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At least 1,135 records · Page 63

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NASA NTRS1960

Effects of Sweep Angle on the Boundary-Layer Stability Characteristics of an Untapered Wing at Low Speeds

An investigation was conducted in the Ames 12-Foot Low-Turbulence Pressure Tunnel to determine the effects of sweep on the boundary-layer stability characteristics of an untapered variable-sweep wing having an NACA 64(2)A015 section normal to the leading edge. Pressure distribution and transition were measured on the wing at low speeds at sweep angles of 0, 10, 20, 30, 40, and 50 deg. and at angles of attack from -3 to 3 deg. The investigation also included flow-visualization studies on the surface at sweep angles from 0 to 50 deg. and total pressure surveys in the boundary layer at a sweep angle of 30 deg. for angles of attack from -12 to 0 deg. It was found that sweep caused premature transition on the wing under certain conditions. This effect resulted from the formation of vortices in the boundary layer when a critical combination of sweep angle, pressure gradient, and stream Reynolds number was attained. A useful parameter in indicating the combined effect of these flow variables on vortex formation and on beginning transition is the crossflow Reynolds number. The critical values of crossflow Reynolds number for vortex formation found in this investigation range from about 135 to 190 and are in good agreement with those reported in previous investigations. The values of crossflow Reynolds number for beginning transitions were found to be between 190 and 260. For each condition (i.e., development of vortices and initiation of transition at a given location) the lower values in the specified ranges were obtained with a light coating of flow-visualization material on the surface. A method is presented for the rapid computation of crossflow Reynolds number on any swept surface for which the pressure distribution is known. From calculations based on this method, it was found that the maximum values of crossflow Reynolds number are attained under conditions of a strong pressure gradient and at a sweep angle of about 50 deg. Due to the primary dependence on pressure gradient, effects of sweep in causing premature transition are generally first encountered on the lower surfaces of wings operating at positive angles of attack.

Boltz, Frederick W.↗

NASA NTRS1989

Unsteady blade pressure measurements on a model counterrotation propeller

In an exploratory effort an advanced counterrotation propeller instrumented with blade-mounted pressure transducers was tested in the NASA Lewis 9- by 15-Foot Anechoic Wind Tunnel at a simulated takeoff and landing speed of Mach 0.20. The propeller's aft diameter was reduced to investigate possible noise reductions resulting from reduced blade row interaction with the tip vortex. The propeller was tested at three blade row spacings at fixed blade setting angles, at the maximum blade row spacing at higher blade setting angles and at propeller axis angles attack to the flow up to + or - 16 deg. A limited number of unsteady blade surface pressure measurements were made on both rotors of the model counterrotation propeller. Emphasis was placed on determining the effects of rotor-rotor interactions on the blade surface pressures. A unique method of processing the pressure signals was developed that enables even weak interaction waveforms and spectra to be separated from the total signal. The interaction on the aft rotor was many times stronger than that on the forward rotor. The fundamental rotor interaction tone exhibited complicated behavior but generally increased with rotational speed and blade setting angle and decreased with rotor spacing. With the propeller axis at an angle to the flow, the phase response of the aft rotor appeared to be significantly affected by the presence of the forward rotor.

Heidelberg, Laurence J.↗

NASA NTRS1989

Noise produced by turbulent flow into a rotor: Theory manual for atmospheric turbulence prediction and mean flow and turbulence contraction prediction

Prediction of helicopter main rotor noise due to ingestion of atmospheric turbulence was analyzed. The analysis combines several different models that describe the fluid mechanics of the turbulence and the ingestion process. Two models, atmospheric turbulence, and mean flow and turbulence contraction were covered. The third model, covered in a separate report, describes the rotor acoustic mode. The method incorporates the atmospheric turbulence model and a rapid distortion turbulence contraction description to determine the statistics of the anisotropic turbulence at the rotor plane. The analytical basis for a module was provided which was incorporated in NASA's ROTONET helicopter noise prediction program. The mean flow and turbulence statistics associated with the atmospheric boundary layer were modeled including effects of atmospheric stability length, wind speed, and altitude. The turbulence distortion process is modeled as a deformation of vortex filaments (which represent the turbulence field) by a mean flow field due to the rotor inflow.

Simonich, J. C.↗

NASA NTRS1989

Unsteady blade pressure measurements on a model counterrotation propeller

Heidelberg, Laurence J.↗

NASA NTRS1990

Numerical simulation of the viscous flow around a simplified F/A-18 at high angles of attack

A numerical method developed for solving the viscous flow around three-dimensional complex configurations is presently used to simulate the flow around a simplified F/A-18 configuration encompassing forebody, wing, leading-edge extension, faired-over inlet, and deflected wing leading-edge flaps, at Mach 0.243 and 30.3 deg angle of attack. The computational results show the details of the flowfield structure, including primary, secondary, and tertiary separation lines, the development of forebody and leading-edge extension vortex, and the burst of this vortex. A grid-refinement study is conducted to assess the effect of grid characteristics on solution accuracy. Substantial agreement is obtained between these results and flight data.

Rizk, Yehia M.↗

NASA NTRS2014

Gradient Calculation Methods on Arbitrary Polyhedral Unstructured Meshes for Cell-Centered CFD Solvers

A survey of gradient reconstruction methods for cell-centered data on unstructured meshes is conducted within the scope of accuracy assessment. Formal order of accuracy, as well as error magnitudes for each of the studied methods, are evaluated on a complex mesh of various cell types through consecutive local scaling of an analytical test function. The tests highlighted several gradient operator choices that can consistently achieve 1st order accuracy regardless of cell type and shape. The tests further offered error comparisons for given cell types, leading to the observation that the "ideal" gradient operator choice is not universal. Practical implications of the results are explored via CFD solutions of a 2D inviscid standing vortex, portraying the discretization error properties. A relatively naive, yet largely unexplored, approach of local curvilinear stencil transformation exhibited surprisingly favorable properties

Meshes↗

NASA NTRS1986

Laser velocimetry in highly three-dimensional and vortical flows

The need for experimentally determined 3-D velocity information is crucial to the understanding of highly 3-dimensional and vortical flow fields. In addition to gaining an understanding of the physics of flow fields, a correlation of velocity data is needed for advanced computational modelling. A double pass method for acquiring 3-D flow field information using a 2-D laser velocimeter (LV) is described. The design and implementation of a 3-D LV with expanded capabilities to acquire real-time 3-D flow field information are also described. Finally, the use of such an instrument in a wind tunnel study of a generic fighter configuration is described. The results of the wind tunnel study highlight the complexities of 3-D flow fields, particularly when the vortex behavior is examined over a range of angles of attack.

Novak, C. J.↗

NASA NTRS1992

Fluid-flow of a row of jets in crossflow - A numerical study

A detailed computer-visualized flow field of a row of jets in a confined crossflow is presented. The Reynolds averaged Navier-Stokes equations are solved using a finite volume method that incorporates a partial differential equation for incremental pressure to obtain a divergence-free flow field. The turbulence is described by a multiple-time-scale turbulence model. The computational domain includes the upstream region of the circular jet so that the interaction between the jet and the crossflow is simulated accurately. It is shown that the row of jets in the crossflow is characterized by a highly complex flow field that includes a horse-shoe vortex and two helical vortices whose secondary velocity components are co-rotating in space. It is also shown that the horse-shoe vortex is a ring of reversed flows located along the circumference of the jet exit.

Kim, S.-W.↗

NASA NTRS1993

Simulation of tail buffet using delta wing-vertical tail configuration

Computational simulation of the vertical tail buffet problem is accomplished using a delta wing-vertical tail configuration. Flow conditions are selected such that the wing primary-vortex cores experience vortex breakdown and the resulting flow interacts with the vertical tail. This multidisciplinary problem is solved successively using three sets of equations for the fluid flow, aeroelastic deflections and grid displacements. For the fluid dynamics part, the unsteady, compressible, full Navier-Stokes equations are solved accurately in time using an implicit, upwind, flux-difference splitting, finite-volume scheme. For the aeroelastic part, the aeroelastic equation for bending vibrations is solved accurately in time using the Galerkin method and the four-stage Runge-Kutta scheme. The grid for the fluid dynamics computations is updated every few time steps using a third set of interpolation equations. The computational application includes a delta wing of aspect ratio 1 and a rectangular vertical tail of aspect ratio 2, which is placed at 0.5 root-chord length downstream of the wing trailing edge. The wing angle of attack is 35 deg and the flow Mach number and Reynolds number are 0.4 and 10,000, respectively.

Kandil, Osama A.↗

NASA NTRS2006

Unsteady Flow Simulation of High-speed Turbopumps

Computation of high-speed hydrodynamics requires high-fidelity simulation to resolve flow features involving transient flow, cavitation, tip vortex and multiple scales of unsteady fluctuations. One example of this type in aerospace is related to liquid-fueled rocket turbopump. Rocket turbopumps operate under severe conditions at very high rotational speeds typically at thousands of rpm. For example, the Shuttle orbiter low-pressure-fuel-turbopump creates transient flow features associated with reverse flows, tip clearance effects, secondary flows, vortex shedding, junction flows, and cavitation effects. Flow unsteadiness originating from the orbiter Low-Pressure-Fuel-Turbopump (LPFTP) inducer is one of the major contributors to the high frequency cyclic loading that results in high cycle fatigue damage to the flow liners just upstream of the LPFTP. The reverse flow generated at the tip of the inducer blades travels upstream and interacts with the bellows cavity. Simulation procedure for this type high-speed hydrodynamic problems requires a method for quantifying multi-scale and multi-phase flow as well as an efficient high-end computing strategy. The current paper presents a high-fidelity computational procedure for unsteady hydrodynamic problems using a high-speed liquid-fueled rocket turbopump.

Kiris, Cetin C.↗

NASA NTRS1983

Computation of supersonic viscous flows around pointed bodies at large incidence

A recently reported parabolized Navier-Stokes method has been extended to compute turbulent supersonic flows around cones and an ogive-cylinder body at large incidence. The algebraic eddy-viscosity turbulence model contained in the code was modified to properly account for the large regions of cross-flow separation that occur in these flows. Extensive comparisons between computed results and experimentally measured flow fields are presented. The results show good agreement for viscous-layer profiles and details of the external leeward-side vortex structure at angles of attack up to three times the cone half angles. Details of the modified turbulence model are presented and discussed.

Degani, D.↗

NASA NTRS1985

Some design philosophy for reducing the community noise of advanced counter-rotation propellers

Advanced counter-rotation propellers have been indicated as possibly generating an unacceptable amount of noise for the people living near an airport. This report has explored ways to reduce this noise level, which is treated as being caused by the interaction of the upstream propeller wakes and vortices with the downstream propeller. The noise reduction techniques fall into two categories: (1) reducing the strength of the wakes and vortices, and (2) reducing the response of the downstream blades to them. The noise from the wake interaction was indicated as being reduced by increased propeller spacing and decreased blade drag coefficient. The vortex-interaction noise could be eliminated by having the vortex pass over the tips of the downstream blade, and it could be reduced by increased spacing or decreased initial circulation. The downstream blade response could be lessened by increasing the reduced frequency parameter omega or by phasing of the response from different sections to have a mutual cancellation effect. Uneven blade to blade spacing for the downstream blading was indicated as having a possible effect on the annoyance of counter-rotation propeller noise. Although there are undoubtedly additional methods of noise reduction not covered in this report, the inclusion of the design methods discussed would potentially result in a counter-rotation propeller that is acceptably quiet.

Dittmar, J. H.↗

NASA NTRS1987

Severe winds in the DFW microburst measured from two aircraft

Advanced analytical methods are applied to the digital flight data and Air Traffic Control radar records from two airliners, Delta flight 191 and American flight 539, that penetrated a microburst at the Dallas/Ft. Worth airport on August 2, 1985. The results for Delta 191 (which crashed) show that the aircraft encountered a strong downburst followed by a strong outflow accompanied by large and rapid changes in the vertical wind. The rapid changes in the vertical wind detected near the ground are attributed to vortex-induced turbulence. The next following aircraft, American 539, made a go-around 110 sec after Delta 191 and traversed the microburst at an altitude of about 2500 ft above the ground. The measured winds during this go-around indicate a broad pattern of downflow in the microburst, with regions of upflow at the extreme edges. The combined results indicate a microburst that was increasing in size with vortex-induced turbulence embedded in a strong outflow near the ground. These wind measurements provide a realistic description of the microburst wind environment that can be used in simulator studies and pilot training, and provide a standard of comparison for ongoing experiments and modeling of microbursts involving vortex rings.

Wingrove, R. C.↗

NASA NTRS1992

Effect of planform and body on supersonic aerodynamics of multibody configurations

An experimental and theoretical investigation of the effect of the wing planform and bodies on the supersonic aerodynamics of a low-fineness-ratio, multibody configuration has been conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Force and moment data, flow-visualization data, and surface-pressure data were obtained on eight low-fineness-ratio, twin-body configurations. These configurations varied in inboard wing planform shape, outboard wing planform shape, outboard wing planform size, and presence of the bodies. The force and moment data showed that increasing the ratio of outboard wing area to total wing area or increasing the leading-edge sweep of the inboard wing influenced the aerodynamic characteristics. The flow-visualization data showed a complex flow-field system of shocks, shock-induced separation, and body vortex systems occurring between the side bodies. This flow field was substantially affected by the inboard wing planform shape but minimally affected by the outboard wing planform shape. The flow-visualization and surface-pressure data showed that flow over the outboard wing developed as expected with changes in angle of attack and Mach number and was affected by the leading-edge sweep of the inboard wing and the presence of the bodies. Evaluation of the linear-theory prediction methods revealed their general inability to consistently predict the characteristics of these multibody configurations.

Mcmillin, S. Naomi↗

NASA NTRS1977

A numerical method for the exact calculation of airloads associated with impulsively started wings

A numerical method is developed to calculate three-dimensional potential flows due to the steady and impulsive motion of isolated wing and wing-wing interaction problems. The velocity potential is represented by a discrete set of constant-doublet quadrilaterals on wing and wake surfaces. The exact surface boundary condition is enforced, and the solution is obtained in a step-by-step fashion, configurations being impulsively started from rest. Free-wake geometries are generated for each time step with Rankine or Lamb viscous vortex segments used in wake-velocity calculations. Sample results include calculated performance to steady state for a thick wing and indicial lift of a wing-wing interaction problem.

Summa, J. M.↗

NASA NTRS1980

Fast Generation of body conforming grids for 3-D

A fast algorithm was developed for accurately generating boundary conforming, three dimensional, consecutively refined, computational grids applicable to arbitrary axial turbomachinery geometry. The method is based on using a single analytic function to generate two dimensional grids on a number of coaxial axisymmetric surfaces positioned between the hub and the shroud. These grids are of the "O" type and are characterized by quasi-orthogonality, geometric periodicity, and an adequate resolution throughout the flowfield. Due to the built in additional nonorthogonal coordinate stretching and shearing, the grid lines leaving the trailing of the blade end at downstream infinity, thus simplifying the numerical treatment of the three dimensional trailing vortex sheet.

Dulikravich, O.↗

NASA NTRS1986

Method for the prediction of the installation aerodynamics of a propfan at subsonic speeds: User manual

This document is the user's guide for the method developed earlier for predicting the slipstream wing interaction at subsonic speeds. The analysis involves a subsonic panel code (HESS code) modified to handle the propeller onset flow. The propfan slipstream effects are superimposed on the normal flow boundary condition and are applied over the surface washed by the slipstream. The effects of the propeller slipstream are to increase the axial induced velocity, tangential velocity, and a total pressure rise in the wake of the propeller. Principles based on blade performance theory, momentum theory, and vortex theory were used to evaluate the slipstream effects. The code can be applied to any arbitrary three dimensional geometry, expressed in the form of HESS input format. The code can handle a propeller alone configuration or a propeller/nacelle/airframe configuration, operating up to high subcritical Mach numbers over a range of angles of attack. Inclusion of a viscous modelling is briefly outlined. Wind tunnel results/theory comparisons are included as examples for the application of the code to a generic supercritical wing/overwing Nacelle with a powered propfan. A sample input/output listing is provided.

Chandrasekaran, B.↗

NASA NTRS1992

Supersonic jet mixing enhancement by delta-tabs

The results of a continuing investigation of the effect of vortex generators, in the form of small tabs at the nozzle exit, on the evolution of a jet are reported. Primarily, tabs of triangular shape are considered, and the effect is studied up to an equivalent jet Mach number of 1.8. By changing the orientation of the tab with respect to the nozzle exit plane, streamwise vortex pairs of opposite sign were generated. This resulted in either an outward election of jet core fluid into the ambient or an inward indentation of the mixing layer into the core of the jet. A triangular shaped tab with its apex leaning downstream, referred to as a delta tab, was found to be the most effective in influencing the jet evolution. Two delta tabs, spaced 180 degrees apart, completely bifurcated the jet. Four delta tabs increased jet mixing substantially, more than by various other methods tried previously; the mass flux at fourteen jet diameters downstream from the nozzle increased by about 50 percent over that for the no tab case. The tabs were found to be effective in jets with laminar or turbulent boundary layers as well as in jets with low or high core turbulence intensities.

Zaman, K. B. M. Q.↗