Showing papers on "Freestream published in 1970"

An Experimental Investigation of Turbulent Slot Injection at Mach 6

Abstract: The use of film cooling for protecting a surface exposed to high temperature air at hypersonic speeds is investigated experimentally. The tests were conducted in a Mach 6 contoured axisymmetric nozzle with a streamlined centerbody. The Reynolds number in the test section was in the range of 1-3.6 X 10 6/in.; and a wall to freestream temperature ratio of 0.635. Heat-transfer distributions downstream of the slot were obtained for various mass flow rates and the effect of injection on the velocity, temperature and Mach number profiles was studied. Correlations for the cooling lengths with the blowing rate parameters X = pjUj/peUe for the various coolants-—air, helium, hydrogen, and argon were obtained. Correlations for the heattransfer rates in the form (1 — q/go.») an

Influence of measured freestream disturbances on hypersonic boundary-layer transition

Abstract: Free stream disturbances influence on hypersonic boundary layer transition Reynolds number in heated and unheated flows

Leading edge flow by the Monte Carlo direct simulation technique

Abstract: The flow near the leading edge of a sharp flat plate is studied by means of a Monte Carlo molecular simulation. This technique consists of following, by digital computation, the motion of a representative set of molecules flowing past the body while collisions are computed by statistical sampling. Flowfield properties and plate surface fluxes are presented for a monatomic gas for several molecular models: hard spheres and point centers of inverse power repulsion. The gas-surface interaction law used was a prescribed mixture of diffuse and specular reflection. Mach number was varied from 5.5 to 29.2 and both insulated and cooled plates were studied. Plate length relative to the freestream mean free path was varied from 10 to 88. For the longer plates the flow over much of the plate is that for a semiinfinite plate. Comparisons are made between these results and experimental data.

Near wake of a hypersonic blunt body with mass addition

Abstract: An experimental investigation of the steady, laminar near-wake flowfield of a two-dimensional, adiabatic, circular cylinder ·with surface mass transfer has been made at a freestream Mach number of 6.0. The pressure and mass- concentration fields associated with the transfer of argon, nitrogen, or helium into the near wake were studied for mass transfer from the forward stagnation region, and from the base. For sufficiently low mass transfer rates from the base, for which a recirculating zone exists, the entire near-wake flowfield correlates with the momentum flux, not the mass flux, of the injectant, and the mass-concentration field is determined by counter-current diffusion into the reversed flow. For mass addition from the forward stagnation region, the pressure field is undisturbed and the mass- concentration field is nearly uniform in the region of reversed flow. The axial decay of argon mass concentration in the intermediate wake, downstream of the neck, is explained with the aid of an integral solution in the incompressible plane, from which the location of the virtual origin for the asymptotic far-wake solution has been derived as one result.

Experimental investigation of flush-mounted electrostatic probes

Abstract: The behavior of flush-mounted electrostatic probes has been investigated in a pressuredriven, arc-heated shock tube over a wide range of shock tube freestream conditions and probe bias voltage. Measurements were made with large one-dimensional, flush electrostatic probes at initial shock tube pressures of 0.1 and 1.0 torr. Freestream electron densities ranged from 10 to almost 10 elec/cm. The flush electrostatic probes were biased at —3, — 15 and —90 v. The experimental results support predictions for which theories are available over the range of conditions corresponding to the case of sheath dimension small compared to the velocity boundary-layer thickness and frozen chemistry in the boundary layer. However, even at electron densities an order of magnitude below that for which the thinsheath assumption is valid, the deviation of the experimental data from theoretical predictions does not exceed the data scatter. The saturated ion current density collected by the probe was found to vary with the bias voltage raised to the one-half power.

Effects of engine exhaust flow on boattail drag

Abstract: Wind-tunnel tests were conducted to investigate the effects of jet plume shape and entrainment on hoattail pressure drag. The results were used to determine nozzle drag levels at various engine operating conditions as well as at conditions related to airplane force models. An isolated nozzle model with a pressure-tapped exterior and changeable internal parts was tested subsonically to examine changes in drag due to alterations in internal geometry and nozzle pressure ratio. In addition, tests were run with solid plume-shaped sleeves as a means to separate plume-shape effects from jet entrainment effects. Large differences in drag were measured with changes in plume shape, and, in certain regimes, jet entrainment also had a significant effect. The results of this study include boattail pressure distributions, integrated drag coefficients, and a comparison of test data with analytically predicted drag levels. Nom enclatur e A = cross-sectional area CD = boattail pressure drag coefficient (drag/g<^4.m) Cp = boattail pressure coefficient [(P — Po)/qQ] D,d = diameter h = boundary-layer height L = length of boattail MQ = freestream Mach number NPR = nozzle pressure ratio (PTJ/PO) P = local static pressure Po = freestream static pressure PTJ = exhaust jet total pressure q = dynamic pressure R = radius /3 = boattail trailing-edge angle Subscripts boattail terminal plane jet B,b J M,m T maximum freestream total

Effect of finite plate length on supersonic turbulent boundary layerwith large distributed surface injection

Abstract: The two-dimensional, supersonic, turbulent boundary layer on a finite length flat plate with large injection is analyzed using integral moment methods which include the interaction between the viscous and inviscid flows. Data obtained at Mm = 2.6 are presented and show agreement with the analysis. As injection is increased, the velocity profiles become inflected, the sonic line moves away from the wall and the flow becomes "subcritical." Under this condition, the effect of termination of injection can be felt upstream of the end of injection. In particular, as injection rates approaching the maximum value which can be entrained by a constant pressure mixing layer are approached, the analysis predicts that virtually the entire blowing region experiences a falling pressure due to the effect of finite length. It is postulated that this effect provides for a smooth transition from a boundary-layer flow to one where mixing is negligible, except in a thin layer near the streamline which divides the injected and freestream gas.

Blowing from a porous cone with an embedded shock wave

Abstract: : Massive blowing from a porous cone in a supersonic flow is considered under the assumptions of inviscid, conical flow. The injection is assumed to be uniform, normal, and supersonic. This last assumption requires a straight shock wave in the injected-flow field. The numerically obtained solutions have two noteworthy features. First, the contact surface angle relative to the cone's axis decreases sharply when the embedded shock wave moves off the body. Second, it is possible to have a completely supersonic flow from the outer shock wave to the body, so that any upstream effect due to the presence of the cone's base is eliminated. Detailed solutions are presented and a model for the porous wall is used to relate the position of the embedded shock wave to freestream and plenum conditions. (Author)

Substructure heating on cracked ablative heat shields

Abstract: = crack width, typically 0.01 in. to 0.1 in. w = wall enthalpy, Btu/lb = stagnation enthalpy, Btu/lb = crack height, in. = ablation rate, lb/ft sec = blowing rate parameter = local heat transfer to crack wall, Btu/ftsec = surface heat transfer, Btu/f t/sec = momentum thickness Reynolds number, p^ujd/^^ = friction Reynolds number, (rw/p)(pb/^} = Stanton number = freestream velocity, ft/sec = boundary-layer displacement thickness, in. = local density = freestream density = wall shear stress, lb/ft

Transitional Flow Separation Upstream of a Compression Corner

Abstract: Experimental measurements of transitional separation, caused by a compression corner on a two-dimensional model were analyzed in an attempt to correlate the extent of separation A previously published correlation, based on the wetted length Reynolds number at the separation point, failed when the plate length alone was changed A re-examination of the data showed that the true correlating parameter was the freestream unit Reynolds number Examination of the equation of motion applicable to the dividing streamline between the boundary layer and the recirculating flow suggested that the unit Reynolds number determined the growth of turbulence in the transition region The separated flow transition could then be related to attached-flow , boundary-layer transition on a plate, with the same external flow condition A new correlation showed that the length of transitional separation was a function only of freestream Mach number, inviscid flow pressure rise at the compression corner, and the length over which transition develops in an attached-flow, flat-plate boundary layer with the same external flow conditions

Flow with M/infinity/equals 1 past thin airfoils

Abstract: Pressure distribution on thin nonlifting airfoils in steady two dimensional flow with freestream Mach number at or near unity

On the critical point of the Crocco-Lees mixing theory in the laminar near wake

Abstract: The application of the multimoment integral method to the study of the base pressure behind a supersonic vehicle is examined. The primary purpose is to understand the exact nature of the Crocco-Lees critical point, which provides a uniqueness condition in the problem. The analysis is carried out mainly in Poincare phase space, where the singularities of the differential equations are investigated. Two singular curves are found. The one which is physically meaningful for the near wake flow is located downstream of the rear stagnation point. This singular curve consists of saddle points which yield “wake” solutions, and focal points or saddle-foci. Only those saddle points which yield “wake” solutions correspond to Crocco-Lees critical points. Thus, the integration of the differential equations should be started from a saddle point both for the upstream and the downstream solutions. The current analysis then requires that one deal with only a one-parameter (freestream Mach number) family of solutions, rather than the two-parameter (freestream Mach number and Reynolds number based on thickness of the viscous shear layer at the rear stagnation point) family of solutions obtained in previous works. Finally, this paper clarifies anomalous details of previous numerical investigations carried out for the near wake by Webb, Golik, Vogenitz, and Lees.

Errors in Freestream Reynolds Number of Helium Wind Tunnels

Abstract: Correction factor for free stream Reynolds number errors resulting from low temperature viscosity in He wind tunnels

Study of base pressure fluctuations behind a cone in supersonic flow

Abstract: Such fluctuations are observed during supersonic flow past axisymmetric bodies with a spike mounted in the nose section [1, 2]. The oscillations of the dimensions of the stagnant zone were studied with the aid of a high speed movie camera. Measurements were made in [3] using a piezoelectr ic sensor of the pressure oscillations on thewal l at the separation point and in the stagnant zone ahead of a twodimensional step for a freestream Mach number M~ = 2.9. Data were obtained in [4] on the fluctuation of the base pressure behind a sphere in subsonic flow. Mention should also be made of [8], devoted to the study of pressure fluctuations on a plane pointed plate with a turbulent boundary layer. These experiments give an idea of the pulsations of the freestream parameters. Below we present the remits of measurement of the base pressure behind cones for a freestream Mach number M~o = 8. The cone with hal f -angle 10 ~ and interchangeable afterbodies with base diameters of d = 10G 180, and 150 mm was mounted in the working section of the setup on two slender pointed sweptback stings located in the vert ical plane. In addition, there were fabricated two afterbodies of 150 mm diameter which had at the basetwo internalcups of 90 mm diameter; the small cup was 85 mm long, the large cup was was 70 mm long. Two static pressure taps of 1.8 mm diameter were located on each base section of the models and in the case of thecups similar taps were located in the cups. The dependence of corona discharge parameters on the pressure was used for the measurements of the pressure f}ucmations. The following design of pressure sensor was developed. A platinum point, mounted in a porcelain tube, was located at the center of an aluminum chamber. The inlet opening was shifted relat ive to the sensor axis so that the sharp edges of the opening did not fall in the act ive region of the discharge. The sequence of sensor calibration was as follows. The variation of the discharge current with discharge vokage was recorded in a vacuum chamber for various pressures. The resulting volt-ampere characteristics have a positive slope, i . e . , the sensor is stable with respect to the external circuit. The family of vol t -ampere curves determines that level of the sensor vokage for which the specified pressure range is covered. A distinctive feature of the corona discharge is that for definite relationships of the radius of the discharging electorde, the pressure, and the potential, the occurrence of regular relaxation oscillations in the audio-frequency range is possible. In the case of the occurrence of such oscillations it is necessary to alter the discharge regime with the aid of a resistance connected in the discharge circuit. A dynamic check of the sensor in the high-frequency region was made in a vacuum chamber. The pulsating flow was created by interrupting an air jet from a nozzle by a toothed disk. With a flow interruption frequency of 2000 Hz there was an area of steady-state pressure regime, while the amplitude of the pulses remained unchanged. For al l disc speeds the slope of the pulse edges was 70-80%. The temperature effect was verified by cooling the sensor by liquid nitrogen on a calibration stand. A variation of the average current through the sensor by a factor of 1.6-1.8 was noted, but the slope of the volt-ampere curves and the pressure calibration remain unchanged. Since the sensor was not intended for measuring average values, but only for measuring the pulsations, the temperature effect on the experiment results may be neglected. The output signal from the discharge sensor was fed through a cathode follower to an SZCh audiorfrequency spectrum analyzer with a range from 40 to 28 000 Hz with one-third-octave parallel filters; at the same t ime a VZ-18 quadratic vokmeter was used to measure the mean-square value of the pulsations and a peak-voltmeter VLI-3 was used to measure the maximal amplitude of the pulsations over the operating t ime period. Along with the paral lel analyzer, in some of the experiments we used an ASChKh-1 sequential heterodyne analyzer. Prior to each experiment a calibration was made of the ASChKh-1 analyzer with respect to amplitude and frequency. To do this, a signal of known frequency was applied to the analyzer input from an EG-10 audio oscillator. A voltmeter was used to monitor the amplitude of this signal. Only an amplitude calibration was made for the SZCh analyzer. The monitor photos were used for interpreting a given experiment. In paral lel with the type SZCh spectrometer these was connected an infrasonic spectrometer of the SICh type with a frequency range from 1 to 127 Hz. In the process of the conduct of the experiments the nose level of the tunnel background was measured, which was recorded by the analyzers and voltmeters. A significant background level was obtained, approximately the same on the wall of the tunnel working section and on the side wall of the cone. Over the entire range of the measured spectrum the background level was lower than the signal level from the sensor located on the base section. It was found that in the range of the low frequencies, measured by the SICh analyzer, the pulsation level at the base exceeded the tunnel background several t imes.

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.

SYNOPTIC: Influence of Measured Freestream Disturbances on Hypersonic

Abstract: This paper presents the results of a study of boundary-layer transition in hypersonic flow over a wide range of test conditions in the Langley M = 20 helium tunnel. The experiment includes: surveys of the nozzle wall boundary layer to establish if laminar, transitional, or turbulent flow exists on the nozzle wall, direct measurements of the freestream disturbances with a constant current hot-wire anemometer, and model boundary-layer transition detection using surface heattransfer rates. Boundary-layer surveys and surface Pitot pressures show that the nozzle wall boundary layer changes from laminar to turbulent over the operating stagnation pressure range of the facility. To study the attendant effects on the level of the freestream disturbances, hot-wire measurements were made in the test section near the tunnel center line; the disturbance level was nearly constant across the test core outside of the tunnel wall boundary layer. Signal spectra were recorded, and Kovasznay (and Morkovin) mode diagrams were generated over the operating pressure range. These mode diagrams, supported by the other experimental observations, identify the freestream disturbances as sound waves radiated from the turbulent boundary layer on the nozzle wall. The disturbances are lowest with a laminar boundary layer on the nozzle wall and highest when the boundary layer is transitional. When the boundary layer is turbulent the sound

Breakup of liquid sheets and jets in a supersonic gas stream

Abstract: Experimental and analytical results are presented with the objective of denning the mechanism of liquid sheet and j et breakup when subj ected to a supersonic gas stream. Liquid sheets are studied with photomicrographs and high-speed movies of the activity of a liquid layer maintained upon a porous plate test model in a parallel Macli 2.2 freestream. Tests with several different liquids show wave motion, with droplet and ligament shedding across the liquid surface. Numerical results from a liquid surface stability analysis are used to explain these observations. Liquid jets are studied with spark shadowgraphs, high-speed movies and photomicrographs of the normal injection of various liquids into a Mach 2.1 freestream. The results show that the breakup mechanism is characterized by gross jet fracture, as opposed to surface disintegration. 'the degree of breakup at a given streamwise location and jet spread after injection are found to be related to injection diameter and dynamic pressure, and certain liquid properties.

Transitional flow separation upstream of a compression corner

Abstract: Experimental measurements of transitional separation, caused by a compression corner on a two-dimensional model were analyzed in an attempt to correlate the extent of separation. A previously published correlation, based on the wetted length Reynolds number at the separation point, failed when the plate length alone was changed. A re-examination of the data showed that the true correlating parameter was the freestream unit Reynolds number. Examination of the equation of motion applicable to the dividing streamline between the boundary layer and the recirculating flow suggested that the unit Reynolds number determined the growth of turbulence in the transition region. The separated flow transition could then be related to attached-flow , boundary-layer transition on a plate, with the same external flow condition. A new correlation showed that the length of transitional separation was a function only of freestream Mach number, inviscid flow pressure rise at the compression corner, and the length over which transition develops in an attached-flow, flat-plate boundary layer with the same external flow conditions.