The response of the mean and turbulent flow structure of a supersonic high-Reynolds-number turbulent boundary layer flow subjected to local and global mechanical distortions was experimentally examined. Local disturbances were introduced via small-scale wall patterns, and global distortions were induced through streamline curvature-driven pressure gradients. Local surface topologies included k-type diamond and d-type square elements; a smooth wall was examined for comparison purposes. Three global distortions were studied with each of the three surface topologies. Measurements included planar contours of the mean and fluctuating velocity via particle image velocimetry, Pitot pressure profiles, pressure sensitive paint and Schlieren photography. The velocity data were acquired with sufficient resolution to characterize the mean and turbulent flow structure and to examine interactions between the local surface roughness distortions and the imposed pressure gradients on the turbulence production. A strong response to both the local and global distortions was observed with the diamond elements, where the effect of the elements extended into the outer regions of the boundary layer. It was shown that the primary cause for the observed response was the result of local shock and expansion waves modifying the turbulence structure and production. By contrast, the square elements showed a less pronounced response to local flow distortions as the waves were significantly weaker. However, the frictional losses were higher for the blunter square roughness elements. Detailed quantitative characterizations of the turbulence flow structure and the associated production mechanisms are described herein. These experiments demonstrate fundamental differences between supersonic and subsonic rough-wall flows, and the new understanding of the underlying mechanisms provides a scientific basis to systematically modify the mean and turbulence flow structure all the way across supersonic boundary layers.

In the present study, the effects of large-scale periodic surface roughness on a high-speed (M = 2.86), high Reynolds number (Reo approximate to 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 (similar to - 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 similar to 100 %. This result demonstrated a mechanism for altering the turbulence in supersonic boundary layers.

Oblique wing aircraft hold the promise of combining efficient supersonic and subsonic flight with excellent low speed endurance. For this reason, there has been recent renewed interest in developing oblique wing aircraft. This paper will provide a historical review of oblique wing demonstrator aircraft and other major oblique wing research. A significant amount of early theoretical work was done by R.T. Jones of NASA Ames, beginning in the 1950s. In the past 35 years, a number of small test aircraft were flown to prove the feasibility of oblique wing aircraft flight control. Small gliders, several small remotely piloted aircraft and the manned AD-1 aircraft, were flown by NASA. In the 1990s, Steve Morris of Stanford University built and flew two oblique flying wing aircraft – the first powered oblique flying wings to fly. Numerous conceptual design studies and research papers have addressed oblique wing-body-tail or oblique flying wing designs. A number of wind tunnel tests have been performed on both oblique wing and oblique flying wing designs. This paper provides an overview of the research, testing and flight demonstrations related to oblique wing and oblique flying wing aircraft over the past half-century.

θ ) supersonic turbulent boundary layers were examined. Large localized distortions ( d = -0.5 to 0.4) were generated from the shock and expansion structure produced by diamond roughness elements. Weak ( dmax = 0.05) and strong global favorable pressure gradients ( dmax = 0.25) were studied. The results were compared to similar flows with canonical surface patterns (smooth and square roughness). The measurements included planar contours of the mean and fluctuating velocity, Pitot pressure profiles, pressure sensitive paint and schlieren photography. The canonical flows followed established trends. However, their inclusion provides (1) a basis for comparison for the non- canonical flows and (2) new high-speed experimental data with turbulence. The diamond roughness element produced substantially different flows that were characterized by strong local distortions ( d = -0.5 to 1.8 across the element) and highly varying turbulence properties, where the shear stress levels varied by ~100%. The present data showed that combining the global pressure gradient to the local gradients associated with the diamond roughness element produced regions of flow over a rough surface with turbulence levels reduced to 70% of the undisturbed zero pressure boundary layer. These data and trends have important implications in controlling the turbulence in high-speed boundary layers.

Introduction S CHLIEREN imaging systems are widely used for both qualitative and quantitative flow visualization in compressible flows and active index of refraction flowfields such as two-index mixing problems or combustion flowfields. A common implementation of schlieren imaging uses a pulsed light source to provide instantaneous measurements in unsteady flowfields. A variety of pulsed light sources has been used in the past with schlieren imaging systems, including arc lamps, incandescent bulbs, flash tubes, spark gaps, and light-emitting diodes.1 Lasers have been used to provide a narrow linewidth illumination source, which is useful for filtering broad spectral emissions from plasmas or flames, but at the cost of image degradation due to laser speckle. A recent technical note described the use of a laser-induced spark as a point source, which could be inserted in the flowfield avoiding the need to integrate through the density fluctuations associated with the boundary layers on the wind-tunnel walls.2 Recent papers have also described the use of a laser-induced spark as a light source for schlieren imaging in a plasma flow3 and an exploding wire bridge,4 both applications that benefit from a very high-intensity schlieren light source. Our objective in this Technical Note is to characterize a laserinduced spark schlieren imaging technique that provides a very highintensity light source, with short time duration and with repeatable temporal and spatial characteristics. Spatial and temporal variations in intensity are reported for this light source, as well as a comparable light source using a laser discharge in air. This light source has

A Planar Doppler Velocimetry system has been developed for multi-component velocity measurements in large scale subsonic wind tunnel facilities. System components, methodologies and improvements are evaluated and discussed. Data is presented on two component measurements conducted in the flow field above a 70 degree delta wing at an angle of attack of 23 degrees in a Mach 0.2 (69 m/s) free stream. Although only two components of velocity could be resolved, axial and spanwise measurements were made characterizing the velocity field associated with the vortex cores as they developed downstream. Detector placement and system performance was improved in a second test to measure the three dimensional velocity field above a Boeing UCAV model again at 20 degrees angle of attack operated in a Mach 0.2 free stream. The evolution of the three dimensional velocity field created by vortices from the sharp leading edge of the body and the outboard vortex above the wing was characterized. Further optimization of the current system indicated that current levels of uncertainty can be improved by proper camera placement to reduce the sensitivity to laser frequency fluctuations and by improving the linear independence of the measured velocity components as characterized by the condition number of the coefficient matrix. Other considerations of camera placement and considerations to make measurements are also discussed.

Hypersonic vehicles are subject to intense local aerothermodynamic loading associated with specific local flow phenomena. An example is the well-known Edney IV interaction which can result in peak heat transfer rates which are orders of magnitude above the already high stagnation point heat transfer. Failure to alleviate or control such local aerothermodynamic loading can leading to serious degradation of vehicle aerodynamic integrity and performance, and even to catastrophic failure. In this paper, we explore the capability of pulsed laser energy deposition to favorably modify the shock structure associated with an Edney IV interaction generated by the intersection of an oblique shock with the blunt body shock associated with a sphere at Mach 3.45. Surface pressure measurements indicate a signicant modification of the surface pressure during the interaction including substantial reduction in peak surface pressure.

An experimental investigation was conducted to examine the effect of a pulsed Nd:YAG laser energy addition on the shock structures and surface pressure in a Mach 3.45 flow past a sphere. Two configurations were considered: 1) a sphere in a uniform freestream and 2) an Edney IV interaction generated by impingement of an oblique shock on the bow shock of the sphere

A combined computational and experimental study was performed to investigate the effect of a single laser energy pulse on the transition from a Mach Reflection (MR) to a Regular Reflection (RR) in the Dual Solution Domain (DSD). The freestream Mach number is 3.45 and two oblique shock waves are formed by two symmetric \(22^\circ\) wedges. These conditions correspond to a point midway within the DSD wherein either an MR or an RR is possible. A steady MR was first obtained experimentally and numerically, then a single laser pulse was deposited above the horizontal center plane. In the experiment, the laser beam was focused resulting in a deposition volume of approximately 3 mm3, while in the simulation, the laser pulse was modeled as an initial variation of the temperature and pressure using Gaussian profile. A grid refinement study was conducted to assess the accuracy of the numerical simulations. For the steady MR, the simulation showed the variation of Mach stem height along the span due to side effects. The predicted spanwise averaged Mach stem height was 1.96 mm within 2% of the experimental value of 2 mm. The experiment showed that the Mach stem height decreased to 30% of its original height due to the interaction with the thermal spot generated by the laser pulse and then returned to its original height by \(300\;\mu\)s. That the Mach stem returned to its original height was most likely due to freestream turbulence in the wind tunnel. The numerical simulation successfully predicted the reverse transition from a stable MR to a stable RR and the stable RR persisted across the span. This study showed the capability of a laser energy pulse to control the reverse transition of MR \(\rightarrow\) RR within the Dual Solution Domain.

A planar Doppler velocimetry (PDV) system was developed and demonstrated in a small-scale facility (Mach 1.36 freejet) and then applied in a large-scale subsonic wind tunnel, where measurements were made over a delta wing at a 23-deg angle of attack. This PDV system utilized a pulsed, injection-seeded, frequency-doubled Nd:YAG laser to interrogate the flow. Back-illuminated charge-coupled device (CCD) cameras in conjunction with an iodine filter were used to record images produced by the scattered laser light, permitting the determination of the velocity at each CCD pixel, The PDV instrument also included custom software and a frequency-monitoring system composed of photodiodes, gated integrators, and a second iodine cell. With this setup, we recorded the shot-to-shot iodine-filtered and reference images and the associated laser frequency. In the freejet, mean velocities in the core were measured by PDV to within 6.4 m/s (out of similar to 260 m/s) of the value obtained by laser Doppler velocimetry. In the wind tunnel, freestream empty-tunnel measurements indicated bias and random errors of less than 2 and 4 m/s, respectively. The dominant source of random error arose from laser speckle, and the dominant source of bias error came from the characterization of the iodine filters. Measurements over the delta wing showed similar velocity ranges but smaller vortex cores when compared to the velocity field predicted by a computational fluid dynamics model.

Molecular filter based diagnostics are continuing to gain popularity as a research tool for investigations in areas of aerodynamics, fluid mechanics, and combustion. This class of diagnostics has gone by many terms including Filtered Rayleigh Scattering, Doppler Global Velocimetry, and Planar Doppler Velocimetry. The majority of this article reviews recent advances in Planar Doppler Velocimetry in measuring up to three velocity components over a planar region in a flowfield. The history of the development of these techniques is given with a description of typical systems, components, and levels of uncertainty in the measurement. Current trends indicate that uncertainties on the order of 1m/s are possible with these techniques. A comprehensive review is also given on the application of Planar Doppler Velocimetry to laboratory flows, supersonic flows, and large scale subsonic wind tunnels. The article concludes with a description of future trends, which may simplify the technique, followed by a description of techniques which allow multi-property measurements (i.e. velocity, density, temperature, and pressure) simultaneously.

Doppler Global Velocimetry (DGV) is a new laser-based diagnostic which complements existing measurement capabilities for wind tunnel applications. Measurements of velocity are based on determining the Doppler shift of single frequency laser light scattered off particles in the flow field. This technique is capable of making simultaneous measurements over an entire laser sheet. This paper presents results obtained with the DGV instrument in both calibration and wind tunnel test setups. Characterization and analysis of isystem accuracy is addressed with the calibration setups. Example applications of the DGV technique for measurements on a wind tunnel model are also presented.

Velocity measurements using the Planar Doppler Velocimetry (PDV) technique were made in a small jet facility and in a large wind tunnel to study the feasibility of the technique including sources of error and means to minimize them. PDV is a non-intrusive technique which allows the measurement of one or more velocity components everywhere in a plane within a flow field. Its potential to yield instantaneous, simultaneous velocities throughout a plane and at high spatial resolution in a single measurement makes it a promising tool for turbulence studies. PDV measurements are made by detecting the Doppler shift produced when laser light is scattered from moving particles in a flow field. The current PDV system utilizes a pulsed, injection-seeded, frequency-doubled Nd:YAG laser capable of producing narrow linewidth light (-140 MHz) that can be tuned to frequencies associated with the absorption lines of iodine. An iodine filter may therefore be used to discriminate the frequency of the laser light. A system utilizing two scientific-grade CCD cameras was used to record images produced by the scattered laser light. An unusual feature of the experimental arrangement was a frequency monitoring system which tracked the laser set point frequency on an image to image basis, providing resolution to better than 4 MHz. In a Mach 1.36 free jet, mean flow velocities in the core (averaged from three streamwise stations) were measured by PDV to within 2.4% of the value obtained by LDV (to within 6.4 m/s). For the high speed core of the jet, instantaneous velocity fluctuations greater than 3.4% (about 9 m/s) were measurable by PDV, but lesser fluctuations were masked by PDV system noise. In the wind tunnel (3.05 x 2.13 m test section), uniform flow runs at 68 and 96 m/s indicated bias and random errors of less than 2 and 4 m/sec, respectively. Mean measurements of the flow over a delta wing at a 23° angle-of-attack were similar to the results of a CFD calculation by Rizzetta (1996). The dominant source of random error in the velocity measurements arose from the phenomenon of laser speckle and the dominant source of bias error came from the characterization of the iodine filter profiles.