Showing papers on "Freestream published in 1975"

Linear Stability Theory and the Problem of Supersonic Boundary- Layer Transition

Abstract: Compressible linear stability theory is first reviewed and then used to calculate the amplitude ratio of constant-frequency disturbances as a function of Reynolds number for insulated and cooled-wall flat-plate boundary layers between Mach numbers 1.3 and 5.8. These results are used to examine the consequences of using a fixed disturbance amplitude of the most unstable frequency as a transition criterion. The effect of the freestream Mach number M1 on the transition of insulated-wall boundary layers is calculated using two different assumptions concerning the initial boundary-layer disturbance amplitude A0. It is found that the shape of the transition Reynolds number Ret vs MI curve observed in wind tunnels can be closely duplicated. As a second example, the effect of wall cooling at MI = 3.0 is calculated. A much faster increase of Re, with cooling is obtained than is observed experimentally. However, when A0 is determined from the forced response of the boundary layer to irradiated sound and from the measured freestream power spectrum, a rise in Re, similar to what is observed is obtained for a certain amplitude criterion.

Boundary-Layer Transition on Supersonic Cones in an Aeroballistic Range

Abstract: Research was undertaken with the purpose of determining the effect of the unit Reynolds number on boundary-layer transition under conditions where disturbances associated with wind tunnel flows would not be present. The location of boundary-layer transition was determined from shadowgrams of nominally sharp, 4° and 10° semi-angle cones in an aeroballistic range at freestream Mach numbers of 23 and 5.0 and unit Reynolds numbers of 0.3 x 10 to 8 x 10 per in. Owing to constant and equal freestream and cone skirt temperatures, the average ratio of cone wall-to-adiabatic recovery temperature was 0.52 at Mach 2.3 and 0.19 at Mach 5.0. Features of free-flight experimentation that may be suspected of influencing boundary-layer transition were investigated. These included 1) oscillatory motion and finite angles of attack, 2) surface roughness, 3) vibration of the model, and 4) non-uniform (hot-tip) surface temperature. There was no evidence that any of these conditions influenced the major results. The data show local Reynolds number of transition increasing with unit Reynolds number for both Mach numbers. A siren was used to elevate the fluctuating sound pressure ratio by a factor of 200, but that produced no measurable effect on transition locations.

Nozzle Wall Boundary-Layer Transition and Freestream Disturbances at Mach 5

Abstract: One of the principal design objectives for a "quiet" hypersonic tunnel, where low disturbance levels are required, is to maintain laminar boundary layers on the nozzle wall at sufficiently high Reynolds numbers to obtain transition on test models Tests were conducted in a small conventional Mach 5 nozzle to investigate the effects of several factors on transition in the nozzle wall boundary layer The profiles of mean pitot pressure were measured and compared with theoretical predictions The freestream disturbance levels and spectra were also measured using a constant current hot-wire anemometer and a fluctuating pitot pressure probe The results of these measurements were used to determine whether the boundary-layer on the nozzle wall was laminar, transitional, or turbulent For the present tests, roughness in the vicinity of the nozzle throat was the dominant factor controlling transition

Comparison of Wind Tunnel Transition and Freestream Disturbance Measurements

Abstract: Boundary-layer transition measurements have been made on two 5° half-angle cones at M^ « 7 in the Ames 3.5-ft Hypersonic Wind Tunnel and the Langley Variable Density Wind Tunnel. Although there were differences between the measured freestream disturbance scales and pressure fluctuation levels, the choice of consistent locations within the transition region, using either thin-film fluctuation or surface heat-transfer data results in excellent agreement between the transition Reynolds number in both facilities. However, there is an apparent connection between changes in freestream pressure fluctuation levels and movement of boundary-laye r transition location in the two facilities. Nomenclature D — wind-tunnel test section diameter / = frequency p = pressure pc — surface pressure on cone p02 — pitot pressure Re — local Reynolds number based on distance from cone apex Rxx, RZZ — spatial correlation coefficients along the wind-tunnel axis and normal to the axis, respectively ST = Stanton number T0 = stagnation temperature Tw — wall temperature u — velocity p — density 6C = cone half-angle x,z — separation distances along the wind-tunnel axis and normal to the axis, respectively

Effect of Injector Shape on Penetration and Spread of Liquid Jets

Abstract: L injection into a supersonic air stream finds applications in supersonic combustion ramjets (scramjets), transpiration cooling of re-entry bodies and thrust vector control of rockets. The injector geometry will have a significant role in liquid injection applications. The injection of liquid into a supersonic air stream produces an interaction shock and a freestream boundary-layer separation zone upstream of the injector. The separation zone plays an important role during combustion due to the high rate of heat transfer to the wall in this region. The shock system associated with each injector in a practical supersonic combustor has two important effects: 1) it reduces the total pressure of the freestream and thus adversely affects the overall performance of the engine; 2) static temperature and pressure of the freestream rise through the injector shock system thus creating better conditions from the viewpoint of chemical reaction rates. The shape and strength of the shock also affect the forces on the liquid column and thus penetration. In general, the shock system is a strong function of injector geometry. The influence of injector geometry on penetration and spread of the jet were given major emphasis in the present investigation. The purpose was to obtain experimental penetration and spread data suitable for engineering use and to seek theoretical correlations incorporating and governing parameters. The motivation for the present work comes from the work of Kush and Schetz who observed that a liquid jet through a rectangular slot aligned with the flow gives significantly higher penetration than through a circular hole of the same area.

Internal cowl-separation at high incidence angles

Abstract: Internal flow separation at large incidence angles is studied on the basis of wind-tunnel pressure data for six axisymmetric inlet geometries. The inlet geometric variables investigated are the angle of incidence, the throat Mach number, the internal lip contraction ratio, 'sharpness' of the external shape, and freestream velocity. It is shown that an increase in lip contraction ratio delays internal flow separation, while an increase in the sharpness of the external profile tends to reduce the angle of incidence at which complete flow separation occurs. It is also shown that inlet flow separation is subject to scale effects associated with the boundary layer on the cowl surface. The scale effects are particularly pronounced at very high throat Mach numbers.

On Laminar Boundary Layers in Oscillatory Flow

Abstract: This paper describes an experimental and theoretical investigation of laminar boundary layer response to harmonic oscillations in velocity associated with a travelling wave imposed on an otherwise constant freestream velocity and convected in the freestream direction. An oscillatory flow wind tunnel and recording system were set up to produce and measure laminar boundary layer velocity profiles over frequencies of 2–10 H z for freestream amplitudes of up to 14 % of the mean velocity. An analysis on the lines of Lighthill’s theory but applying for any travelling wave convection velocity has been developed for both the low and high frequency cases. The experiments show that for a wide range of amplitude and frequency of oscillation the mean flow characteristics are the same as those of steady flow. This supports a major assumption of the theory which is linear in terms of the oscillating perturbations. Comparisons between theory and experiment show satisfactory agreement although the experimental results are largely for frequencies between the extreme ranges of the theories and they display features which are special to that intermediate frequency range. It is found that the boundary layer response is predominantly affected by the travelling wave convection velocity and frequency. In the experiments the freestream oscillation amplitudes increased with downstream distance but the effects of this increase are shown to be negligible.

Effect of free-stream turbulence on film cooling

Abstract: Film-cooling experiments were conducted at four levels of free-stream turbulence to test the hypothesis that the film-cooling effectiveness is inversely related to the free-stream turbulence level. The hot-gas operating conditions were held constant at a temperature of 590 K, a pressure of 1 atmosphere, and a velocity of 62 m/sec. The film-cooling air was at ambient inlet temperature, and the film-cooling flow rates were 2.5, 5.0, 7.5, and 10.0 percent of the total airflow. Blockage plates with blockage areas of 0, 52, 72, and 90 percent were placed upstream of the film-cooling slot and produced axial turbulence intensities of 7, 14, 23, and 35 percent, respectively. The film-cooling effectiveness decreased as much as 50 percent as the freestream turbulence intensity was increased from 7 to 35 percent. The value of the turbulent mixing coefficient used in previous work was compared with the axial turbulence intensity. The turbulent mixing coefficient was found to be 10 to 40 percent of the axial turbulence intensity.

Aerodynamic Heating to Corrugation Stiffened Structures in Thick Turbulent Boundary Layers

Abstract: The results of an experimental program to evaluate heat transfer and pressure distributions on corrugation roughened flat plates in thick turbulent boundary layers are presented. The experimental program consisted of tests in the tunnel wall boundary layers of the Langley Unitary Plan Wind Tunnel (UPWT) and Continuous Flow Hypersonic Tunnel (CFHT) at freestream Mach numbers of 2.5, 3.5, 4.5, and 10.3. Tests in the UPWT were conducted at a freestream Reynolds number/m of 10.8 x 106 and in the CFHT, at Reynolds numbers/m of 1.3 to 5.8 x 106. The test configurations consisted of 50.8 cm x 50.8 cm panels with corrugated beads of two different peak amplitudes, 0.61 cm and 0.29 cm. The angle of the corrugated beads relative to the flow direction was varied between 0° (aligned) and 90° (normal). The measured peak and average heat transfer are analyzed and correlated in terms of the bulk boundary layer, internal boundary layer, and geometric parameters. The data are also compared with similar data for thinner boundary layers, and with previously published correlation techniques.

Skin Friction Reduction by Injection Through Combinations of Slots and Porous Sections

Abstract: TT'LUID injection for thermal protection and skin -F friction reduction on flight vehicles has received considerably study. Energy conservation has now elevated the importance of skin friction reduction. The information available for supersonic speeds is limited, but it indicates drag reductions large enough to be interesting from a systems viewpoint. The competing configurations are the porous wall and the tangential slot, but there have been few studies when a direct comparison of the two schemes was made at the same nominal conditions. Further, combinations of the two schemes might be expected to be beneficial as a result of synergistic interactions. This report presents the results of a comparative study of slot injection, porous wall injection through a short strip and combinations of the two at freestream conditions of Mach 2.9, stagnation pressure of 6.9 N/m (100 psia) and total temperature of 290K. A "flat plate," solid wall configuration was also studied as a reference point. The principal data obtained were: 1) schlieren photographs; 2) wall pressure; 3) Mach number profiles; and 4) wall shear measured with a floating element balance. Wall shear was also inferred from Preston Tube measurements.

Supersonic flow past circular cones at large angles of attack

Abstract: Flow past sharp-nosed circular cones is investigated for a broad range of freestream Mach numbers M∞>1 and cone half-angles θc at angles of attack from zero to the value at which conical flow breaks down. Several new results are obtained with regard to the position of the Ferri point, the shape of the local supersonic zones and internal shock wave, and the nonmonotonicity of the windward shock slope as a function of the angle of attack. The existence of flow regimes in which the radial velocity on the windward side is directed toward the apex of the cone is demonstrated. The investigation is carried out numerically with relaxation of the solution in a fictitious time coordinate.

Flow-establishment times for blunt bodies in an expansion tube

Abstract: Flow establishment results are presented as obtained from shock standoff distance, pressure, and heat transfer measurements in the Langley expansion tube. The models tested were flat-faced cylinders with varying radius and a sphere with a constant radius, and they were positioned at the acceleration section exit and tested in the open jet at zero angle of attack. The experimental results were obtained as spinoff from various studies using helium, air, and CO2 test gases at freestream velocities in the range 5-7 km/sec. Time histories of shock detachment distance illustrate that the shock formation about the smaller-radii flat-faced cylinders and the sphere is symmetrical, whereas a complex, asymmetric formation is observed for the larger-radii cylinders. Flow is shown to establish more readily about the sphere than a flat-faced cylinder of the same diameter. A quasi-steady flow exists about relatively large blunt models during two-thirds of the approximate 250-microsec expansion tube test period.

Numerical computation of three-dimensional blunt body flow fields with an impinging shock

Abstract: A time-marching finite-difference method was used to solve the compressible Navier-Stokes equations for the three-dimensional wing-leading-edge shock impingement problem. The bow shock was treated as a discontinuity across which the exact shock jump conditions were applied. All interior shock layer detail such as shear layers, shock waves, jets, and the wall boundary layer were automatically captured in the solution. The impinging shock was introduced by discontinuously changing the freestream conditions across the intersection line at the bow shock. A special storage-saving procedure for sweeping through the finite-difference mesh was developed which reduces the required amount of computer storage by at least a factor of two without sacrificing the execution time. Numerical results are presented for infinite cylinder blunt body cases as well as the three-dimensional shock impingement case. The numerical results are compared with existing experimental and theoretical results.

Potential and viscous flow in VTOL, STOL or CTOL propulsion system inlets

Abstract: A method was developed for analyzing the flow in subsonic axisymmetric inlets at arbitrary conditions of freestream velocity, incidence angle, and inlet mass flow. An improved version of the method is discussed and comparisons of results obtained with the original and improved methods are given. Comparisons with experiments are also presented for several inlet configurations and for various conditions of the boundary layer from insignificant to separated. Applications of the method are discussed, with several examples given for specific cases involving inlets for VTOL lift fans and for STOL engine nacelles.

Nonequilibrium Viscous-Layer Computations of Space Shuttle TPS Requirements

Abstract: The thermal protection system requirements on the windward pitchplane of a typical shuttle orbiter are evaluated using a recently developed computational procedure. This procedure treats chemically nonequilibrium inviscid and viscous flows for surfaces of arbitrary catalycity. Surface catalycities for typical shuttle materials were evaluated from arc plasma generator data. Computations for the shuttle orbiter demonstrated the significance of entropy swallowing on heat transfer and show that for representative catalycity values, nonequilibrium effects are small though not insignificant. It is shown that the reduced heat transfer rates, caused by surface effects, observed in arc plasma facilities, should be reviewed carefully since these reduced rates will not be as significant for the larger orbiter dimensions. Nomenclature HT = total enthalpy kw (O) = surface catalycity for O recombination kw(N) = surface catalycity for N recombination Me = boundary-layer edge Mach number MOO = freestream or flight Mach number P = pressure Pt2 = total pressure behind a normal shock wave q = heat flux

Dynamic behavior of angle-of-attack vane assemblies

Abstract: Nomenclature A = aspect ratio, w/S b = semichord, i.e., chord = 2b CLa =wing lift curve slope, dCL/da = 7rA/2 for very small A fn = natural frequency in Hz, co „/2ir h — transverse velocity of vane pivot axis, positive down h = transverse acceleration of vane pivot axis / = mass rotary moment of inertia of the vane assembly K = emperical stiction factor in dry friction term <£ = distance between vane pivot axis and vane center of pressure M = Mach number q = freestream dynamic pressure S = vane planform area U = freestream velocity, usually VEQV = (2q/p 0) Vl w = spanwise width a = angular displacement with respect to freestream direction, positive clockwise OL = angular velocity, da/dt a. = angular acceleration, da/dt \LD = dry friction coefficient \LV = viscous friction coefficient TT -3.14159... P0 = sea level density of air f = damping coefficient un = natural frequency, rad/sec

Transport Processes in the Two-Dimensional Near Wake

Abstract: •TPESTS were conducted to characterize the two-dimensional JL recirculation zone downstream of circular cylinders and wedges, in the Reynolds number range 8 x 10 to 3 x 10. In this effort, a tracer material consisting of NO2 in a mixture of NO 2 — N204 was injected through holes in the bluff-body wall into the recirculation zone where tracer residence times and concentration distributions were determined using a fiberoptic probe. Vortex shedding and wake geometrical parameters were found from spark schlieren photographs taken with helium tracer injected into the zone. Measurements were taken at numerous approach flow velocities in the range 15 150 m/sec, at the three freestream densities of 0.41, 0.67, and 1.20 kg/m, and at several initial levels of turbulent intensity up to 11%. The data indicate that the vortex shedding phenomenon associated with two-dimensional bluff-bodies exerts a major influence on the transport of mass in the nearwake.

An experimental investigation of heat transfer to reusable surface insulation tile array gaps in a turbulent boundary layer with pressure gradient

Abstract: An experimental investigation was performed to determine the effect of pressure gradient on the heat transfer to space shuttle reusable surface insulation (RSI) tile array gaps under thick, turbulent boundary layer conditions. Heat transfer and pressure measurements were obtained on a curved array of full-scale simulated RSI tiles in a tunnel wall boundary layer at a nominal freestream Mach number of 10.3 and freestream unit Reynolds numbers of 1.6, 3.3, and and 6.1 million per meter. Transverse pressure gradients were induced over the model surface by rotating the curved array with respect to the flow. Definition of the tunnel wall boundary layer flow was obtained by measurement of boundary layer pitot pressure profiles, and flat plate wall pressure and heat transfer. Flat plate wall heat transfer data were correlated and a method was derived for prediction of smooth, curved array heat transfer in the highly three-dimensional tunnel wall boundary layer flow and simulation of full-scale space shuttle vehicle pressure gradient levels was assessed.

Two-Dimensional Turbulent Boundary Layer Separation on a Flat Plate with Ramp at Freestream Mach Numbers of 3.7 and 6.3 in Supersonic and Hypersonic Flow.

Abstract: : This investigation describes surface pressure distribution, heat transfer, and turbulent flow separation and reattachment on a flat plate with variable ramp angle at Mach numbers averaging 3.66 and 6.30. The freestream Reynold's number varied from 3.33 to 8.34 million per foot for the former case and from 5.243 million to 14.6 million per foot for the latter case. Ramp angle was varied from 0 to 35 degrees for these experiments. Stagnation temperatures averaged 849 degrees Rankine, with wall temperature assumed constant at 535 degrees Rankine for all tests. The knowledge of effective wedge angles and freestream conditions makes it possible to determine shocks so that conditions can be evaluated at all positions on the model.

Film Cooling by Oblique Slot Injection in High-Speed Laminar Flow

Abstract: Wall-cooling effectiveness is investigated for oblique injection of coolant through single or multiple wall slots into a high-speed laminar compressible boundary layer by numerical solutions of the boundary-layer equations. A grid control procedure which maintains a constant flow rate between grid lines is found to be well suited to the present injection calculations wherein the boundary-layer growth in the slot is as much as a hundred-fold and the logitudinal component of the injection velocity is in some cases as large as the freestream velocity. Film-cooling effectiveness is reported for a variety of injection configurations so that the effects of coolant mass flow rate, injection angle, upstream boundary-layer thickness, slot width, and the presence of upstream cooling slots can be investigated. For the coolant mass flow rates considered, normal injection provides better cooling than tangential injection, particularly when frictional heating effects caused by tangential injection become a dominant consideration. However, the excessive boundary-layer growth which accompanies normal injection may reduce aerodynamic performance, thus making inclined injection a desirable compromise.

The slip-dominated merged layer regime for newtonian hypersonic flow past a flat plate*

Abstract: The rarefied uniform hypersonic flow past the leading edge of a sharp flat plate at zero angle of attack is analyzed on the basis of a continuum model consisting of the Navier–Stokes equations and the velocity-slip and temperature-jump plate boundary conditions. The model fluid is a perfect gas having constant specific heats, a constant Prandtl number of order unity, and first and second viscosity coefficients varying as a power of the absolute temperature. For this flow, it is taken that the Newtonian parameter, $\varepsilon = ( {\gamma - 1} )/( {\gamma + 1} )$, goes to zero, and that the freestream Mach number, $M = ( \rho _\infty u_\infty ^2 /\gamma p_\infty )^{1/2} $, the stagnation temperature parameter, $\theta _S = \{ 1 + \varepsilon )/( 1 - \varepsilon )\}\varepsilon M^2 $, and the free-stream Reynolds number (based on the characteristic axial length from the leading edge), $R_L = \rho _\infty u_\infty L/\mu _\infty $, go to infinity.For the viscosity-temperature exponent, $\omega $, satisfying $1...

STOL aircraft transient ground effects. Part 1: Fundamental analytical study

Abstract: The first phases of a fundamental analytical study of STOL ground effects were presented. Ground effects were studied in two dimensions to establish the importance of nonlinear effects, to examine transient aspects of ascent and descent near the ground, and to study the modelling of the jet impingement on the ground. Powered lift system effects were treated using the jet-flap analogy. The status of a three-dimensional jet-wing ground effect method was presented. It was shown, for two-dimensional unblown airfoils, that the transient effects are small and are primarily due to airfoil/freestream/ground orientation rather than to unsteady effects. The three-dimensional study showed phenomena similar to the two-dimensional results. For unblown wings, the wing/freestream/ground orientation effects were shown to be of the same order of magnitude as for unblown airfoils. This may be used to study the nonplanar, nonlinear, jet-wing ground effect.

Rough wall turbulent heat transfer with variable velocity, wall temperature, and blowing

Abstract: An integral prediction method is presented which accurately describes Stanton number behavior for a fully rough turbulent boundary layer flowing over a uniformly rough surface. The kernel function which represents the response of such a system to an unheated starting length is shown to also describe the response to variations in freestream velocity, surface temperature and blowing. Predictions are compared with experimental data for cases of variable wall temperature, favorable pressure gradients and variable blowing. Agreement is excellent in all cases. (auth)