Analytical and experimental studies of shock interference heating in hypersonic flows

An analytical and experimental study is presented of the aerodynamic heating resulting from six types of shock interference patterns encountered in high speed flow. Centerline measurements of pressure and heat transfer distributions on basic bodies were obtained in four wind tunnels for Mach numbers from 6 to 20, specific heat ratios from 1.27 to 1.67, and free stream Reynolds numbers from 3 million to 25.6 million per meter. Peak heating and peak pressures up to 17 and 7.5 times stagnation values, respectively, were measured. In general, results obtained from semiempirical methods developed for each of the six types of interference agreed with the experimental peaks

Evaluation of a technique to generate artificially thickened boundary layers in supersonic and hypersonic flows

The feasibility of using a contoured honeycomb model to generate a thick boundary layer in high-speed, compressible flow was investigated. The contour of the honeycomb was tailored to selectively remove momentum in a minimum of streamwise distance to create an artificially thickened turbulent boundary layer. Three wind tunnel experiments were conducted to verify the concept. Results indicate that this technique is a viable concept, especially for high-speed inlet testing applications. In addition, the compactness of the honeycomb boundary layer simulator allows relatively easy integration into existing wind tunnel model hardware

Use of surface heat transfer measurements as a flow separation diagnostic in a two dimensional reflected oblique shock/turbulent boundary layer interaction

The feasibility of using streamwise surface heat transfer measurements to detect the presence of flow separation in a two-dimensional reflected oblique shock/turbulent boundary layer interaction is reported. Surface heat transfer and static pressure data are presented for attached and separated flows for a free stream nominal Mach number range of 2.5 to 3.5 and shock generator angles of 2 to 8 degrees. The static pressure data do show the characteristic triple inflection point distribution for the strongly separated flow cases. The corresponding surface heat transfer data show unique trends that correlate well with the static pressure determination of the extent of the separated flow region. For the incipient or weakly separated flow cases, the static pressure data do not exhibit the characteristic triple inflection point distribution. However, the same trends in the heat transfer data that are seen for the strongly separated flow cases are evident for the weakly separated flows. Hence, the heat transfer data can be used to determine the extent of weakly separated flows when the surface static pressure distributions often can not

Investigation of the double ramp in hypersonic flow using luminescent measurement systems

Compression ramp flows in supersonic and hypersonic environments present unique flow patterns for shock wave-boundary layer interaction studies. They also represent the generic geometry of two-dimensional inlets and deflected control surfaces for re-entry vehicles. Therefore, a detailed knowledge of the flow behaviour created by such geometries is critical for optimum design. The flow is made more complicated due to the presence of separation regions and streamwise Görtler vortices. The objective of the current research is to study the behaviour and characteristics of the flow over the double ramp model placed in hypersonic flow at freestream Mach number of 5. Three different incidence angles of 0°, −2°, and −4° are studied using colour Schlieren and luminescent paints consisting of anodized aluminium pressure-sensitive paint (AA-PSP) and the temperature-sensitive paint (TSP) technique. The colour Schlieren provides description of the external flow while the global surface pressure and temperature distribution is obtained through the AA-PSP and TSP methods. The TSP technique also proves that it is very effective in identifying the location and properties of the Görtler vortices; revealing the effect of incidence on the magnitude and pattern of Görtler vortices formed

X-ray measurements of growth rates at a gas interface accelerated by shock waves

A new experimental technique to measure the density of a high atomic number gas at a shock-accelerated interface has been developed and demonstrated. It is based on the absorption of x rays by the high atomic number gas, and it was implemented in a vertical square shock tube. The object of the study was the turbulent entrainment and mixing of shock-accelerated air/xenon interfaces prepared by retracting a metal plate, initially separating the two gases, prior to the release of the shock wave. Interfaces of two types, quasi-sinusoidal and nominally flat, were examined. The amplitude of large wavelength (25–100 mm) perturbations on the interface, and the thickness of the interface were measured. An integral definition for the interface mean line was adopted, making it possible to study and time evolution of the individual Fourier modes of the perturbations. A new integral definition for the interface thickness was proposed, making it feasible to study for the first time the time evolution of the thickness of quasi-sinusoidal interfaces. Images of interfaces after interacting with a series of weak waves reverberating between the interface and the shock tube end wall were obtained. The perturbations are studied at the late stages of their evolution, when their amplitude is no longer small compared to their wavelength. Consequently, the measured growth rates of the modal amplitudes are smaller than those predicted by the impulsive model based on the small amplitude approximation. In the case of nominally flat interfaces, the thickness is observed to grow linearly at rates comparable to values previously reported

Growth of shocked gaseous interfaces in a conical geometry

The results of experiments on Richtmyer-Meshkov instability growth of multimode initial perturbations on an air-sulfur hexafluoride (SF6) interface in a conical geometry are presented. The experiments are done in a relatively larger shock tube. A nominally planar interface is formed by sandwiching a polymeric membrane between wire-mesh frames. A single incident shock wave ruptures the membrane resulting in multimode perturbations. The instability develops from the action of baroclinically deposited vorticity at the interface. The visual thickness delta of the interface is measured from schlieren photographs obtained in each run. Data are presented for delta at times when the interface has become turbulent. The data are compared with the experiments of Vetter [Shock Waves 4, 247 (1995)] which were done in a straight test section geometry, to illustrate the effects of area convergence. It is found from schlieren images that the interface thickness grows about 40% to 50% more rapidly than in Vetter's experiments. Laser induced scattering is used to capture the air-helium interface at late times. Image processing of pictures is also used to determine the interface thickness in cases where it was not clear from the pictures and to obtain the dominant eddy-blob sizes in the mixing zone. Some computational studies are also presented to show the global geometry changes of the interface when it implodes into a conical geometry in both light-heavy and heavy-light cases

A semi-quantitative schlieren high-speed flow diagnostic : analysis of high-pressure-ratio, overexpanded planar flow in rocket nozzles

This work introduces a semi-quantitative schlieren (SQS) method which is used to qualitatively and quantitatively analyze complex, unsteady, compressible flows in a small, planar convergent-divergent nozzle. A basic schlieren system is used to image the evolution in time of complex supersonic flow structures, including Prandtl-Meyer expansion fans, internal shocks, near-wall oblique shocks, quasi-normal shocks, shock/boundary layer interactions, shock/shock interactions, and shock trains. The first images of shock trains in high nozzle-pressure-ratio flows are shown, and the underlying processes are described. A flow-field decomposition method is presented which allows the entire flow field to be separated into unit processes and analyzed. Various methods of analysis are presented, including a new method for the determination of node locations along a defined nozzle wall geometry using the method of characteristics. A numerical solution is developed for the analysis of a blow-down process. Computer programs which implement these solutions are presented

An inventory of aeronautical ground research facilities. Volume 1: Wind tunnels

A survey of wind tunnel research facilities in the United States is presented. The inventory includes all subsonic, transonic, and hypersonic wind tunnels operated by governmental and private organizations. Each wind tunnel is described with respect to size, mechanical operation, construction, testing capabilities, and operating costs. Facility performance data are presented in charts and tables

An investigation of passive control methods for shock-induced separation at hypersonic speeds

Peer Reviewedhttp://deepblue.lib.umich.edu/bitstream/2027.42/76765/1/AIAA-1992-2725-533.pd