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At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

An experimental study of the properties of surface pressure fluctuations for separating turbulent boundary layers

Noise generated by helicopter and turbomachine rotors is a nuisance that designers would like to predict and to minimize within other design constraints. A key element for the noise calculation procedure is knowledge relating the flowfield structure to the surface pressure fluctuation structure. Surface pressure fluctuation data for zero-pressure-gradient and accelerating turbulent boundary layers were obtained. The zero-pressure-gradient turbulent boundary layers were examined with freestream velocities of 72 and 105 fps. Mean and fluctuation velocity profiles and streamwise velocity spectra and wavespeeds were obtained for momentum thickness Reynolds numbers up to 18000 for the zero-pressure-gradient case and up to 4000 for the favorable-pressure-gradient case. The wall shearing stress was estimated from a Clauser plot of the near wall data. It is clear that turbulent pressure fluctuations are produced by turbulent velocity fluctuations. Detailed simultaneous measurements of all of these fluctuations are needed to determine in more detail the structural relationships between velocity and pressure fields. Although some measurements were made for unseparated flows, none were made for separated flows.

Simpson, Roger L.↗
NASA NTRS1975

Recent developments in dynamic stall

Tests were performed on an airfoil in a low-turbulence subsonic wind tunnel at fixed incidence for different speeds of the wind, the airfoil being in translatory motion normal or parallel to the wind. The resulting aerodynamic forces and moment on the airfoil were measured along with local pressure and skin friction. When the incidence is small and the boundary layer attached, no unsteady effects were observed. However, for a large incidence (16 deg), in the case of boundary layer separation, strong unsteady effects are revealed when the boundary layer is reattached in periodic flow. Reattachment can be explained assuming that for the same value of free stream velocity, the gradient of the modulus of the periodic component of the relative airspeed at the upper surface of the airfoil in a direction normal to this surface is greater than the gradient of the relative air speed at the upper surface when the airfoil is not oscillating.

Mccroskey, W. J.↗
NASA NTRS1977

Supersonic separated turbulent boundary - layer over a wavy wall

A prediction method is developed for calculating distributions of surface heating rates, pressure and skin friction over a wavy wall in a two-dimensional supersonic flow. Of particular interest is the flow of thick turbulent boundary layers. The surface geometry and the flow conditions considered are such that there exists a strong interaction between the viscous and inviscid flow. First, using the interacting turbulent boundary layer equations, the problem is formulated in physical coordinates and then a reformulation of the governing equations in terms of Levy-Lees variables is given. Next, a numerical scheme for solving interacting boundary layer equations is adapted. A number of modifications which led to the improvement of the numerical algorithm are discussed. Finally, results are presented for flow over a train of up to six waves at various flow conditions.

Polak, A.↗
NASA NTRS1977

The separated turbulent boundary layer over a wavy wall

A study and application of the fourth order spline collocation procedure, numerical solution of boundary layer like differential equations, is presented. A simple inversion algorithm for the simultaneous solution of the resulting difference equations is given. Particular attention is focused on the boundary condition representation for the spline second derivative approximations. Solutions using the spline procedure, as well as the three point finite difference method, are presented for several model problems in order to assess and improve the spline numerical scheme. Application of the resulting algorithm to the incompressible laminar self similar boundary layer equations is presented.

Polak, A.↗
NASA NTRS1987

Characteristics of a separating confluent boundary layer and the downstream wake

Measurements of pressure and velocity characteristics are presented and analyzed for flow over and downstream of a NACA 4412 airfoil equipped with a NACA 4415 single-slotted flap at high angle of attack and close to maximum lift. The flow remained attached over the main element while a large region of recirculating flow occurred over the aft 61 percent of the flap. The airfoil configuration was tested at a Mach number of 0.09 and a chord Reynolds number of 1.8x10 to the 6th power in the NASA Ames Research Center 7- by 10-Foot Wind Tunnel. Measurement of mean and fluctuation velocities were obtained in regions of recirculation and high turbulence intensity using 3-D laser velocimetry. In regions where the flow had a preferred direction and relatively low turbulence intensity, hot-wire anemometry was used. Emphasis was placed on obtaining characteristics in the confluent boundary layer, the region of recirculating flow, and in the downstream wake. Surface pressure measurements were made on the main airfoil, flap, wind tunnel roof and floor. It is thought likely that because the model is large when compared to the wind tunnel cross section, the wind tunnel floor and ceiling interference should be taken into account when the flow field is calculated.

Adair, Desmond↗
NASA NTRS1961

Radiation Emission Effects of the Equilibrium Boundary Layer in the Stagnation Region

The laminar compressible boundary layer in chemical equilibrium is analysed to show the effect of emission of radiation by the boundary layer on solutions of the energy equation and on the resulting heat transfer. Solutions are obtained at one flight condition, but for several nose radii, in a regime where absorption is negligible. The consequent effects on heat transfer and boundary-layer thickness are determined. The concept of a separate boundary layer and shock layer is discussed in the light of the results obtained.

Howe, John Thomas↗
NASA NTRS1999

Aeroacoustic Resonance with Convergent-Divergent Nozzles

Convergent-divergent nozzles, when run at off-design conditions, often undergo flow resonance accompanied by the emission of a tone. Apart from screech occurring at higher operating pressures, resonance is also common at lower Mach numbers near transonic as well as subsonic conditions. With data from six nozzles of different size and design Mach number, the present paper documents the characteristics of the latter phenomenon that is morphologically quite different from conventional screech. The resonance is due to a feedback loop internal to the nozzle and is apparently driven by unsteady laminar boundary layer separation near the throat of the nozzle. Appropriate boundary layer tripping prior to the throat is found to eliminate or alter most of the tones. The Helmholtz number of the resonance, based on the throat-to-exit length, is found to attain a value of approximately 0.15 at M(sub j)=1 for all nozzles. However, its variation with M(sub j) may be different and depend on the nozzle geometry. With nozzles having larger throat-to-exit angle of divergence, the frequency is found to increase, in some cases having stage jumps to lower frequencies, with increasing operating pressure. With nozzles having smaller angle of divergence, the frequency variation exhibits an increase followed by a decrease involving one prominent stage occurring around transonic (M(sub j)= 1) condition. While the mechanisms remain far from completely clear, a model involving downstream propagating aerodynamic disturbance together with acoustic feedback explains the overall frequency characteristics for most cases.

Zaman, K. B. M. Q.↗
NASA NTRS2016

Changes in Flat Plate Wake Characteristics Obtained With Decreasing Plate Thickness

The near and very near wake of a flat plate with a circular trailing edge is investigated with data from direct numerical simulations. Computations were performed for four different Reynolds numbers based on plate thickness (D) and at constant plate length. The value of θ/D varies by a factor of approximately 20 in the computations (θ being the boundary layer momentum thickness at the trailing edge). The separating boundary layers are turbulent in all the cases. One objective of the study is to understand the changes in wake characteristics as the plate thickness is reduced (increasing θ/D). Vortex shedding is vigorous in the low θ/D cases with a substantial decrease in shedding intensity in the largest θ/D case (for all practical purposes shedding becomes almost intermittent). Other characteristics that are significantly altered with increasing θ/D are the roll-up of the detached shear layers and the magnitude of fluctuations in shedding period. These effects are explored in depth. The effects of changing θ/D on the distributions of the time-averaged, near-wake velocity statistics are discussed.

Wakes↗
NASA NTRS2016

Simulation of Sweep-Jet Flow Control, Single Jet and Full Vertical Tail

This work is a simulation technology demonstrator, of sweep jets used to suppress boundary layer separation and increase maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate stream wise direction. It also generate turbulent eddies at the oscillation frequency, which are typically large relative to boundary layer turbulence, and which augmenting mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from the oscillator, and the suppression of boundary layer separation by an array of sweep jets are performed. Simulation results are compared to data from a dedicated CFD validation experiment of a single oscillator and its sweep jet, and from a study of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets.2, 20 A critical step in the work is the development of realistic time-dependent sweep-jet in flow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the Over flow CFD solver, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

flow control↗
NASA NTRS2016

Simulation of Sweep-Jet Flow Control, Single Jet and Full Vertical Tail

This work is a simulation technology demonstrator, of sweep jet flow control used to suppress boundary layer separation and increase the maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate streamwise direction. It also generates turbulent eddies at the oscillation frequency, which are typically large relative to the scales of boundary layer turbulence, and which augment mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from a nozzle downstream of the oscillator, and an array of sweep jets which suppresses boundary layer separation are performed. Simulation results are compared to data from a dedicated validation experiment of a single oscillator and its sweep jet, and from a wind tunnel test of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets. A critical step in the work is the development of realistic time-dependent sweep jet inflow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the computational fluid dynamics (CFD) solver Overow, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

flow control↗
NASA NTRS1992

Measurements in a leading-edge separation bubble due to a simulated airfoil ice accretion

The separation bubble formed on an airfoil at low Reynolds number behind a simulated leading-edge glaze ice accretion is studied experimentally. Surface pressure and split hot-film measurements as well as flow visualization studies of the bubble reattachment point are reported. The simulated ice generates an adverse pressure gradient that causes a laminar separation bubble of the long bubble type to form. The boundary layer separates at a location on the ice accretion that is independent of angle of attack and reattaches at a downstream location 5-40 percent chord behind the leading edge, depending on the angle of attack. Velocity profiles show a large region of reverse flow that extends up from the airfoil surface as much as 2.5 percent chord. After reattachment, a thick distorted turbulent boundary layer exists. The separation bubble growth and reattachment are clearly seen in the plots of boundary-layer momentum thickness vs surface distance. Local minima and maxima in the boundary-layer momentum thickness development compare well with the shear layer transition point as indicated by the surface pressures and the reattachment point as measured from surface oil flow, respectively.

Bragg, M. B.↗
NASA NTRS1983

Assessment of NASA and RAE viscous-inviscid interaction methods for predicting transonic flow over nozzle afterbodies

The Langley Research Center of the National Aeronautics and Space Administration and the Royal Aircraft Establishment have undertaken a cooperative program to conduct an assessment of their patched viscous-inviscid interaction methods for predicting the transonic flow over nozzle afterbodies. The assessment was made by comparing the predictions of the two methods with experimental pressure distributions and boattail pressure drag for several convergent circular-arc nozzle configurations. Comparisons of the predictions of the two methods with the experimental data showed that both methods provided good predictions of the flow characteristics of nozzles with attached boundary layer flow. The RAE method also provided reasonable predictions of the pressure distributions and drag for the nozzles investigated that had separated boundary layers. The NASA method provided good predictions of the pressure distribution on separated flow nozzles that had relatively thin boundary layers. However, the NASA method was in poor agreement with experiment for separated nozzles with thick boundary layers due primarily to deficiencies in the method used to predict the separation location.

Putnam, L. E.↗
NASA NTRS1985

Direct numerical simulation of equilibrium turbulent boundary layers

This paper describes the simulation of turbulent boundary layers by direct numerical solution of the three-dimensional, time-dependent Navier-Stokes equations, using a spectral method. The flow is incompressible, with Re sub delta = 1000. The equations are written in the self-similar coordinate system and periodic streamwise and spanwise boundary conditions are imposed. A family of nine 'equilibrium' boundary layers, from the strongly accelerated 'sink' flow to Stratford's separating boundary layer is treated. Good general agreement with experiments is observed. The effects of pressure gradients on the structures and statistics, both in the wall and wake regions are discussed.

Spalart, P. R.↗
NASA NTRS1983

Experimental study of the separating confluent boundary-layer. Volume 2: Experimental data

An experimental low speed study of the separating confluent boundary layer on a NASA GAW-1 high lift airfoil is described. The airfoil was tested in a variety of high lift configurations comprised of leading edge slat and trailing edge flap combinations. The primary test instrumentation was a two dimensional laser velocimeter (LV) system operating in a backscatter mode. Surface pressures and corresponding LV derived boundary layer profiles are given in terms of velocity components, turbulence intensities and Reynolds shear stresses as characterizing confluent boundary layer behavior up to and beyond stall. LV derived profiles and associated boundary layer parameters and those obtained from more conventional instrumentation such as pitot static transverse, Preston tube measurements and hot-wire surveys are compared.

Braden, J. A.↗
NASA NTRS2018

Scale-Resolving Simulations of Low-Pressure Turbine Cascades with Wall Roughness Using A Spectral-Element Method

The accurate prediction of wall-roughness effects in turbomachinery is becoming critical as turbine designers address airfoil surface quality and degradation concerns arising from the shift to advanced ceramic matrix composite (CMC) or additively-manufactured airfoils operating in higher temperature environments. In this paper, a recently developed computational capability for accurate and efficient scale-resolving simulations of turbomachinery is extended to analyze the boundary- layer separation and transition characteristics in a rough-wall low-pressure turbine (LPT) cascade. The computational capability is based on an entropy-stable discontinuous-Galerkin spectral-element approach that extends to arbitrarily high orders of spatial and temporal accuracy, and is implemented in an efficient manner for a modern high performance computer architecture. Results from the scale-resolving simulations of both smooth and rough airfoil cascades are presented and compared to previous experiments and numerical simulations. The results show that the suction surface boundary layer undergoes laminar separation, transition, and turbulent reattachment for the smooth airfoil cascade, while in the presence of roughness the separation and transition behavior of the suction surface boundary layer is substantially modified. The differences between the smooth and rough airfoil cascades are then highlighted by a detailed analysis of their respective turbulent flow fields.

Spectral-Element↗
NASA NTRS1999

Static Performance of a Fixed-Geometry Exhaust Nozzle Incorporating Porous Cavities for Shock-Boundary Layer Interaction Control

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a fixed-geometry exhaust nozzle incorporating porous cavities for shock-boundary layer interaction control. Testing was conducted at static conditions using a sub-scale nozzle model with one baseline and 27 porous configurations. For the porous configurations, the effects of percent open porosity, hole diameter, and cavity depth were determined. All tests were conducted with no external flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable, shock-induced, boundary-layer separation at over-expanded conditions. Porous configurations were capable of controlling off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. The ability of the porous nozzle concept to alternately alleviate separation or encourage stable separation of exhaust flow through shock-boundary layer interaction control offers tremendous off-design performance benefits for fixed-geometry nozzle installations. In addition, the ability to encourage separation on one divergent flap while alleviating it on the other makes it possible to generate thrust vectoring using a fixed-geometry nozzle.

Asbury, Scott C.↗
NASA NTRS2014

Computational Aerodynamic Simulations of a 1484 ft/sec Tip Speed Quiet High-Speed Fan System Model for Acoustic Methods Assessment and Development

Computational Aerodynamic simulations of a 1484 ft/sec tip speed quiet high-speed fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, low-noise research fan/nacelle model that has undergone experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating points simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, which includes a core duct and a bypass duct that merge upstream of the fan system nozzle. As a result, only fan rotational speed and the system bypass ratio, set by means of a translating nozzle plug, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. Computed blade row flow fields at all fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive or critical boundary layer separations or related secondary-flow problems, with the exception of the hub boundary layer at the core duct entrance. At that location a significant flow separation is present. The region of local flow recirculation extends through a mixing plane, however, which for the particular mixing-plane model used is now known to exaggerate the recirculation. In any case, the flow separation has relatively little impact on the computed rotor and FEGV flow fields.

Tweedt, Daniel L.↗