Showing papers on "Freestream published in 1968"

Effects of shock impingement on the heat transfer around blunt bodies.

Abstract: An extraneous shock impinging on a blunt body in hypersonic flow is observed to alter the flow around the body and increase the local heat-transfer rate near the impingement point. A novel, quasi-static technique is developed to study this phenomenon. A hemispherical, glass model, equipped with platinum thin-film thermometers, is injected into a hypersonic tunnel through a slot in a variable-incidence flat plate. Analog networks provide graphs of the heat-transfer rates at various points on the model as a function of the model's position relative to the extraneous shock. Peaks in local heat-transfer rates up to 10 times the local, unperturbed, freestream values are recorded as the model traverses the shock. The peak heating is severest on the side of the model nearest the plate and increases with increasing shock strength, occurring over a narrow region where a shear layer or jet, originating at the intersection of the bow shock and the impinging shock, meets the model surface. A physical model is set up which predicts variations in shock interference patterns, in surface pressure distributions, and in the intensity and extent of the peak heating in accordance with experiment. Nomenclature M = Mach number P = pressure normalized with respect to freestream pressure q = heat-transfer rate, cal-cm~2-sec~1 r = radius of hemisphere, mm t = time, sec T = temperature, °C U = velocity normalized with respect to freestream velocity V = injection speed, m-sec"1 x = distance along surface measured from stagnation point, mm

Analysis and design of wing-body combinations at subsonic and supersonic speeds.

Abstract: The method of aerodynamic influence coefficients has proved to be an effective tool for the analysis and design of wings, bodies, and wing-body combinations at supersonic speeds. This paper describes the extension of this method into the subsonic flow regime, and correlates the theory with experiment over a wide speed range. The method may be applied to the calculation of the pressures and forces acting on arbitrary wing-body combinations in steady flight, including aeroelastic effects, and to the design of wing camber surfaces in the presence of a body. Nomenclature A = aspect ratio b = span c =. chord C = aerodynamic coefficient d = distance / = singularity distribution function F = distribution function K = constant I = body length L — panel sweep M = Mach number r = radius Re = real part u, v, w = perturbation velocities U = freestream velocity x, y, z = Cartesian coordinates a. = angle of attack 0 = angular coordinate, panel inclination X = taper ratio A = leading edge sweep £, rj = integration variables

Effects of freestream disturbances on boundary-layer transition.

Abstract: T effects of freestream disturbances on boundary-layer transition in flows with zero pressure gradient have been studied by several investigators. Some of the notable results have been published by Dryden, Hall and Hislop, and Schubauer and Skramstad. A study similar to that of Schubauer and Skramstad was performed and reported by Wells. In the two latter studies the transition Reynolds number was found to increase systematically with decreasing freestream disturbance intensity until a maximum was reached which was not affected by further reductions in disturbance intensity. The maximum transition Reynolds number found by Wells was 80% greater than that reported by Schubauer and Skramstad, being 5.0 X 10 and 2.8 X 10, respectively, for disturbance intensities less than 0.1% of the freestream velocity Examination of the energy spectra of the freestream disturbances indicated that most of the energy in Wells' experiments occurred at frequencies below 150 cps with acoustic content less than 10% of the total energy. The tests of Schubauer and Skramstad involved significant energy levels out to 400 cps, and in addition the spectrum exhibited large acoustic energy peaks at 60 and 95 cps which accounted for approximately 90% of the total disturbance energy that was measured for intensities less than about 0.05%. These results suggest an influence of both the spectral energy distribution and the nature of the disturbances in determining the transition Reynolds number. The effects of acoustical noise are particularly interesting in the light of recent studies by Pfenninger, Bacon, and Carlson" and Jackson and Heckl concerning the stability of laminar boundary layers in the presence of intense acoustical fields. This report describes a recent series of experiments designed to extend the study reported in Ref. 4. In particular, the effects of acoustic noise fields of discrete frequencies and the broad-band turbulence created by coarse-mesh grids placed in the freestream have been investigated. These experiments are meant to be only illustrative of the problem and do not attempt to explain completely the relation between transition and free-stream disturbance energy. The results do emphasize the importance of the frequency spectra and nature of the disturbance, as well as the rms intensity of disturbance energy in predicting transition.

Similarity Parameter for Two-Stream Turbulent Jet-Mixing Region

Abstract: u = velocity component in .XT-direction Ua = velocity in faster freestream Ub = velocity in slower freestream X, Y = coordinates in the reference coordinate system 77 = dimensionless position; rj = aY/X = dimensionless velocity; = u/Ua 06 = velocity ratio of the two streams; <£& = Ub/Ua

Experimental Investigation of Laminar Near Wakes Behind 20 Wedges M=6

Abstract: An experimental investigation at Mm = 6 has been conducted to determine mean-flow properties in near wakes behind several 20° included- angle wedges at zero angle of attack. One cold-wall (H = 0.3 in., TW/TQ = 0.19) and two adiabatic-wall (H = 0.15 in., H = 0.3 in.) configurations were tested. Freestream Reynolds numbers were varied from 0.5 X 10 5 to 2 X 105 per in. for each model. Flowfield mappings and flow-property profiles were obtained in the base region for the wedge of 0.3-in. base height with and without cooling by combining Pitot- pressure data with total temperature and mass flux results from hot-wire measurements. The variation of total pressure along streamlines was initially negligible during the shearlayer turning process. Downstream boundaries of these isentropic turns corresponded to viscous-layer edges that were positioned in the outer portions of the shear layers, indicating that wake shocks originated from within viscous regions of the shear layer.

Pressure-loading functions for oscillating wings with control surfaces.

Abstract: In the linearized formulation of the oscillating-surf ace problem, singularities in the lift distribution occur at subsonic leading edges, at control surface leading edges, and in general wherever the up wash prescribed by the wing deformations is discontinuous. These singularities are examined by use of the method of matched asymptotic expansions. It is shown that for a known discontinuity in upwash (as in the case of the control surface) the form, as well as the strength, of the singularity are determined uniquely. For subsonic leading edges only the form, but not the strength, of the singularity can be determined. A discussion is also given of the proper shape of the loading functions near side edges. Nomenclature b = reference length (root semi chord) Cp = pressure coefficient k = reduced frequency, ub/Um M = Mach number of freestream MN = M cos A, Mach number normal to edge p — pressure amplitude, Cp = peikt t = time Um = freestream velocity w = amplitude of prescribed upwash on the wing X) y} z = Cartesian coordinates with x in the freestream and in the span wise direction xc = location of control-surfa ce leading edge and hinge Xj y, z = stretched variables

High-speed monorail rocket sleds for aerodynamic testing at high Reynolds numbers.

Abstract: The use of a monorail rocket-sled system is detailed and recommended as a method of obtaining aerodynamic data in a regime of high Mach and Reynolds numbers. Sled designs and the data collection system are discussed in detail. The system has been used successfully to acquire pressure data on cone-cylinders. These data are compared with data from free-flight rockets, wind-tunnel tests, and computer calculations. The sled-testing technique appears to be the most promising method of simulating flight conditions at low altitudes at Mach numbers up to M = 8. A principal advantage is the high degree of accuracy possible in determining actual test conditions, which include Mach number, static pressure, and static temperature. Exploratory tests indicate that sleds may be used to collect other types of experimental data such as heat-transfer and boundary-layer transition. To date, excellent data have been obtained at 7430 fps (M = 6.53), which to our knowledge is the highest velocity ever attained by a rocket sled. At this speed, and the Holloman Air Force Base track altitude, the Reynolds number was 42 X 106/ft, and the dynamic pressure was 55,000 psf. Nomenclature CD, CL = drag and lift coefficients, respectively, based on sled cross-sectional area M = Mach number Pm = measured model surface pressure POO = freestream ambient pressure Re = freestream Reynolds number To, Tw, Tm = total, wall, and freestream temperatures, respectively U = freestream velocity, fps

Research on slot injection into a supersonic air stream

Abstract: : The results of an experimental and analytical study of tangential slot injection into a supersonic stream are presented. The experiments were performed in an atmospheric intake wind tunnel with freestream Mach numbers of 2.85 and 4.19. Injection of air, helium and carbon dioxide at various subsonic Mach numbers and one supersonic condition was considered. Experiments for the flow over of wedges with turning angles between 5 and 25 deg. located on the wall downstream of the injection are also reported. The principal data are in the form of spark schlieren photographs, interferograms and wall static pressure distributions. Density profiles at several axial stations determined from interferograms are also presented. The transition to turbulence in the shear layer and the character of the turbulence were observed from the spark schlieren photographs. The presence of separation zones was detected by small tufts or threads on the surface. With subsonic injection, it is found that the initial slot exit conditions are not arbitrary for a given injectant mass flow but are determined by the downstream interaction between the two streams. The flow field has many of the features of the now well-known base-flow problem. A relatively simple analysis is developed which predicts the initial jet exit conditions. Very good agreement with the experimental observations is achieved.

A theoretical and experimental investigation of the compressible laminar wake behind a long slender cylinder.

Abstract: A theoretical and experimental investigation of the compressible laminar wake behind a long slender cylinder has been made. The theoretical investigation has led to approximate solutions for two-dimensional and axially symmetric flows which are valid within the entire wake region downstream of the rear stagnation point; the approximation is based on a modified Oseen-type linearization. An experiment to produce a thick laminar compressible wake behind a long thin circular cylinder was devised. The freestream Mach number and Reynolds number were 4.02 and 2300/in., respectively. The electron beam-fluorescence technique was used to obtain the density and rotational temperature profiles for stations extending from -J to 42 diam downstream from the base of the cylinder. Impact pressures were also obtained at these stations. Measurements in the separated region indicated the existence of large gradients in static pressure. Comparison of theory and experiment showed very good agreement.

An Experimental Investigation of Downstream Flow-Field Properties Behind a Forward Located Sonic Jet Injected into Transonic Freestream from a Body of Revolution

Abstract: : Transonic wind-tunnel investigation was conducted to determine the nature of the flow field downstream of a lateral sonic jet on a body of revolution The survey was made in a plane normal to the body centerline 925 body diameters aft of the lateral jet nozzle Velocity measurements were made by a remotely driven Pitot-static probe at wind-tunnel Mach numbers of 09 and 12 The data are presented in the form of Mach number vectors mapped in the normal plane for three pressure ratios and for model angles of attack of 00 and 10 deg

Linearized inviscid-flow theory of two-dimensional thin jet penetration into a stream.

Abstract: The potential flow of a stream that interacts with a two-dimensional thin jet of a different total head, being injected into the stream from an infinite plane surface at an arbitrary angle, is analyzed using natural coordinates. The velocity magnitudes along the interface and the nondimensional shape of the interface between the jet and the stream are obtained as functions of the injection angle and the ratio of the freestream velocity to the velocity in the jet at infinity downstream. Results are presented for several cases when the jet issues at oblique angles from the surface, and also the limiting case when the jet opposes the freestream. The latter case corresponds to the flow due to one branch of a translating two-dimensional jet after the jet has been split into two branches by impingement on the ground. It might also correspond to the flow of a two-dimensional thrust reverser with 180° flow reversal. The calculations show a deeper jet penetration than indicated by previous theories and by a single set of experimental results.

On the effects of a periodic, spanwise variation of suction upon asymptotic suction flow

Abstract: The effects of superposing streamwise vorticity, periodic in the lateral direction, upon two-dimensional asymptotic suction flow are analyzed. Such vorticity, generated by prescribing a spanwise variation in the suction velocity, is known to play an important role in unstable and turbulent boundary layers. The flow induced by the variation has been obtained for a freestream velocity which (i) is steady, (ii) oscillates periodically in time, (iii) changes impulsively from rest. For the oscillatory case it is shown that a frequency can exist which maximizes the induced, unsteady wall shear stress for a given spanwise period. For steady flow the heat transfer to, or from a wall at constant temperature has also been computed.

Statics and aerodynamics of lifting decelerators

Abstract: At super and hypersonic speeds, lifting deeelerators may take the form of parawings or twodimensional "sails." For freestream Mach numbers between 10 and 4, an analysis is made of the profiles assumed and isentropic waves produced in nonviscous flows by two-dimensional sails, under pure tension and of finite weight. At the higher freestream Mach numbers, large parts of the compression flow are virtually centered, and even for long sails (e.g., 100 ft chord) at a high Mach number (e.g., 10) and low stress (e.g., 5 tons/in.), the weight of such a membrane need not exceed 1 lb/ft. The two-dimensional analysis can include the effects of skin friction, and is extended to singly-curved "caret" sails, which allow leading edges to be swept but can still produce two-dimensional waves; equilibrium can still be maintained by appropriately applied tensile forces. Experimental evidence on two-dimensional, rectangular sails tends to support the theoretical predictions that much of the sail compression flow will be nearly centered.

Laminar-compressible boundary layer on a blunt body.

Abstract: T equations for the compressible two-dimensional laminar boundary layer, with heat transfer and arbitrary pressure gradient, originally set forth by Stewartson, have been solved by Cohen and Reshotkofor a number of pressure gradients and wall-to-f reestream stagnation temperature ratios. The purpose of this note is to compare some experimental data with the velocity profiles and displacement thicknesses computed from these solutions for an actual body, with nonsimilar pressure distribution and adiabatic wall. This experiment was conducted during a study of the effects of three-dimensional roughness on transition of the laminar boundary layer on a cooled blunt body. It was necessary during the course of the study to determine accurately the laminar boundary-layer profile and thickness over the body surface; however, a search of the literature failed to provide an experimental verification of the accuracy of using the Cohen and Reshotko solutions to calculate velocity profiles for the conditions of interest. Kemp et al. found that the laminar heat-transfer rates to blunt, highly cooled bodies could be predicted by assuming local similarity, but data for velocity profiles, and for conditions near adiabatic wall temperatures, are not available. The test body used was a hemisphere-cylinder with an isothermal wall temperature equal to the freestream stagnation temperature Tw = T0, in a flow of air having a freestream Mach number of 2.01. This case satisfies approximately the conditions set for the solutions given in Ref. 1: Prandtl number equal to one, linear viscosity-temperature relation across the boundary layer, and an isothermal surface. The solutions were applied in a point-by-point fashion along the surface of the body using the measured static pressure distribution. This pressure distribution, for the hemisphere-cylinder model used in the experiment, can be seen in Fig. 1. The measured Joressure distribution is compared with the Newtonian prediction,