Showing papers on "Freestream published in 1984"

Disruptive and Micro-Explosive Combustion of Free Droplets in Highly Convective Environments

Abstract: Experinlental evidence is presented showing the existence of disruptive burning for two-component fuel droplets and micro-explosive combustion for water-in-singlc fuel component emulsificd drops subjected to high relative velocities ranging from 16 to 19 m/sec. Reynolds numbers, based on the high temperature (1200-1400 K) convective freestream conditions and droplet diameter, exceedcd 40. Calculations using the one-third averaging rule suggested by other investigators for evaluating the relevant properties lead to Reynolds numbers exceeding 135. The fuels considered in this study were a solution of 50 vol.% 11-hexane-50 vol.% n-hexadecane and water-in-11-hexadecane cniulsions, with 9 and 18 percent (volume) of water. The substantial decrease i n time from ignition-to-disruption (with increased relative gas-droplet velocity) in combination with the observcd "droplet-shoulder-orientcd disruptions" for. Ihe multi-component solulion supports the existence of a toroidal vortex-like structure within th...

Supersonic axisymmetric flow over boattails containing a centered propulsive jet

Abstract: The influence of underexpanded jets on a supersonic afterbody flowfield is investigated using computational techniques. The thin-shear-layer formulation of the compressible, Reynolds-averaged, Navier-Stokes equations is solved using a time-dependent, implicit numerical algorithm. Solutions are obtained for supersonic flow over an axisymmetric conical afterbody containing a centered propulsive jet where the freestream Mach number is 2.0 and the jet exit Mach number is 2.5. Exhaust-jet static pressures are considered in the range of 2 to 9 times the freestream static pressure and with nozzle-exit half-angles from 15 to 43 deg. Comparisons are made with experimental results for base pressure, separation distance, afterbody pressure distribution, and flowfield structure. Although good quantitative agreement with experimental separation distance and base pressure level is not observed, the parametric trends induced by exhaust-jet pressure level and nozzle-exit angle are well predicted, as well as the flowfield details in the vicinity of the afterbody and in the exhaust plume.

Hot-wire investigation of an unseparated shock-wave/turbulent boundary-layer interaction

Abstract: Detailed hot-wire measurements of the longitudinal component of the mass-flow fluctuations have been made for an adiabatic fully attached shock-wave/turbulent boundary-layer interaction. The shock wave is induced by an 8-deg compression corner at a Mach number of 2.85 and a unit Reynolds number of 6.3 X10 m ~ l . The data indicate that the absolute mass-flow turbulence intensity increases abruptly through the shock wave, and continues to increase with further distance downstream. When nondimensionalized by the local freestream massflow rate the maximum turbulence intensity actually exceeds the upstream equilibrium value at the last measurement station. Measurements of the probability density of the fluctuations reveal only minor difference between the boundary-layer behavior upstream and downstream of the shock. A significant dip in the kurtosis was observed near the point of maximum intensity, which also coincided with the onset of intermittency.

A new consistent spatial differencing scheme for the transonic full-potential equation

Abstract: A new spatial differencing scheme for the transonic full-potential equation in conservative form has been developed. Three consistency conditions for the full-potential equations are derived and are satisifed by the new scheme. This scheme guarantees zero truncation error on any curvilinear mesh for freestream flows in either two- or three-space dimensions. Solutions obtained with this new differencing scheme, away from freestream regions, exhibit greatly improved accuracy, especially for nonsmooth or singular meshes. The computing times associated with the new scheme are approximately the same as the less accurate old scheme when computations are performed on the same mesh.

Analysis of large solid propellant rocket engine exhaust plumes using the direct simulation Monte Carlo method

Abstract: A new solution procedure has been developed to analyze the flowfield properties in the vicinity of the Inertial Upper Stage/Spacecraft during the 1st stage (SRMI) burn. Continuum methods are used to compute the nozzle flow and the exhaust plume flowfield as far as the boundary where the breakdown of translational equilibrium leaves these methods invalid. The Direct Simulation Monte Carlo (DSMC) method is applied everywhere beyond this breakdown boundary. The flowfield distributions of density, velocity, temperature, relative abundance, surface flux density, and pressure are discussed for each species for 2 sets of boundary conditions: vacuum and freestream. The interaction of the exhaust plume and the freestream with the spacecraft and the 2-stream direct interaction are discussed. The results show that the low density, high velocity, counter flowing free-stream substantially modifies the flowfield properties and the flux density incident on the spacecraft. A freestream bow shock is observed in the data, located forward of the high density region of the exhaust plume into which the freestream gas does not penetrate. The total flux density incident on the spacecraft, integrated over the SRM1 burn interval is estimated to be of the order of 10 to the 22nd per sq m (about 1000 atomic layers).

Thermofluiddynamic experiments with a heated and rotating circular cylinder in crossflow

Abstract: Two-dimensional boundary layer profiles of heated and rotating circular cylinders in crossflow were investigated for subcritical freestream-Reynolds-numbersRe = 34 · 104 andRe = 48 · 104 The peripheral speed of the cylinder surface corresponds to velocity ratios 0 ≦ ξ ≦ 2(α, circular/freestream-velocity) Special attention was focused on the location of separation points, which was determined as a function of the velocity ratio a whereby a strong influence of wall temperature was revealed

Effect of geometry on airfoil icing characteristics

Abstract: A droplet trajectory computer code is used to predict the water droplet impingement characteristics of several low- and medium-speed airfoils. The maximum impingement efficiency, total collection efficiency, and limits of impingement are analyzed as functions of the airfoil geometry and freestream conditions. The airfoil geometry is represented by leading edge radius, maximum thickness, maximum camber, and angle of attack. The analysis shows that the primary effects are an increase in maximum impingement efficiency with a decrease in leading edge radius, a reduction in total collection efficiency for thicker airfoils, and a change in the limits of impingement for airfoils of different maximum camber.

Shock-induced dynamic stall

Abstract: At freestream Mach numbers above M= 0 3 shock/boundary layer interaction begins to complicate the unsteady airfoil stall characteristics The present paper shows how theoretical relationships can be developed for the interdependence between unsteady and steady characteristics to provide the means whereby the shock induced dynamic stall characteristics can be determined if the static characteristics are known, e g , from ex periments

Mixing layer produced by a screen and its dependence on initial conditions

Abstract: This paper aims to elucidate the structure of the turbulent mixing layers, especially, its dependence on initial disturbances. The mixing layers are produced by setting a woven-wire screen perpendicular to the freestream in the test section of a wind tunnel to obstruct part of the flow. Three kinds of model geometry are treated; these model screens produced mixing layers which may be regarded as the equivalents of the plane mixing layer and of two-dimensional and axisymmetric wakes issuing into ambient streams of higher velocity. The initial disturbances are imposed by installing thin rods of various sizes along the edge of the screen or at the origin of the mixing layer. Flow features are visualized by the smoke-wire method. Statistical quantities are measured by a laser-Doppler velocimeter. In all cases large-scale transverse vortices seem to persist, although comparatively small-scale vortices are superimposed on the flow field in the mixing layer. The mixing layers are in self-preserving state at least up to third-order moments, but the self-preserving state is different in each case. The growth rates of the mixing layer are shown to depend strongly on the initial disturbance imposed.

Characteristic boundary conditions for three-dimensional transonic unsteady aerodynamics

Abstract: Characteristic far-field boundary conditions for the three-dimensional unsteady transonic small disturbance potential equation have been developed. The boundary conditions were implemented in the XTRAN3S finite difference code and tested for a flat plate rectangular wing with a pulse in angle of attack; the freestream Mach number was 0.85. The calculated force response shows that the characteristic boundary conditions reduce disturbances that are reflected from the computational boundaries.

Comparison of viscous-shock-layer heating analysis with Shuttle flight data in slip flow regime

Abstract: Comparison of STS-2 Shuttle flight heating data along the windward centerline has been made with two-dimensional nonequilibrium viscous shock-layer solutions obtained with shock and wall-slip conditions at an altitude range of 90 to 110 km. The shock slip condition used is the modified Rankine-Hugoniot relations of Cheng as used by Davis, and the wall-slip conditions are based on the first order consideration derived from kinetic theory as given by Scott and Hendricks. The results indicate that the calculated heating distributions with slip boundary conditions agree better with the flight data than those without slip conditions. The agreement improves when the accommodation coefficient or freestream density is decreased to one-half, suggesting the possibility of less than full accommodation for the tile surface and (or) an overestimate of freestream density using the Jacchia-Roberts model. Heating reduction due to the slip effect becomes very pronounced as the flow becomes more rarefied, and the effect is more significant for the stagnation region than the aft region of the vehicle.

Supersonic Compressive Ramp Without Laminar Boundary-Layer Separation

Abstract: Similar, laminar, boundary-layer theory for a two-dimensiona l or axisymmetric body is extended by a derivation that yields explicit equations for the wall shape. Isentropic edge conditions are assumed, which results in a differential equation for the pressure gradient parameter ft. The solution of this equation, when ft is constant, parametrically yields the wall shape with the edge Mach number as the parameter. A two-dimensional wall shape is determined when the freestream is supersonic. For a compressive turn, the boundary layer does not separate if ft is not too negative. In this case, the magnitude of ft depends on the ratio of wall temperature to the freestream stagnation temperature. For this application a criterion is provided for the validity of the isentropic edge assumption. A transformation is given for axisymmmetric body shapes.

Treatment of supersonic configurations by an updated low-order panelmethod

Abstract: A recently updated version of the Woodward linearized subsonic/supersonic panel method (USSAERO) has been applied to the calculation of the supersonic characteristics of wing body combinations The use of a new singularity with directional properties (triplet) for representing body effects and the newly developed extension to the supersonic case of a nonplanar boundary condition over the wing(s) overcomes many of the shortcomings exhibited by former USSAERO versions when analyzing complex configurations, such as the fighter type air plane at supersonic speeds Examples of applications to significant test cases are presented and discussed Comparisons of the results with other theoretical and/or experimental data demonstrate capabilities and limitations of the present method Nomenclature b = span length c = chord length CD = drag coefficient CL = lift coefficient CM = pitching moment coefficient Cp = pressure coefficient d = body diameter length L = body length M = freestream Mach number p = pressure Poo = freestream pressure Re - Reynolds number u v w = perturbation velocity components V = total velocity magnitude Fa, = freestream total velocity magnitude x y z = Cartesian coordinates a = angle of attack j8 = Prandtl Glauert number, = Vl -M2 ACp = load coefficient 7 = specific heat ratio 0 = perturbation potential

Numerical simulation of two-dimensional inlet flowfields

Abstract: Inlet flowfields for airbreathing missiles are calculated by a two-dimensional computational method. A supersonic freestream is assumed to allow the forebody calculation to be uncoupled from the inlet calculation. The inlet calculation employs an implicit, time-marching, finite difference procedure to solve the Euler equations formulated in body-fitted coordinates. The method can be used for a flowfield with both subsonic and supersonic regions and is found to converge rapidly for supercritical inlet operation. For subcritical inlet operation, however, convergence to steady state is slow.

Flowfield computation of an aerobrake orbital transfer vehicle

Abstract: Viscous real gas and ideal gas supersonic flowfields over the forebody of a aeroassisted orbital transfer vehicle are determined using a unsteady factored implicit algorithm. Air in chemical equilibrium is considered and its local thermodynamic properties are computed by an equilibrium composition method. Numerical solutions are obtained for both real and ideal gases at a Mach number of 30 and at angles of attack up to 20 degrees. Shock stand-off distances and surface pressure distributions are presented for the gas models with and without viscous effects. For the freestream conditions considered, viscous effects dominate the flow.

Evaluation of NCOREL, PAN AIR, and W12SC3 for Supersonic Wing Pressures

Abstract: The ability of three recently developed computer programs to predict pressures on a supersonic maneuver wing has been evaluated. The NCOREL program was the only code capable of predicting the nonlinear leeward leading edge supercritical crossflow and resultant shock wave formation at the higher angles of attack Both linear panel methods, PAN AIR and W12SC3, performed reasonably well at the lower angles of attack Im plementing the freestream axis as the compressibility axis, PAN AIR overpredicted the windward pressures in comparison to the W12SC3 program at higher angles of attack

The Effects of Freestream Turbulence on the Profile Boundary Layer and Losses in a Compressor Cascade

Abstract: The turbulent profile boundary layer on a one-foot chord compressor cascade blade has been measured with varying levels of freestream turbulence. Increased levels of freestream turbulence were found to increase the fullness of the velocity profiles, with a consequent decrease in displacement thickness and an increase in the skin friction coefficient. A small increase in freestream turbulence causes the cascade total-pressure loss to increase initially, while at the higher turbulence levels boundary layer separation was delayed, resulting in a decrease in the total-pressure loss and deviation angle.Copyright © 1984 by ASME

Flow Separation Induced by Periodic Aerodynamic Interference

Abstract: : Analysis of experimental surface pressure distributions on an NACA 0012 airfoil has revealed four flow states: attached, separated, borderline, and dynamically separated. The important parameters that determine the flow state are Reynolds number, reduced frequency, airfoil angle of attack, and surface condition at the leading edge. Testing was done at Re=700,000, .5=k=6.4, and 0=alpha=18 degrees. For this flow the dynamically separated state takes the form of an alternation between attached and separated flow. It has a period that ranges from 1 to 30 times that of the unsteady perturbation. In the separated state a convected surface pressure disturbance was identified, and found to propagate downstream from a location near the leading edge at a phase speed of 1/3 to 1/2 that of the freestream velocity.

Large-Eddy Structures in a Turbulent Mixing Layer

Abstract: Structures of turbulent mixing layers, especially of large-eddies, are investigated by visual observation by means of a smoke-wire method. Mixing layers are produced by a screen set perpendicular to freestream in a test section. Flow features of the mixing layers are observed by means of motion pictures taken with a 16 mm high-speed camera. Results show that 1) large-eddies form from the amplification of unstable small-amplitude waves, 2) vortex pairing process as well as the growth of large-eddies without vortex pairing, found by Hernan & Jimenez, play an important role in the growth of the turbulent mixing layer, and 3) the mixing layers are fundamentally two-dimensional.

Transient Ablation of Blunt Bodies at Angles of Attack

Abstract: An experimental investigation of transient ablation is presented, with particular emphasis on the transient shape changes of blunt bodies at angles of attack. Models were made of pure Teflon and Teflon mixed with glass powder, respectively. Initial body shapes were the flat-faced cylinder and the hemisphere-cylinder, respectively. The high-temperature region in which ablation occurs appears on the windward side of a body at angles of attack, and has great influence on the apparent recession depth results. It is revealed that the enhancement of heat transfer on the windward surface, as well as that on the front surface of the body, plays an important part during entry processes. Nomenclature D - model diameter / = distance from the nozzle exit MO, = freestream Mach number Po = reservoir pressure Re = Reynolds number T = temperature t = time (x,y, 0) = cylindrical coordinates a = angle of attack &X - recession depth

Performance tests for the NASA Ames Research Center 20 cm x 40 cm oscillating flow wind tunnel

Abstract: An evaluation is presented of initial tests conducted to assess the performance of the NASA Ames 20 cm x 40 cm oscillating flow wind tunnel. The features of the tunnel are described and two aspects of tunnel operation are discussed. The first is an assessment of the steady mainstream and boundary layer flows and the second deals with oscillating mainstream and boundary layer flows. Experimental results indicate that in steady flow the test section mainstream velocity is uniform in the flow direction and in cross section. The freestream turbulence intensity is about 0.2 percent. With minor exceptions the steady turbulent boundary layer generated on the top wall of the test section exhibits the characteristics of a zero pressure gradient turbulent boundary layer generated on a flat plate. The tunnel was designed to generate sinusoidal oscillating mainstream flows. Experiments confirm that the tunnel produces sinusoidal mainstream velocity variations for the range of frequencies (up to 15 Hz). The results of this study demonstrate that the tunnel essentially produces the flows that it was designed to produce.

A Numerical Study of Gas-Particle Supersonic Flow Past Blunt Bodies : The Case of Two-Dimensional Flow

Abstract: A numerical method (inverse method) was developed for a gas-particle supersonic flow past two-dimensional blunt bodies. This method is based on two transformations (von Mises and additional one), which are convenient for determining the shock layer flow fields and the body shapes. Using the present method, the pure gas flow field around a circular cylinder was first solved numerically for the freestream Mach numbers M∞=2.0 and 3.0. Then the gas-particle flow in the shock layer around blunt bodies (nearly circular cylinders) was solved for freestream Mach number M∞=3.0, with particle diameter dp 2, 5, 10 μm and freestream loading ratios α=0, 0.2, 0.5 and 1.0, respectively. The effects of dp and α on the shock stand-off distance, the body surface pressure and flow quantities along the stagnation streamline are discussed, and the flow patterns in the shock layers are also shown.

Heat-transfer distributions for biconics at incidence in hypersonic-hypervelocity real-gas flows

Abstract: Laminar heat-transfer distributions were measured on spherically blunted, 13/7 deg straight and bent biconics at freestream velocities from 4.5 to 6.9 km/s and Mach numbers from 6 to 9. The flows were generated in the NASA's Langley Expansion Tube using helium, nitrogen, air, and carbon dioxide; angle of attack, referenced to the axis of the aft cone, was varied from zero to 20 deg. The penalty in windward heating to the fore cone due to the 7-deg nose bend diminished rapidly with increasing angle of attack and was only 10 to 20 percent at the design trim angle of attack of 20 deg. Leeward heating initially decreased, then increased, with increasing angle of attack. Windward heating rates predicted with a computer code that solves the parabolized Navier-Stokes equations were in good agreement with measurements for helium and air. The study used a 1.9-percent scale model of the proposed generic planetary vehicle and is directly applicable to the orbital transfer vehicle, which incorporates a spherically blunted biconic.

Atmospheric boundary layer testing: hot-wire anemometry measurements of turbulent boundary layer flow past a two-dimensional obstacle. 1982-1983 equipment loan report. Supplement 2

Abstract: Measurements of a zero-pressure-gradient turbulent boundary layer flow past a two-dimensional obstacle were made in the present study. Measurements were made for both smooth and rough surfaces using single and X hot-wire probes. The Reynolds number based on obstacle height and freestream velocity was about 15,302. Profiles of mean velocity, turbulent intensity and probability density functions in two dimensions were determined. Also, Reynolds stress profiles, energy spectra and second moments of energy spectra were obtained. From the results evidence emerged that upstream, over, and downstream of the obstacle there zones of recirculating flow. The flow-field was dominated by the obstacle presence, such that no distinction between smooth-and rough-surface measurements could be made.

The role of freestream turbulence scale in subsonic flow separation

Abstract: The clarification of the role of freestream turbulence scale in determining the location of boundary layer separation is discussed. Modifications to the test facility were completed. Wind tunnel flow characteristics, including turbulence parameters, were determined with two turbulence generating grids, as well as no grid. These results are summarized. Initial results on the role of scale on turbulent boundary layer separation on the upper surface of an airfoil model are also discussed.

An investigation of the influence of freestream turbulence on the thermal characteristics around a single tube

Abstract: Experiments have been carried out to study the influence of freestream turbulence on the transport of heat and momentum in the boundary layers on a single tube in cross flow. The ratio of the integral scale of turbulence to tube diameter was kept small, less than 0.14, while the turbulence intensity was ranging from 0.1 − 11.5 %. Special emphasis was given to the behaviour of the fluctuating components within the boundary layers. The investigation was carried out for the two Reynolds numbers 5.104 and 1.105.