fraction measurement; however, it implicitly assumed that the tracer molecules were seeded at the same fraction in both the main and the secondary flows. In the present study, we generalized the Hartfield et al. method by considering differences in the tracer-seeding rates for obtaining planar images of mole fraction and density. Experimental validation of the new method was carried out in a mixing flowfield formed by sonic transverse injection into a Mach 1.9 supersonic airstream. The injectant mole-fraction distribution obtained from planar laser-induced fluorescence data processed by our new approach showed better agreement with the gas-sampling data than one based on the Hartfield et al. method. The density distribution was verified by comparison with the theoretical density ratio across the oblique shock wave.

In the present study, the effects of large-scale periodic surface roughness on a high-speed (M = 2.86), high Reynolds number (Re θ ≈ 60,000), supersonic turbulent boundary layer was examined. Two roughness topologies (square and diamond) were compared with an aerodynamically smooth wall. The measurements included planar contours of the mean and fluctuating velocity, pitot pressure profiles, pressure-sensitive paint, and schlieren photography. The local strain-rate distortion parameters for the square roughness pattern were small (∼-0.01), and the mean and turbulent flow properties followed the canonical rough-wall boundary-layer trends. The diamond-shaped roughness topology produced a pattern of attached oblique shocks and expansion waves that led to strong distortion parameters. The distortions varied from -0.3 to 0.4 across the roughness elements, which resulted in localized extra turbulence production that generated large periodic variations in the turbulence levels across individual roughness elements that spanned the boundary-layer thickness; for example, the Reynolds shear stress varied by ∼100%. This result demonstrated a mechanism for altering the turbulence in supersonic boundary layers.

Supersonic inlet buzz in a rectangular, mixed-compression inlet has been simulated on a 20 x 10 6 points mesh using the delayed detached-eddy simulation method, a version of detached-eddy simulation that ensures the attached boundary layers are treated using Reynolds-averaged Navier-Stokes equations. The results are compared with experimental data obtained during a previous campaign of wind-tunnel experiments. The comparison of unsteady data is performed thanks to phase averages, Fourier transforms, and wavelet transforms. The buzz observed at Mach 1.8, which occurred at a frequency of 18 Hz, is well reproduced. The shock oscillations, as well as the different flow features experimentally observed, are present in the simulation. The buzz frequency, as well as higher frequencies existing in the experimental pressure signals, are correctly predicted. The data issued from the simulation (time history of pressure fluctuations, pseudo-Schlieren, and three-dimensional visualizations) allow a better investigation of the inlet flowfield during buzz and a detailed description and physical analysis of this phenomenon. A description and an explanation of the mechanism at the origin of secondary oscillations that occur at a higher frequency during buzz are proposed. The crucial role of acoustic waves moving through the duct is shown.

This paper aims at investigating the effect of discharges generated near steady bow shocks and their possible use for localized heat deposition leading to drag reduction. We report the observation of steady discharges generated in front of a blunted model in a M = 3 supersonic flow. The test model is designed as a pair of coaxial electrodes, and set up in the ONERA R1Ch M = 3 blowdown wind tunnel. A high voltage power supply is used to generate negative discharges. A corona regime, a glow regime, and a filamentary arc regime are observed. The negative corona regime consists of Trichel pulses dissipating less than 200 mW of electrical power. The glow discharge absorbs powers up to 0.5 kW. It displays a strong light emission in the vicinity of the shock front, followed by a darker region downstream of the shock. The drag coefficient is measured and shows no measurable change when a glow discharge is switched on. To explain this and further investigate the effect of localized heat deposition at the shock front, modified Rankine-Hugoniot jump relations are computed, taking into account a volumetric heat source. This allows one to compute a fair estimate of the drag coefficient and shows that drag reduction by localized heating in the shock front is possible. However, it also shows that in our experiment, the plasma thermal power is too small to appreciably reduce the drag, possibly because of the role of the electron impact excitation of N 2 vibrations, whose fairly long relaxation time could shift downstream the effective gas heating. More generally, the model shows that power-efficient plasma-induced drag reduction requires high plasma heating efficiency.

An experimental study of supersonic inlet buzz is presented. This study was carried out on a mixed compression rectangular inlet model; tests were done at Mach numbers ranging from 1.8 to 3, with and without bleed. Inlet flows were analyzed thanks to Schlieren videos and signal processing of unsteady pressure recordings. Two kinds of buzz were observed: "little buzz," which is thought to correspond to an acoustic resonance phenomenon excited by the presence of a shear layer under the cowl lip (Ferri criterion), and "big buzz," which seems to be triggered by a boundary layer separation on the compression ramps (Dailey criterion). In some cases, little buzz could be virtually suppressed by the introduction of a bleed. The frequency of big buzz is shown to be already present in the flow before the onset of large-amplitude oscillations, which suggests that the underlying mechanism of big buzz, probably linked to acoustics, already exists before buzz starts.

L AMINAR-turbulent transition in boundary layers is important for the prediction and control of skin friction, heat transfer, and other boundary layer properties. Therefore, it is important to have reliable capabilities in predicting boundary layer transition to optimize hypersonic vehicle performance [1]. The mechanisms leading to transition are poorly understood [2]. Transition experiments have been carried out in conventional ground testing facilities for many decades. However, most of the experimental data obtained from these facilities are not reliable because they have much higher disturbance levels compared with actual flight conditions [3]. Quiet-flow wind tunnels are intended to replicate the low noise conditions of actual flight at hypersonic speed. Reaching quiet flow requires the maintenance of a laminar boundary layer on the nozzle wall to avoid acoustic fluctuations generated by boundary layer turbulence. One method to reduce noise is to delay boundary layer transition using a bleed slot before the nozzle throat. The Boeing/Air Force Office of Scientific Research (AFOSR)Mach 6 wind tunnel at Purdue University has been designed as a quiet tunnel with a bleed slot for which the noise level is an order of magnitude lower than that in conventional wind tunnels. It is a Ludwieg tube that is a long pipe having a converging-diverging nozzle followed by a test section as shown in Fig. 1. A close-up view of the tunnel geometry around the bleed slot lip is shown in Fig. 2. However, the tunnel (which has been operational since 2001) is not yet quiet for the desired range of stagnation pressures of up to 150 psi. Two different nozzles have been fabricated and tested. The tunnel is quiet up to a stagnation pressure of 8 psi with the original electroformed nozzle. The original design of the outer surface of the bleed slot has been modified, and eight different bleed slot designs together with the original one have been tested [4]. A second nozzle throat has been fabricated from aluminum [5]. The tests on the tunnel with this aluminum surrogate throat show that the tunnel is quiet up to a stagnation pressure of 93 psi. Early transition of the nozzle wall boundary layer has been identified as the cause of the test section noise for the tunnel at Purdue University. Separation bubbles on the bleed lip and associated fluctuations induced near the bleed lip were identified as the most likely cause of early transition [4]. The experimental study of Klebanoff and Tidstrom [6] showed that the presence of a separation bubble of sufficient size destabilizes the laminar boundary layer downstream of reattachment thereby leading to an earlier transition to turbulence, i.e., the location of transition moves upstream relative to where it would occur without the separation bubble. This hypothesis regarding a separation bubble was supported by the measurements showing an increase in quiet-flow stagnation pressure from 8 to 93 psia when the electroformed nozzle throat was replaced with the aluminum throat [5]. The bleed lip of the electroformed throat has a 0.001 in. kink that is not present in the aluminum throat, and it appears that the kink in the electroformed throat exacerbates a natural tendency to form a separation bubble near the lip. This separation bubble is highly unsteady and can lead to early transition downstream [7]. However, the separation bubble apparently still exists even at 93 psia, according to the computations presented herein. To achieve quiet flow above 93 psia, and to make the quiet flow less sensitive to the exact shape of the bleed lip, it is desirable to eliminate the separation bubble completely. The situation in the hypersonicwind tunnel at PurdueUniversity is an example illustrating the importance of the bleed lip geometry and the effects of separation bubbles that form around the bleed lip on the quality of the flow at the test section. The objective of this study is to demonstrate the effect of separation bubbles on flow structure by numerically investigating the existence of steady and unsteady separation bubbles on the main-flow or the bleed-flow side of the nozzle lip of the Boeing/AFOSR Mach 6 wind tunnel at Purdue University, and to design a new geometry to eliminate or reduce the size of the separation bubbles.

This investigation details the design, analysis, and testing of a new, two-component wall shear gauge for three-dimensional, high-temperature flows. This gauge is a direct-measuring, nonnulling design with a round head surrounded by a small gap. Two flexure rings are used to allow small motions of the floating head. Strain gauges are mounted on the flexures, and fiber-optic displacement sensors measure how far polished faces of counterweights move in relation to a fixed housing. The strain gauges are for validation of the newer fiber optics. The sensor is constructed of Haynes® 230®, a high-temperature nickel alloy. All components, in pure fiber-optic form, can survive to a temperature of 1073 K. The dynamic range of the sensor is from 0-500 Pa. Higher shear forces can be measured by changing the floating head size. No damping or water cooling of the sensor is required. Finite element modeling was used during the design and analysis of the sensor. Static structural, modal, and thermal analyses were performed using the ANSYS finite element package. Repeated cold-flow tests at Mach 2.4 and Mach 4.0 under high-Reynolds-number conditions have been accomplished in the Virginia Tech Supersonic Wind Tunnel. Experimental results are in excellent agreement with semiempirical prediction methods.

THE flowfield due to transverse injection of a round sonic jet into a supersonic flowis a configuration of interest in the design of supersonic combustors or thrust vector control of supersonic jets. The flow is also of fundamental interest because it presents separation from a smooth surface, embedded subsonic regions, curved shear layers, strong shocks, an unusual development of the injected jet into a kidney-shaped streamwise vortex pair, and a wake behind the jet. Although the geometry is simple, the flow is complex and is a good candidate for assessing the behavior of turbulence models for high-speed flow, beginning with the corresponding two-dimensional flow shown in Fig. 1. At the slot, an underexpanded sonic jet expands rapidly into the supersonic crossflow. Expansion waves reflect at the jet boundary, coalesce, and give rise to a Mach surface (Mach disk for round jets).

A new blowdown nonequilibrium plasma magnetohydrodynamic (MHD) supersonic wind tunnel operated at complete steady state has been developed and tested at Ohio State. The wind tunnel can be operated at Mach numbers up to M = 3-4 and mass flow rates of up to 45 g/s at a stagnation pressure of 1 atm

The effects of nonaxisymmetrical fuselage design for reducing the drag of a low-sonic-boom airplane are investigated by computational fluid dynamics (CFD) analyses and verified in wind-tunnel tests. The nonaxisymmetrical fuselage design, in which the upper side of a fuselage is designed for low drag whereas the lower side is designed for low sonic boom, is applied to the design of a Mach 1.7 scaled supersonic experimental airplane. The designed airplane is compared with a low-drag airplane and a low-sonic-boom airplane with an axisymmetrical fuselage