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Rearward-facing steps in laminar supersonic flows with and without suction.
It was found in the experimental study conducted that, for both suction and no-suction cases, an increase in the step height caused a sharp drop in the initial heating rates which then gradually recovered to less or near the attached-flow values. The height of the step controlled the heating rates at the step base. Mass suction from the separated area increased the local heating rates, the relative increase being most significant immediately behind the step. In general, however, the effect of mass suction on heat-transfer at all laminar flows was relatively weak.
A study of methods which predict supersonic flow fields from body geometry, distance, and Mach number
A study of seven methods for predicting flow-field pressure signatures from the parameters Mach number, body geometry, and field-path distance has been made. The methods included the method of characteristics, which served as a standard of comparison; a shock-capturing method; three Whitham theory methods; a modified characteristics method; and a bicharacteristics method. Results from each method were also compared with recently obtained wind-tunnel data for a cone-cylinder model at Mach numbers of 2.96 and 4.63 with ratios of radial distance to cone length of 2 and 5. The comparisons at a Mach number of 2.96 showed that signatures from all the methods correlated well with wind-tunnel data and with the signatures predicted by the method of characteristics. At a Mach number of 4.63, however, the agreement between the signatures obtained in the wind tunnel and those predicted by theory varied from good to poor, as did the agreement between the signatures obtained by the method of characteristics and the other six methods. It should be noted that these results and comparisons indicate pressure prediction capabilities only for the near-field flow about bodies of revolution.
Fluid dynamic studies on scattering aerosol and its generation for application as tracer particles in supersonic flow measurements utilizing laser Doppler velocimeter
An experimental study on the particle-fluid interactions of scattering aerosols was performed using monodisperse aerosols of different particle sizes for the application of laser Doppler velocimeters in subsonic turbulence measurements. Particle response was measured by subjecting the particles to an acoustically excited oscillatory fluid velocity field and by comparing the ratio of particle velocity amplitude to the fluid velocity amplitude as a function of particle size and the frequency of oscillation. Particle velocity was measured by using a differential laser Doppler velocimeter. The test aerosols were fairly monodisperse with a mean diameter that could be controlled over the size range from 0.1 to 1.0 micron. Experimental results on the generation of a fairly monodisperse aerosol of solid particles and liquid droplets and on the aerosol response in the frequency range 100 Hz to 100 kHz are presented. It is indicated that a unit density spherical scatterer of 0.3 micron-diameter would be an optimum choice as tracer particles for subsonic air turbulence measurements.
Rearward-facing steps in laminar supersonic flows with and without suction
An experimental investigation of heat-transfer and pressure distributions within regions of laminar separated flows produced by two-dimensional rearward-facing steps has been carried out at freestream Mach numbers of around 4 in the range of step height-to-boundary layer thickness varying from 0.1 to 2.4. With no suction from the separated area, the ratio of the maximum post-step heat transfer to the attached-flow values was less than unity. The maximum heating-rate region was located far downstream of the reattachment plate stagnation point. Mass suction from the separated area increased the local heating rates, this effect was however relatively weak for purely laminar flow conditions and the competing effect of the step height clearly predominated. At step heights comparable with boundary-layer thickness, even removing the entire approaching boundary layer was not sufficient to raise the post-step heating rates above the flat-plate values.
Effect of interjet spacing on mixing of multiple coaxial jets in supersonic flow
An experimental and analytical investigation of strut-mounted, four-nozzle, downstream facing, gaseous fuel injector assembly was conducted in a Mach 4 air-stream with P = 154 psia and T sub o 520 R. Helium was used as the injectant, and the interjet spacing (S/D) was the main parameter varied. The principal data are in the form of Helium concentration profiles at seven axial stations, Mach number distributions at three axial stations and schlieren photographs of the flow field at different interjet spacings. An approximate analysis was developed based upon a linearization in the Von Mises plane, Crocco integrals for the temperature and concentrations fields and an eddy viscosity model. Good agreement with the data was achieved. It was found that, at these conditions, interjet spacing has an effect on mixing only up to S/D of approximately 2.5, with concentration of injectant becoming smaller as S/D is increased, and the Mach number downstream of the injector tends to decrease with decreasing S/D.
Normal injection of helium from swept struts into ducted supersonic flow
Recent design studies have shown that airframe-integrated scramjets should include instream mounted, swept-back strut fuel injectors to obtain short combustors. Because there was no data in the literature on mixing characteristics of swept strut fuel injectors, the present investigation was undertaken to provide such data. This investigation was made with two swept struts in a closed duct at Mach number of 4.4 and nominal jet-to-air mass flow ratio of 0.029 with helium used to simulate hydrogen fuel. The data is compared with flat plate mounted normal injector data to obtain the effect of swept struts on mixing. Three injector patterns were evaluated representing the range of hole spacing and jet-to-freestream dynamic pressure ratio of interest. Measured helium concentration, pitot pressure, and static pressure in the downstream mixing region are used to generate contour plots necessary to define the mixing region flow field and the mixing parameters.
Skin friction reduction in supersonic flow by injection through slots, porous sections and combinations of the two
An experimental study of skin friction reduction in a Mach 3.0 air steam with gaseous injection through a tangential slot, a porous wall section, and combinations of the two was conducted. The primary data obtained were wall shear values measured directly with a floating element balance and also inferred from Preston Tube measurements. Detailed profiles at several axial stations, wall pressure distributions and schlieren photographs are presented. The data indicate that a slot provides the greatest skin friction reduction in comparison with a reference flat plate experiment. The porous wall section arrangement suffers from an apparent roughness-induced rise in skin friction at low injection rates compared to the flat plate. The combination schemes demonstrated a potential for gain.
Effects of geometric variables on the performance of a probe for direct measurement of free-steam stagnation pressure in supersonic flow
An investigation was conducted at Mach numbers of 1.41, 1.83, and 2.20 to determine the effects of parametric variations in the height of the pitot-tube center line from the probe surface and in the radius of the surface curvature on the pressure recovery of a probe designed to measure the free-stream stagnation pressure. The probe consists of a pitot tube mounted on the surface of a curved cylinder of circular cross section. The pitot tube senses the pressure of the stream tube which has been slowed by isentropic compression along the curved surface. Pressure recovery, greater than or equal to 99.8 percent of the free-stream stagnation pressure, was obtained for a wide range both of angle of attack and of yaw for probes satisfying the predetermined optimum design criteria.
A short static-pressure probe design for supersonic flow
A static-pressure probe design concept was developed which has the static holes located close to the probe tip and is relatively insensitive to probe angle of attack and circumferential static hole location. Probes were constructed with 10 and 20 deg half-angle cone tips followed by a tangent conic curve section and a tangent cone section of 2, 3, or 3.5 deg, and were tested at Mach numbers of 2.5 and 4.0 and angles of attack up to 12 deg. Experimental results indicate that for stream Mach numbers of 2.5 and 4.0 and probe angle of attack within + or - 10 deg, values of stream static pressure can be determined from probe calibration to within about + or - 4 percent. If the probe is aligned within about 7 deg of the flow experimental results indicated, the stream static pressures can be determined to within 2 percent from probe calibration.
Experimental investigation of supersonic flow past double-wedge configurations
Viscous-inviscid interactions characteristic of those which occur when the fuselage-generated shock wave interacts with the wing-generated shock wave of a shuttle orbiter were studied experimentally. Surface-pressure measurements and schlieren photographs were obtained to define the flowfield generated when a Mach 4.97 stream encounters a double-wedge configuration. The deflection angles for the two wedge surfaces were such that the shock interaction pattern was either a Type-V pattern or a Type-VI pattern, as defined by Edney. The correlation between the present data and the theoretical solution for the Type-VI solution is satisfactory. The correlation between the measured Type-V shock-interaction pattern and the theoretical solution is satisfactory up to the interaction region. Downstream of the interaction the Type-V data depend primarily on the shape of the leading-edge shock wave.
An integrated approach to the analysis and design of wings and wing-body combinations in supersonic flow
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Flutter of laminated plates in supersonic flow
A solution procedure was developed using linear small deflection theory for the flutter of simply supported laminated plates. For such plates, the bending and extensional governing equations are coupled and have cross-stiffness terms which do not appear in classical plate theory. An extended Galerkin method is used to obtain approximate solutions to the governing equations, and the aerodynamic pressure loading used in the analysis is that given by linear piston theory with flow at arbitrary cross-flow angle. A limited parametric study was conducted for typical laminated composite plates. The calculations show that both the bending-extensional coupling and the cross-stiffness terms have a large destabilizing effect on flutter. Since classical plate theory does not consider bending-extensional coupling and cross stiffness terms, it usually gives inaccurate and nonconservative flutter boundaries for laminated plates.
A steady and oscillatory kernel function method for interfering surfaces in subsonic, transonic and supersonic flow
The theory, results and user instructions for an aerodynamic computer program are presented. The theory is based on linear lifting surface theory, and the method is the kernel function. The program is applicable to multiple interfering surfaces which may be coplanar or noncoplanar. Local linearization was used to treat nonuniform flow problems without shocks. For cases with imbedded shocks, the appropriate boundary conditions were added to account for the flow discontinuities. The data describing nonuniform flow fields must be input from some other source such as an experiment or a finite difference solution. The results are in the form of small linear perturbations about nonlinear flow fields. The method was applied to a wide variety of problems for which it is demonstrated to be significantly superior to the uniform flow method. Program user instructions are given for easy access.
Boundary layer shear stress in subsonic and supersonic flow
A wide range of shear stress distributions for turbulent boundary layers is examined. A solution for the shear stress in terms of the mean flow is obtained for the limiting case of large Reynolds numbers. Attention is given to turbulent boundary layer shear stress, zero pressure gradient flow, increasing pressure gradient flow, and decreasing pressure gradient flow.
A detailed study of attached and separated compression corner flowfields in high Reynolds number supersonic flow
An experimental study has been carried out to detail the interaction of a compressible turbulent boundary layer with shock waves of varying strengths. The interaction was produced by two-dimensional compression corners of 8, 16, 20, and 24 deg angles. The incoming boundary layer had an edge Mach number of 2.85 and a Reynolds number of 1.7 million based on overall thickness. Detailed mean-flow and surface measurements are presented for the four corner angles. The 8 deg corner flow was found to be fully attached, while the 16 deg case was near incipient separation. Both the 20 deg and 24 deg corners produced significant flow separation regions. In the discussion of these results, emphasis is placed on the development of flowfield properties from attached to separated conditions. Comparisons made with a computational solution of the Navier-Stokes equations show good agreement when the corner flow is not separated.
Interfering surfaces in subsonic, transonic, and supersonic flow
Computer program provides analysis method based on kernel-function technique that uses assumed pressure functions with unknown coefficients. With technique, generalized forces are calculated in unsteady flow, and pressure distributions are obtained in steady and unsteady flow.
Bumblebee program. Aerodynamic data. Part 1: Supersonic flow field, pressure field, and panel load data for validation of computational methods
The parts of an aerodynamics research project of the Bumblebee Program, called Generalized Missile study, is described. The source related, and potential applications are discussed.