Showing papers on "Freestream published in 1993"

Wing flutter boundary prediction using unsteady Euler aerodynamic method

Abstract: Modifications to an existing three-dimensional, implicit, upwind Euler/Navier-Stokes code (CFL3D Version 2.1) for the aeroelastic analysis of wings are described. These modifications, which were previously added to CFL3D Version 1.0, include the incorporation of a deforming mesh algorithm and the addition of the structural equations of motion for their simultaneous time-integration with the government flow equations. The paper gives a brief description of these modifications and presents unsteady calculations which check the modifications to the code. Euler flutter results for an isolated 45 degree swept-back wing are compared with experimental data for seven freestream Mach numbers which define the flutter boundary over a range of Mach number from 0.499 to 1.14. These comparisons show good agreement in flutter characteristics for freestream Mach numbers below unity. For freestream Mach numbers above unity, the computed aeroelastic results predict a premature rise in the flutter boundary as compared with the experimental boundary. Steady and unsteady contours of surface Mach number and pressure are included to illustrate the basic flow characteristics of the time-marching flutter calculations and to aid in identifying possible causes for the premature rise in the computational flutter boundary.

Zonal Two Equation Kappa-Omega Turbulence Models for Aerodynamic Flows

Abstract: Two new versions of the kappa-omega two-equation turbulence model will be presented The new Baseline (BSL) model is designed to give results similar to those of the original kappa-omega model of Wilcox, but without its strong dependency on arbitrary freestream values The BSL model is identical to the Wilcox model in the inner 50% of the boundary-layer but changes gradually to the standard kappa-epsilon model (in a kappa- omega formulation) towards the boundary-layer edge The free shear layers The second version of the model is called Shear-Stress Transport (SST) model It is a variation of the BSL model with the additional ability to account for the transport of the principal turbulent shear stress in adverse pressure gradient boundary-layers The model is based on Bradshaw's assumption that the principal shear-stress is proportional to the turbulent kinetic energy, which is introduced into the definition of the eddy-viscosity Both models are tested for a large number of different flowfields The results of the BSL model are similar to those of the original kappa-omega model, but without the undesirable freestream dependency The predictions of the SST model are also independent of the freestream values but show better agreement with experimental data for adverse pressure gradient boundary-layer flows

Multi-zone behavior of transverse liquid jet in high-speed flow

Abstract: The behavior of a liquid-fuel spray transversely injected into a uniform high-speed crossflow has been characterized using a laser-sheet imaging technique. The dependence of jet penetration upon jet-to-crossflow momentum ratio was studied by varying the ratio from 3 to 45. The static pressure inside the test section was varied from 14.7 to 30 psia, while the freestream Mach number was held constant at 0.4. A detailed comparison of the jet trajectories (penetration profiles) measured in this study with those predicted by currently available correlation functions revealed gross discrepancies. These discrepancies were attributed to the fact that the spray plume consists of several zones, i.e., a liquid column adjacent to the injector and ligament and droplet regions, exhibiting different characteristics which cannot be adequately described by empirical functions. A composite functional form which takes into account the behavior of these fundamcntally different spray regions has been formulated to provide a more accurate description of the penetration profile of the spray plume. The proposed empirical formula also describes the maximum (asymptotic) penetration of the spray plume in the far field. The dependence of asymptotic penetration hcight upon momentum ratio was analyzed to yield a general formula for predicting the spray trajectory over a wide range of momentum ratios. d’

Experimental Supersonic Hydrogen Combustion Employing Staged Injection Behind a Rearward-Facing Step

Abstract: An experimental investigation of a Mach 2 combustor has been conducted in order to characterize flow properties in a supersonic reacting flowfield. Hydrogen was injected transversely as staged, underexpanded jets behind a rearward-facing step into a ducted Mach 2 air freestream. The effects of the chemical reaction on the supersonic flowfield was investigated using shadowgraphs, broadband flame emission photography, and planar laser-induced fluorescence of OH. The shadowgraphs indicated that the wave pattern in the combustor along with flowfield unsteadiness was strongly affected by the heat release. The broadband flame emission photographs revealed large regions of no combustion in the vicinity of the fuel injectors where fuel/air mixing was insufficient to support combustion. These regions decreased in size as the freestream stagnation temperature was decreased for fixed hydrogen mass flow rate, consistent with an increase in the effective g-ratio with combustion. The size of the zones containing OH in the planar fluorescence images also increased as the main flow stagnation temperature was decreased. Reaction zones were found in the planar fluorescence images away from regions containing inject ant in a nonreacting study of the same geometry, indicating that the pressure rise associated with the reaction forced a large redistribution of the fuel.

Fluorescence imaging of OH and NO in a model supersonic combustor

Abstract: Recent results are presented from an experimental study to develop Planar Laser-Induced Fluorescence (PLIF) diagnostics for application to scramjet combustor development. The measurements are made in a reacting flow shock-tunnel facility which generates Mach 3 air at static conditions of 1500 K and 0.3 atm. Hydrogen or other gases may be injected into the rectangular test section downstream of a rear-facing step. PLIF measurements of NO, naturally present in the reflected shock-heated, freestream air, are used to determine basic flow features, and a technique is demonstrated for determining static temperature variations in nonreacting, non-mixing portions of the flow. pLIF measurements of OH are used to reveal the instantaneous reaction zone interface from a dual round jet injector configuration

Measurement and analysis of nitric oxide radiation in an arc-jet flow

Abstract: On the bases of the centerline enthalpy value deduced from heat transfer measurements and the NOZNT code, it is possible to predict the freestream conditions in an arcjet wind tunnel flow. The translational-rotational and vibrational temperature of NO is nearly reproducible by NOZNT. Relative to the electron and electronic temperatures, the vibrational temperature of N2 and NO are significantly lower at enthalpies of less than 45 MJ/kg. The enthalpy deduced from spectroscopic measurements is in rough agreement with that deduced from heat transfer measurements.

Simulation of ablation in Earth atmospheric entry

Abstract: The process of ablation for Earth atmospheric entry is simulated using a computational approach that allows thermo-chemical nonequilibrium of the flow field and ablation gases The heat pulse into the heat shield is modeled The flowfield and graphite heat shield are coupled through surface mass and energy balances The surface thermochemistry involves the oxidation of graphite and allows for catalytic recombination of diatomic oxygen Steady-state simulations are performed on a one meter nose radius sphere at an altitude of 65/km and at freestream velocities of 8 km/s and 10 km/s A transient simulation is performed at 65 km altitude and a freestream velocity of 10 km/s

Parameters governing the turbulent wall jet in an external stream

Abstract: Mean velocity distributions in a plane, turbulent, and fully developed wall jet embedded in a uniform stream were measured for a variety of initial velocity ratios and Reynolds numbers. It was determined that the bulk of the flow is self-similar, provided the maximum velocity in the jet is twice as large as the freestream velocity. The normalized velocity profile depends on two velocity scales and on two length scales that, in turn, depend on the momentum flux at the nozzle, the viscosity, and the initial velocity ratio between the jet and the freestream defined by R≡(U j -U ∞)/(U j +U ∞). The width of the nozzle that was commonly used to reduce these data has no part in the similarity considerations

Effectiveness of the Gas Turbine Endwall Fences in Secondary Flow Control at Elevated Freestream Turbulence Levels

Abstract: A secondary flow management technique which employs a boundary layer fence on the endwall of a gas turbine passage is evaluated under freestream turbulence conditions that are representative of turbine conditions. A turbulence generator, which was able to reproduce the characteristics of the combustor exit flow, was used. The horseshoe and passage vortices observed in previous tests with low turbulence level remain coherent and strong within the cascade passage when the intensity is elevated to 10 percent. A boundary layer fence on the endwall remains effective in changing the path of the horseshoe vortex and reducing the influence of the vortex on the flow near the suction wall at the high freestream turbulence level. The fence is more effective in reducing the secondary flow for the high turbulence case than for a low 11 case, probably because the vortex which has been deflected into the core flow diffuses and dissipates faster in the more turbulent flow. The fence decreases aerodynamic losses for streamlines within the core of the channel flow. NOMENCLATURE chord� (mm) specific heat of the air (kJ/kgK) total pressure coefficient. (Pt-Ptr)/(0.5pU.2) secondary kinetic energy coefficient, (v2+w2)/UO2 height of the fence� (mm) pressure side leg of horseshoe vortex suction side leg of horseshoe vortex heat transfer coefficient (W/m2K) power spectral density (e.g. u'2(f, df)/di), � (m2/s2) total pressure total pressure in freestream upstream of the cascade Reynolds number Reynolds number based on the chord length curvilinear distance from stagnation line along suction wall (mm)

Measurements of laminar mixed convection flow over a horizontal forward-facing step

Abstract: Measurements and predictions of buoyancy-assisting laminar mixed convection flow over a horizontal, two-dimensional forward-facing step are reported. Laser-Doppler velocimeter (LDV) and cold wire anemometer were used to simultaneously measure the velocity and the temperature distributions, respectively. Flow visualizations were conducted to determine the reattachment lengths for different inlet velocities (u(0) between 0.255 m/s and 0.50 m/s), wall freestream temperature differences (Delta-T between 0 C and 37 C), and step heights (s between 0.79 cm and 1.75 cm). The results reveal that the buoyancy force due to wall heating has a negligible effect on the velocity and temperature distributions and the reattachment lengths, as long as the flow remains stable and two-dimensional. The inlet velocity and the step height, on the other hand, significantly affect the flow and thermal fields. The local heat transfer coefficient is found to increase as the inlet velocity increases and the step height decreases. On the other hand, the length of the recirculation regions upstream and downstream of the step are found to increase as the inlet velocity and the step height increase. Correlation equations are developed to predict the reattachment lengths that appear upstream and downstream of the step. The measured results agree well withmore » numerical predictions. 10 refs.« less

Modeling supersonic inlet boundary-layer bleed roughness

Abstract: Boundary-layer mass removal (bleed) through spanwise bands of holes on a surface is used to prevent or control separation in supersonic inlets. The rough wall algebraic turbulence model of Cebeci and Chang was added to both boundary-layer and Navier-Stokes analyses to simulate the overall effect of bleed on the growth of a boundary layer. Roughness values were determined for seven bleed configurations, a range of Mach numbers between 1.3-4, and bleed rates between zero and choked values. For the bleed experiments considered, the roughness was found to be a function of the fraction of the upstream boundary-layer mass flux removed. Choked bleed flow through holes at a low angle, with respect to the surface, minimized the roughness effect and gave the best improvement in the boundary-layer velocity distribution for separation control. Nomenclature A+ = Van Driest parameter, 26 d ~ bleed hole diameter k = von Karman constant, 0.4 ks = equivalent sand grain roughness Lid — hole aspect ratio / = local turbulent length scale M = Mach number N = number of rows of bleed holes in bleed band P/POO = local static pressure/freestream static pressure R = roughness parameter, in. ulue — ratio of local to freestream velocity within the boundary layer «r = (TJPJ"1 u' = local turbulent velocity scale

Transonic shockwave/turbulent boundary layer interactions on a porous surface

Abstract: Transonic shockwave/turbulent boundary layer interactions on a porous surface above a closed plenum chamber have been studied experimentally in the choked flow of a windtunnel test-section. The equivalent freestream Mach number is 0.76 and results were obtained for three shock strengths. Without the porous surface the Mach numbers ahead of the shock were 1.13, 1.18 and 1.26. The respective shock Mach numbers with the porous surface were 1.10, 1.11 and 1.19. Laser holographic interferometry results are used to measure the density flowfield and examine the nature of the interaction. These results show that the interaction on the porous surface is modified by a thin shear layer adjacent to the surface and the weakening of the Shockwave is attributed to this. The interaction was also studied by solving the two-dimensional Reynolds-averaged Navier-Stokes equations together with the two-layer algebraic eddy-viscosity model of Baldwin-Lomax modified with appropriate corrections for surface transpiration. The computed results show excellent agreement with the experimental data. The examination of these numerical results shows that the surface transpiration occurs at a low subsonic velocity and suggests that the effect of the transpiration through the porous surface on the interaction may be optimised.

Coherent anti-Stokes Raman spectroscopy temperature measurements in a hydrogen-fueled supersonic combustor

Abstract: Coherent anti-Stokes Raman spectroscopy (CARS) thermometry has been used to obtain static temperature cross sections in a three-dimensional supersonic combustor flowfield. Data were obtained in three spanwise planes downstream of a single normal fuel injector which was located downstream of a rearward-facing step. The freestream flow was nominally Mach 2 and was combustion heated to a total temperature of 1440 K (yielding a static temperature of about 800 K in the freestream) to simulate the inflow to a combustor operating at a flight Mach number of about 5.4. Since a broadband probe laser was used an instantaneous temperature sample was obtained with each laser shot at a repetition rate of 10 Hz. Thus root-mean-square (rms) temperatures and temperature probability density functions (pdf's) were obtained in addition to mean temperatures.

On the interaction between the shock wave attached to a wedge and freestream disturbances

Abstract: We present a study of the interaction of small amplitude, unsteady, freestream disturbances with a shock wave induced by a wedge in supersonic flow. These disturbances may be acoustic waves, vorticity waves, or entropy waves (or indeed a combination of all three). Their interactions then generate behind the shock disturbances of all three classes, an aspect that is investigated in some detail. Also, the possibility of enhanced mixing owing to additional vorticity produced by the shock-body coupling is investigated. It is shown that disturbances behind the shock may either decay downstream, or alternatively experience sustained oscillations. The precise regimes under which either behaviour is found are stated.

Airflow and Water-Drop Trajectories at Instrument Sampling Points around the Beechcraft King Air and Lockheed Electra

Abstract: Due to distortion of airflow streamlines, flow velocities and droplet size distributions measured around a moving aircraft can differ from freestream conditions. This can complicate measurements made from aircraft platforms. Potential flow calculations were used to predict airflow characteristics and the spatial distribution of different-sized droplets around the Lockheed Electra L-188 and Beechcraft King Air-200 aircraft at a variety of instrument mounting locations. Large deviations from freestream conditions were found to occur at certain locations on both aircraft near the fuselage and in regions of strong curvature. The number concentration of droplets 100–200 µm in diameter is most seriously affected by flow distortion effects. Calculation results were in reasonable agreement with measurements at a forward mounting location on the King Air.

Linear stability of the compressible reacting mixing layer

Abstract: We investigate the linear stability of mixing layers with special emphasis on the effects of heat release and compressibility. Multiple supersonic modes exist for both nonreacting and reacting flows when the disturbance phase velocity is supersonic relative to the freestream. These supersonic modes become less unstable with increasing Mach number but more unstable with increasing heat release. The most unstable supersonic modes are three-dimensional (oblique) for nonreacting flows but two-dimensional for reacting flows.

Computational analysis of off-design waveriders

Abstract: Comprehensive inviscid and viscous numerical simulations of hypersonic flow past nonconical rounded-nose waveriders are presented. The flow fields and aerodynamic forces at off-design conditions are determined inviscidly by a space-marching CFD code with the initial-data plane provided by a time-marching Navier-Stokes CFD code. Off-design conditions include off-design Mach numbers, angles of attack, and rounded leading edges. A wide range of waverider configurations is investigated and compared. These calculations show the effects of viscous interactions, which are influential near the leading edges, and determine the viscous drag. The inviscid calculations show that L/D decreases as freestream M increases (with alpha = 0). At the on-design Mach numbers, the maximum L/D may occur at slight positive or negative alpha, depending on the shape of the waverider, and zero lift occurs at a negative alpha approximately equal to half of the body thickness. The effects of slight leading-edge blunting produce only local effects in the flow field and small losses in L/D.

A numerical investigation of a subsonic jet in a crossflow

Abstract: A numerical analysis is presented of the flowfield induced by a single subsonic jet exhausting perpendicularly from a flat plate into a subsonic crossflow The analysis used available experimental data from a test case where the jet Mach number was 078 and the freestream Mach number was 013 Time-averaged solutions were obtained using the thin-layer Navier-Stokes equations and two overlapping grids The effect of turbulence model on the solutions was evaluated using two turbulence models: the zero-equation two-layer Baldwin-Lomax (1978) turbulence model and the one-equation Baldwin-Barth (1990) turbulence model It was found that, for some conditions, the zero-equation Baldwin-Lomax turbulence model gave better results than the one-equation Baldwin-Barth model or the laminar case

Inlet turbulence distortion and viscous flow development in a controlled-diffusion compressor cascade at very high incidence

Abstract: Detailed two-component laser Doppler velocimeter (LDV) measurements of the flow through a controlleddiffusion (CD) compressor cascade at a Reynolds number of about 700,000 and at a low Mach number are reported in this article. A very high-incidence angle (8 deg above design) was considered throughout this investigation, which included the full experimental characterization of the turbulence Held. The LDV measurements were fully automated and were all taken in coincidence mode, thus turbulent flow correlations could be determined. Most significant was the measurement of the distortion of the inlet freestream turbulence upstream of the blade leading edges. Such information is important in assessing viscous codes which incorporate transport equations to describe the turbulence within the flowfield. The laminar leading-edge separation bubble, which reattached turbulent, was enlarged on the suction surface of the blade. Consistent with measurements at lower incidence angles, the suction surface boundary layer remained attached over the rear part of the blade. The pressure side boundary layer initially showed little or no growth, however, it finally developed into a profile similar to a wall jet. The wake profiles showed significant asymmetry due to the high loading on the blades at the increased incidence angle.

Boundary-layer transition extent measurements on a cone and flat plate at Mach 3.5

Abstract: Wide excursions of the boundary-layer transition region are expected to occur on the X-30 National Aerospace Plane (NASP) due to the high Mach number, high temperature, and low density environment experienced during flight. Undesirable features of the transition region, such as the peak heat transfer rate, make it important to understand transition region physics. The current study investigates transition extent in 2D and axisymmetric bounday-layer flows. Surface-pitot and recovery temperature data obtained on a cone and flat plate at Mach 3.5 in the Supersonic Low-Disturbance Pilot Tunnel at NASA Langley are presented. Results show the effects of the unit Reynolds number, freestream disturbances, and nose/leading-edge bluntness on the extent of transition.

Evaluation of algebraic eddy-viscosity models in three-dimensional turbulent boundary-layer flows

Abstract: ~'vw/(dW/dy) N = anisotropy constant, ——//ar r /a \ uv/(dU/dy) £/ref__ _ = reference velocity wVv, 2 _ = kinematic normal stresses —7/v, — ww, — "vw = kinematic shear stresses W^o = surface velocity of rotating cylinder a = flow angle with respect to ATFS axis of local freestream coordinates ag = mean velocity gradient vector angle with respect to the #Fs * °f tne local freestream coordinates am = measured shear stress vector angle with respect to the #FS axis of the local freestream coordinates v kinematic viscosity

Transition correlation in subsonic flow over a flat plate

Abstract: A 2D subsonic flow over a flat plate with a freestream Mach number M(infinity) up to 0.8 is considered. The flow can experience continuous uniform suction through the wall and the wall can be heated or cooled continuously with a fixed wall temperature. For a specific combination of M(infinity), suction velocity, and level of heat transfer, the mean flow problem is solved and linear stability calculations are performed to compute the location on the flat plate where the factor representing the integration of growth rates reaches nine. These calculations are repeated for several combinations of flow parameters and the theoretically predicted transition location is presented in the form of a correlation that can account for the effect of wall suction, heat transfer, and Mach number.

CFD Code Calibration and Inlet-Fairing Effects on a 3D Hypersonic Powered-Simulation Model

Abstract: A three-dimensional (3D) computational study has been performed addressing issues related to the wind tunnel testing of a hypersonic powered-simulation model. The study consisted of three objectives. The first objective was to calibrate a state-of-the-art computational fluid dymamics (CFD) code in its ability to predict hypersonic powered-simulation flows by comparing CFD solutions with experimental surface pressure data. Aftbody lower surface pressures were well predicted, but lower surface wing pressures were less accurately predicted. The second objective was to determine the 3D effects on the aftbody created by fairing over the inlet; this was accomplished by comparing the CFD solutions of two closed-inlet powered configurations with a flowing-inlet powered configuration. Although results at four freestream Mach numbers indicate that the exhaust plume tends to isolate the aftbody surface from most forebody flowfield differences, a smooth inlet fairing provides the least aftbody force and moment variation compared to a flowing inlet. The final objective was to predict and understand the 3D characteristics of exhaust plume development at selected points on a representative flight path. Results showed a dramatic effect of plume expansion onto the wings as the freestream Mach number and corresponding nozzle pressure ratio are increased.

Crossflow instability in a spinning disk boundary layer

Abstract: Nomenclature D = diameter of a disk k = constant TV = spinning speed of a disk, rpm n = number of crossflow vortices Re = Reynolds number based on spinning speed r — distance along radius Tu = turbulence intensity U = velocity u = circumferential velocity in the boundary layer v = radial velocity in the boundary layer x = coordinate axis in freestream direction y = coordinate axis in azimuthal direction z = coordinate axis normal to the wall X = wavelength of disturbances X = crossflow parameter

DSMC simulations of Mach 20 nitrogen flows about a 70 degree blunted cone and its wake

Abstract: Numerical results obtained with the direct simulation Monte Carlo (DSMC) method are presented for Mach 20 nitrogen flow about a 70-deg blunted cone. The flow conditions simulated are those that can be obtained in existing low-density hypersonic wind tunnels. Three sets of flow conditions are simulated with freestream Knudsen numbers ranging from 0.03 to 0.001. The focus is to characterize the wake flow under rarefied conditions. This is accomplished by calculating the influence of rarefaction on wake structure along with the impact that an afterbody has on flow features. This data report presents extensive information concerning flowfield features and surface quantities.

Leading Edge Separation Bubbles on Turbomachine Blades

Abstract: Results are presented for separation bubbles of the type which can form near the leading edges of thin compressor or turbine blades. These often occur when the incidence is such that the stagnation point is not on the nose of the aerofoil. Tests were carried out at low speed on a single aerofoil to simulate the range of conditions found on compressor blades. Both circular and elliptic shapes of leading edge were tested. Results are presented for a range of incidence, Reynolds number and turbulence intensity and scale.The principal quantitative measurements presented are the pressure distributions in the leading edge and bubble region, as well as the boundary layer properties at a fixed distance downstream where most of the flows had reattached. Reynolds number was found to have a comparatively small influence, but a raised level of freestream turbulence has a striking effect, shortening or eliminating the bubble and increasing the magnitude of the suction spike. Increased freestream turbulence also reduces the boundary layer thickness and shape parameter after the bubble. Some explanations of the processes are outlined.© 1993 ASME

Near wake structure for a generic ASTV configuration

Abstract: Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using the direct simulation Monte Carlo method. Particular emphasis is given to the near wake flow and its sensitivity to rarefaction and other parametric variations. The flow conditions simulated are attainable in existing low-density hypersonic wind tunnels; that is, Mach 20 nitrogen flow encompassing freestream Knudsen numbers of 0.03 to 0.001. A stable vortex forms in the near wake at and below a freestream Knudsen number of 0.01 and the size of the vortex increases with decreasing freestream Knudsen number. The base region of the flow remains in thermal nonequilibrium for all cases. There is no formation of a lip separation shock or a distinct wake shock at these rarefied conditions.

Separated flow generated by a vectored jet in a crossflow

Abstract: Results of an experimental investigation to determine the characteristics of the separated flow region generated by a subsonic vectored rectangular jet in a subsonic crossflow is presented. Using surface visualization techniques, it was found that the upstream separated flow consist of horseshoe vortices that are formed periodically with a frequency corresponding to that of the vortex shedding behind the jet. The size of the recirculation region around the jet is found to decrease with increasing jet vector angle. The variation of the mean primary separation distance upstream of the jet, with velocity ratio revealed the existence of two different flow regimes. The change from one to the other depends on the velocity ratio and the jet vector angle. It is shown that the Strouhal number, based on the vortex shedding frequency and a combination of jet exit and freestream velocities, varies uniquely with velocity ratio.

Characteristics of Stretched Vortical Structures in Two-Dimensional Stagnation Flow

Abstract: Quantitative aspects on the relationship of the incoming freestream turbulence and the stretched vortical structures developed in a two-dimensional stagnation region are reported in this work The optimal time delay that resulted from the cross correlation of the velocity fluctuations measured in the stagnation region and the wall-pressure fluctuations is successfully normalized by the freestream velocity and the width of the bluff body, implying that this time delay is attributed to the effect of mean-flow convection The statistics regarding the unsteady stretched vortical structures detected at a fixed location in the stagnation region, in terms of the number of events per unit time and the averaged time duration of the flow structures detected, are successfully scaled by Lx and wo, where Lx and wo are the integral length scale and the intensity of grid-generate d turbulence, respectively Nomenclature