Showing papers on "Freestream published in 1979"

Turbulent Shear Stresses in Compressible Boundary Layers

Abstract: Hot-wire anemometer measurements of turbulent shear stresses in a Mach 3 compressible boundary layer were performed in order to investigate the effects of heat transfer on turbulence. Measurements were obtained by an x-probe in a flat plate, zero pressure gradient, two dimensional boundary layer in a wind tunnel with wall to freestream temperature ratios of 0.94 and 0.71. The measured shear stress distributions are found to be in good agreement with previous results, supporting the contention that the shear stress distribution is essentially independent of Mach number and heat transfer for Mach numbers from incompressible to hypersonic and wall to freestream temperature ratios of 0.4 to 1.0. It is also found that corrections for frequency response limitations of the electronic equipment are necessary to determine the correct shear stress distribution, particularly at the walls.

Influence of Roughness and Blowing on Compressible Turbulent Boundary Layer Flow

Abstract: : The determination of the relative and combined effects of surface roughness and mass transfer on turbulent boundary-layer development, and in particular, on skin-friction drag was the prime objective of this study. Wind- tunnel tests were conducted in the NSWC Boundary-Layer Channel at a freestream Mach number of 2.9. The thick nozzle-wall boundary layer in the facility was subjected to a systematic variation of surface roughness and mass transfer conditions. Boundary-layer pressure and temperature surveys were obtained and skin friction was measured directly using a skin-friction balance which had a provision for active blowing through the floating drag element. Data comparisons with skin-friction theories and law-of-the-wall velocity profile correlations are presented.

Effect of Spanwise Blowing on the Aerodynamic Characteristics of the F-5E

Abstract: A 1/lO-scaIe F-5E was tested to investigate the spanwise blowing concept to provide improved aerodynamic characteristics with primary emphasis on high angle-of-attack performance and stability. Test data were obtained in the Northrop 7 X 10-ft low-speed facility at a freestream Mach number of 0.18 for a range of model angle of attack, sideslip, jet momentum coefficient, and leading- and trailing-edge flap deflection angles. Spanwise blowing on the 32 deg-swept wing from the wing/leading-edge extension (LEX) junction at a nozzle sweep angle of 55 deg resulted in LEX and wing leading-edge vortex enhancement and large vortex-induced lift increments and drag polar improvements at the higher angles of attack. Spanwise blowing improved the lateral/directional characteristics by delaying wing stall and maintaining vertical tail effectiveness to higher angles of attack. Deflecting the leading- and trailing-edge flaps down to 24 deg and 20 deg, respectively, delayed to higher model attitudes the more beneficial effects of blowing. Blowing reduces the specific excess power available for maneuvering the F-5E.

Hypersonic Viscous Flows over Sphere Cones at High Angles of Attack

Abstract: A PARABOL1ZED Navier-Stokes (PNS) prediction method developed to treat laminar flow over a sharp cone at incidence to a supersonic or hypersonic freestream has been extended to consider flow over a spherically blunted cone at angles of attack up to 38 deg. The basic method treats the flowfield between the body and the bow shock by one set of equations that are valid in both the viscous and inviscid regions. A description is given of the computational methods used, and results are compared with inviscid flowfield calculations and experimental data. The shock shape predictions were found to be in good agreement with most of the inviscid results as well as in reasonable agreement with some low Reynolds number heat transfer experimental data.

Axial Compressor Blade Optimization in the Low Reynolds Number Regime

Abstract: Original and previously published experimental data are presented in this study of axial compressor blade optimization for use in the subcritical Reynolds number region. Three methods of increasing cascade performance at these low Rc are presented, discussed, and compared: 1) the use of sharp, leading-edge profiles, 2) trip wires or roughness elements, and 3) high freestream turbulence. The original data were collected in the VKI C-l cascade wind tunnel; during these tests, blade chord Reynolds number was varied from 250,000 down to -17,000. Each technique demonstrated an increase in low Rc performance for certain profiles, with the sharp, leading-edge profile as the potential optimum. Axial compressors or pumps that must operate for a significant amount of their design life under low Reynolds number conditions might benefit substantially from these methods.

Wake Vortex Measurements of an Ogive Cylinder at CY = 36 Degrees

Abstract: Three-dimensional laser velocimeter measurements have been made of the wake vortices of a slender tangentogive body which had nose and body fineness ratios of 3.5 and 12, respectively. Data were obtained at an angle of attack of 36 deg, a freestream Mach number of 0.6, and a Reynolds number based upon body diameter of 170,000. Details of the mean flowfield are presented, and features of the turbulent and unsteady nature of the vortex flowfield are discussed. Problems associated with obtaining meaningful vortex measurements in highspeed flows are addressed.

New Rotation-Balance Apparatus for Measuring Airplane Spin Aerodynamics in the Wind Tunnel

Abstract: An advanced rotation-balance apparatus has been developed for the Ames 12-ft Pressure Wind Tunnel to study the effects of spin rate, angles of attack and sideslip, and, particularly, Reynolds number on the aerodynamics of fighter and general aviation aircraft in a steady spin. Angles of attack to 100 deg and angles of sideslip to 30 deg are possible with spin rates to 42 rad/s (400 rpm) and Reynolds numbers to 30x 106/m on fighter models with wing spans that are typically 0.7 m. A complete description of the new rotation-balance apparatus, the sting/balance/model assembly, and the operational capabilities is given. Nomenclature A = body reference area, ird214 b = wing span (0.457 m for Figs. 3-6) Cy = body side force/qA d = diameter of centerbody (0.0762 m for Figs. 3-6) M = freestream Mach number q — freestream dynamic pressure Rd — Reynolds number based on d V — freestream velocity as — sting angle (Figs. 11 and 12) a =angle of attack /3 = angle of sideslip a = angle between freestream velocity vector and body longitudinal axis

Pressure Gradient Effects on Supersonic Boundary Layer Turbulence.

Abstract: : Measurements of mean flow profiles at several streamwise locations in a supersonic turbulent boundary layer growing under a continuous adverse pressure gradient are reported. Tests were performed at a freestream Mach number of 3, for an adiabatic wall, using two curved ramps designed to produce constant pressure gradient flows. The velocity profile data, when transformed to incompressible coordinates, are in good agreement with Coles universal 'wall-wake' velocity profile and they indicate that the boundary layer is in local equilibrium and essentially independent of upstream history. In addition, the Coles wake parameters and Clauser shape factors, characterizing the transformed profiles, are in accord with the results of low speed correlations of adverse pressure gradient flows. The turbulent transport terms were extracted from the mean flow field data and indicate that for a given ramp, the profile of turbulent shear stress normalized by the wall shear, versus distance from the surface, normalized by the local boundary thickness, is severely distored by the pressure gradient although it is apparently insensitive to local conditions.

The Transonic Integral Equation Method With Curved Shock Waves

Abstract: In the existing solutions of the extended integral equation method for transonic flows the shock wave has been assumed normal to the freestream. In this paper the effect of removing this assumption is investigated and flows with curved shock waves are examined.

Performance of a V/STOL tilt nacelle inlet with blowing boundary layer control

Abstract: A scale model of a V/STOL tilt nacelle fitted to a 0.508 m single stage fan was tested in the NASA Lewis 9x15 ft low speed wind tunnel to determine the effect of diffuser blowing on the inlet aerodynamics and aeromechanical performance. The test was conducted over a range of freestream speeds (up to 120 knots) and angles of attack (up to 120 deg). Diffuser blowing had a beneficial affect on all performance parameters. The angle of attack range for separation free flow substantially increased, and the fan face distortion significantly reduced with a corresponding increase in total pressure recovery. Discrete narrow band blade stress peaks which were common to the nonblowing (baseline) configuration were eradicated with diffuser blowing.

Convective heat flux in a laser-heated thruster

Abstract: An analysis is performed to estimate the convective heating to the wall in a laser-heated thruster on the basis of a solution of the laminar boundary-layer equations with variable transport properties. A local similiarity approximation is used, and it is assumed that the gas phase is in equilibrium. For the thruster described by Wu (1976), the temperature and pressure distributions along the nozzle are obtained from the core calculation. The similarity solutions and heat flux are obtained from the freestream conditions of the boundary layer, in order to determine if it is necessary to couple the boundary losses directly to the core calculation. In addition, the effects of mass injection on the convective heat transfer across the boundary layer with large density-viscosity product gradient are examined.

Effect of spanwise blowing on the aerodynamic characteristics of the F-5E

Abstract: A 1/lO-scaIe F-5E was tested to investigate the spanwise blowing concept to provide improved aerodynamic characteristics with primary emphasis on high angle-of-attack performance and stability. Test data were obtained in the Northrop 7 X 10-ft low-speed facility at a freestream Mach number of 0.18 for a range of model angle of attack, sideslip, jet momentum coefficient, and leading- and trailing-edge flap deflection angles. Spanwise blowing on the 32 deg-swept wing from the wing/leading-edge extension (LEX) junction at a nozzle sweep angle of 55 deg resulted in LEX and wing leading-edge vortex enhancement and large vortex-induced lift increments and drag polar improvements at the higher angles of attack. Spanwise blowing improved the lateral/directional characteristics by delaying wing stall and maintaining vertical tail effectiveness to higher angles of attack. Deflecting the leading- and trailing-edge flaps down to 24 deg and 20 deg, respectively, delayed to higher model attitudes the more beneficial effects of blowing. Blowing reduces the specific excess power available for maneuvering the F-5E.

Total Pressure Recovery of Flared Fan Nozzles Used as Inlets

Abstract: A simple expression has been developed which accurately represents the total pressure recovery of a flared exhaust nozzle when it is used as an inlet. Physically such a situation can arise during the reverse-thrust operation of a variable pitch fan jet. The formula is written in terms of the freestream and duct Mach numbers and contains two empirical coefficients. One coefficient is associated with internal flow losses in the nozzle and can be found from a static engine test. The other coefficient is associated with flow losses in the external field of the nozzle and can be approximated from existing cone drag data. The developed expression produces an excellent fit of measured total pressure recoveries; differences between calculated and experimental values were generally less than 0.5%.

An Experimental Study of a Jet Issuing From a Lifting Wing

Abstract: An experimental program was conducted to determine the behavior of a round turbulent jet issuing from a lifting two-dimensional wing in crossflow The jet was located at 65% wing chord on an NACA 0021 airfoil fitted with a 30% chord NACA 4415 flap The flowfield associated with the jet was surveyed extensively with directional pressure probes to determine local velocity vectors and pressures for three different values of lift coefficient at jet effective velocity ratios (square root of the ratio of the jet dynamic pressure to the freestream dynamic pressure) of 4, 6, and 8 Data describing the jet centerline and the path of the contrarotating vortices accompanying the deflected jet are presented and compared with similar data for a round jet issuing from a large flat plate The spacing and strength of the vortices are calculated using a simple vortex model previously proposed for the flat plate case The results show that the penetration of the jet and the vortices increases significantly with increasing lift for the range of test parameters covered in the study The calculated vortex spacing and strength also show an increase with lift

Experimental determination of probe-length requirements for studies of the turbulent wake behind a cylinder.

Abstract: The attenuation of turbulence and mean velocity signals due to the line averaging imposed by hot wires when used in the wake of an isolated circular cylinder has been investigated in a wind tunnel by measurements using several choices of hot-wire length, cylinder diameter, and freestream mean velocity. The results are presented graphically in order to provide a practical method for determining attenuation of the turbulence and mean velocity signals obtained in a wake. The length scale of the wake can be defined as L=0.6[(x-x(o)) d](1/2), where x is the downstream distance from the cylinder, d is the cylinder diameter, and x(o)=25d. For all the wires tested, the attenuation of the measured turbulence signal is limited to within 5% only if the wire length is smaller than 0.1 L. For a wire normal to the cylinder and cross wind, the attenuation of the signal of the mean velocity-defect factor, expressed as (1-u/u(infinity)), where u and u(infinity) are local and free-stream velocities, respectively, is less than 5% only if the wire is less than 0.5 L in length.

Significance of shock and body slip conditions on Jovian entry heating

Abstract: The influence of the body and shock slip conditions on the heating of a Jovian entry body is investigated. The flow in the shock layer is considered to be axisymmetric, steady, laminar, viscous, and in chemical equilibrium. Realistic thermophysical and step-function spectral models are employed and results are obtained by implicit finite-difference and iteractive procedures. The freestream conditions correspond to a typical Jovian entry trajectory point. The results indicate that the effect of the slip conditions is significant when the altitudes are higher than 225 km and that the contribution of a radiative heat-flux term in the energy equation should not be neglected at any altitude.

Technique for Vorticity Measurement in Unsteady Flow

Abstract: A technique is described for determining spatial vorticity distributions in nonstationary fluid flows exhibiting a periodic, time-varying, mean velocity field. The method requires the integration of velocity field data about a spatial array of closed contours to infer the distribution of vorticity using the generalized definition of circulation. A digital data acquisition scheme is suggested for the handling and processing of large quantities of data encountered in typical applications. The method was used to determine the flow characteristics in portions of an unsteady separated region generated by an oscillating spoiler on an airfoil surface. Several data display alternatives are discussed. ORTICITY is a significant physical variable in many real fluid flows, but it is a parameter that is often difficult to determine experimentally. The method to be described provides a means of measuring time-varying mean vorticity components and their distributions in nonstationary flows exhibiting a periodic mean velocity field. Through the years, a number of vorticity measurement techniques have evolved which have been successful to a greater or lesser degree. No single method, however, has proved satisfactory over the entire range of possible flow situations. Attempts to infer the streamwise vorticity component from velocity field measurements date back to the early part of the century.l Limitations on velocity measurements techniques at that time did not permit its application to time-varying fluid motions. A successful method for detecting the freestream-oriented vorticity component was discussed by Kovasznay2 and Kistler.3 A special purpose probe was constructed using an array of four hot-wire sensors of equal length and oriented at the same angle with respect to the freestream direction. The four sensors were operated as elements of a single constant current bridge circuit. In addition to the limitation on its directional capability, the wire array size must be of the order of the flow of Kolmogorov microscale for accurate spectral response.4 Vane-type mechanical vorticity meters have recently become popular, but these devices are also limited to measurement of the mean vorticity component in the freestream direction. The first use of these probes was discussed by May.5 The rotation frequency of the spinner and shaft is converted to an electrical signal using light from a bulb on one side of the shaft. A photocell on the opposite side converts this rotation to a train of pulses whose frequency is a function of the streamwise vorticity component. The resulting spatial resolution is limited by the vane array geometry, while accuracy and linearity are limited by friction in the shaft bearings. Effects of inertia and friction in unsteady flow applications have not yet been documented.

Unsteady thin airfoil theory for transonic flows with embedded shocks

Abstract: Classical unsteady thin airfoil theory fails for low frequencies at the subsonic freestream Mach number, because of the formation of a shock wave that shields the forward region of the airfoil from aft generated disturbances. In the present paper, the classical thin airfoil theory is modified to account for the presence and induced motion of such shocks. The modification consists of taking the steady local Mach number to be a simple step discontinuity, normal to the undisturbed flow, separating two uniform regions. Predicted regions are shown to correlate well both with the experiment and finite difference calculations.

Axisymmetric nonstationary flow around an obstable in a cylindrical tube

Abstract: Questions associated with the interaction between a gas stream and a body in a launching tube, especially the high-speed propulsion of a body by a gas stream [1], have become of great interest in recent years. Partial destruction of the body and the formation of a gap between the body and the launching tube, through which the working gas will flow, inevitably occurs at high velocities. In this case it is possible to consider the ejection of a free body which does not come into contact with the walls of the launching tube as it is accelerated. An analogous problem occurs in the transportation of containers in a tube under the effect of a compressed gas [2], as well as in a gas-dynamic analysis of piston apparatus with different kinds of gas flow through the orifice inmoving or fixed pistons. The interaction between the gas stream and the body or the obstacle in the launching tube must be known for a theoretical investigation of all these problems. The solution is obtained by numerical integration of the nonstationary gas-dynamic equations by means of a through-computation difference scheme [3]. Values of the blockage factors are found for different freestream Mach numbers, for which the reflected shock stands off at infinity upstream. A comparison is given with the one-dimensional approximation obtained under the assumption that the body being streamlined is replaced by two jumps of a strong discontinuity on which the mass, momentum, and energy conservation conditions are satisfied. The results obtained are used in the problem of ejection of a free body from a launching tube under the effect of an unsteady gas flow.

Nonequilibrium viscous flow over Jovian entry probes at high altitudes

Abstract: The viscous chemical nonequilibrium flow around a Jovian entry body is investigated at high altitudes using two different methods. First method is only for the stagnation region and integrates the full Navier-Stokes equations from the body surface to the freestream. The second method uses viscous shock layer equations between the body surface and the shock. Due to low Reynolds numbers, both methods use surface slip boundary conditions and the second method also uses shock slip boundary conditions. The results of the two methods are compared at the stagnation point. It is found that the entire shock layer is under chemical nonequilibrium at higher altitudes and that the slip boundary conditions are important at these altitudes.

Nonequilibrium viscous flow and convective heating of a jovian entry probe at high altitudes

Abstract: The viscous chemical nonequilibrium flow around a Jovian entry body is investigated at high altitudes using two different methods. The first method is only for the stagnation region and integrates the full Navier-Stokes equations from the body surface to the freestream. The second method uses viscous shock-layer equations between the body surface and the shock. Due to low Reynolds numbers, both methods use surfaceslip boundary conditions, and the second method also uses shock-slip boundary conditions. The results of the two methods are compared at the stagnation point. It is found that the entire shock layer is under chemical nonequilibrium at higher altitudes and that the slip boundary conditions are important at these altitudes. Nomenclature C-j = mass fraction of ith species h = nondimensional enthalpy, F/v£ _ _p h. = nondimensional enthalpy of ith species, h-/V ' _ _n ' °° H = nondimensional total enthalpy, J. = mass diffusion flux of species i