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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Viscous-inviscid interaction and local grid refinement via truncation error injection

A methodology is presented which makes it possible to decouple a complex problem having multiple disparate length scales into problems of single length scale so that they can be solved more efficiently on a computer. The method is applied to a viscous transonic flow over an airfoil. It is found that accurate prediction of the flow over an airfoil can be obtained by solving the Euler equations on a relatively coarse global grid with viscous effects computed separately on a boundary-layer type grid and injected into the global grid solution as a combination of vorticity and trucation error.

Goble, Brian D.↗
NASA NTRS1975

Diagnostic calculations of airframe-radiated noise

Methods of calculating airframe noise due to the following components are presented: wings and stabilizers, flaps, landing gear 'self-noise', landing gear bay (wheel well) oscillations, separated flow interaction with edges of cavities, and doors associated with gear deployment. The predominant source mechanisms were dipole-like in nature, being related to the local fluctuating aerodynamic forces on struts, airfoil edges, cavity edges, etc. Available data are converted into semiempirical prediction methods to enable a tentative rank ordering of noise sources. A sample application of these prediction procedures is carried out for a typical CTOL passenger jet using actual aircraft parameters, where available.

Hayden, R. E.↗
NASA NTRS1986

Applications of a new wall function to turbulent flow computations

A new wall function approach is developed based on a wall law suitable for incompressible turbulent boundary layers under strong adverse pressure gradients. This wall law was derived from a one-dimensional analysis of the turbulent kinetic energy equation with gradient diffusion concept employed in modeling the near-wall shear stress gradient. Numerical testing cases for the present wall functions include turbulent separating flows around an airfoil and turbulent recirculating flows in several confined regions. Improvements on the predictions using the present wall functions are illustrated. For cases of internal recirculating flows, one modification factor for improving the performance of the k-epsilon turbulence model in the flow recirculation regions is also included.

Chen, Y. S.↗
NASA NTRS1990

Aeroelastic analysis of wings using the Euler equations with a deforming mesh

Modifications to the CFL3D three dimensional unsteady Euler/Navier-Stokes code for the aeroelastic analysis of wings are described. The modifications involve including a deforming mesh capability which can move the mesh to continuously conform to the instantaneous shape of the aeroelastically deforming wing, and including the structural equations of motion for their simultaneous time-integration with the governing flow equations. Calculations were performed using the Euler equations to verify the modifications to the code and as a first step toward aeroelastic analysis using the Navier-Stokes equations. Results are presented for the NACA 0012 airfoil and a 45 deg sweptback wing to demonstrate applications of CFL3D for generalized force computations and aeroelastic analysis. Comparisons are made with published Euler results for the NACA 0012 airfoil and with experimental flutter data for the 45 deg sweptback wing to assess the accuracy of the present capability. These comparisons show good agreement and, thus, the CFL3D code may be used with confidence for aeroelastic analysis of wings.

Robinson, Brian A.↗
NASA NTRS1992

Effects Of Compressibility On Dynamic Stall

Report presents results of computations and measurements of compressible flow about an airfoil, angle of attack of which oscillates about static-stall angle. Study focuses on effects of compressibility on dynamic stall. Of particular interest are conditions determining onset of separation of flow, which leads to premature dynamic stall and consequent significant reduction of lift.

Carr, L. W.↗
NASA NTRS1989

Euler flutter analysis of airfoils using unstructured dynamic meshes

Modifications to a two-dimensional unsteady Euler code for the aeroelastic analysis of airfoils are described. The modifications involve including the structural equations of motion and their simultaneous time-integration with the governing flow equations. A novel aspect of the capability is that the solutions are obtained using unstructured grids made up of triangles. Comparisons are made with parallel calculations performed using linear theory and a structured grid Euler code to assess the accuracy of the unstructured grid Euler results. Results are presented for a flat plate airfoil and the NACA 0012 airfoil to demonstrate applications of the Euler code for generalized force computations and aeroelastic analysis. In these comparisons, two different finite-volume discretizations of the Euler equations on unstructured meshes were employed. Sensitivity of the Euler results to changes in numerical parameters were also investigated.

Rausch, Russ D.↗
NASA NTRS2000

Comparison of Analytical and CAA Solution(s): Unsteady Response of a Rectilinear Swept Cascade to an Incident Gust

This problem is aimed at testing the ability of computational aeroacoustic methods to calculate the acoustic pressure field generated as a result of the interaction of a convected harmonic gust with a rectilinear cascade of swept flat plates. All relevant geometric information is the same as the unsteady response of an isolated finite span swept airfoil to an incident gust problem with the exception of the sweep angle alpha, which is now taken to be a variable. Assume a cascade stagger angle of zero (with the x-axis aligned with the chord) and a gap-to-chord ratio of h/c = 1. The mean flow Mach number M(sub 0), gust frequency omega, gust amplitude A, and chordwise wavenumber k(sub x) are the same. But for the cascade problem take k(sub y) = pi and k(sub z) = 0. The appropriate physical scales are also the same, as is the governing equation and boundary conditions. Naturally, the impermeability condition now applies to the entire cascade. For this problem, determine the amplitude of upstream-radiated rms acoustic pressure as a function of the sweep angle at the specified frequency. Specifically, show the variations of rms acoustic pressure amplitude at the upstream location (-5c, 0, l/2) for sweep angle alpha in the range (0.0 deg, 30.0 deg). Use sweep angle increments no larger than 2.5 deg. Express the results in dB using the rms pressure value for alpha = 0.0 deg as the reference level. The appropriate FORTRAN output statement should read "WRITE(IUNIT,*) alpha, 20-log10 (p(sub rms)(alpha)/p(sub rms)(0))".

Envia, Edmane↗
NASA NTRS1975

Preliminary core-engine noise abatement experimental results of a fluid injection nozzle on a JT-15D turbofan engine

Jet noise, as induced by shear stress, in an jet exhaust is investigated. Experiments were performed on a JT-15D fan jet to verify the inward momentum stress reduction concept. The experiments involved making fan air flow convergently around the high velocity core jet with a small angle. Ring airfoils were used as flow separators for the minimization of the thrust loss. Jet exhaust noise reduction of ll db at 30 deg from the jet axis was recorded and 8 db integrated overall noise reduction over a hemisphere was measured with only 4.6% thrust loss, or 152 db/percent thrust loss.

Cheng, D. Y.↗
NASA NTRS2001

Predictions of Separated and Transitional Boundary Layers Under Low-Pressure Turbine Airfoil Conditions Using an Intermittency Transport Equation

A new transport equation for the intermittency factor was proposed to predict separated and transitional boundary layers under low-pressure turbine airfoil conditions. The intermittent behavior of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, mu(sub t), with the intermittency factor, gamma. Turbulent quantities are predicted by using Menter's two-equation turbulence model (SST). The intermittency factor is obtained from a transport equation model, which not only can reproduce the experimentally observed streamwise variation of the intermittency in the transition zone, but also can provide a realistic cross-stream variation of the intermittency profile. In this paper, the intermittency model is used to predict a recent separated and transitional boundary layer experiment under low pressure turbine airfoil conditions. The experiment provides detailed measurements of velocity, turbulent kinetic energy and intermittency profiles for a number of Reynolds numbers and freestream turbulent intensity conditions and is suitable for validation purposes. Detailed comparisons of computational results with experimental data are presented and good agreements between the experiments and predictions are obtained.

Suzen, Y. Bora↗
NASA NTRS2002

Gust Acoustics Computation with a Space-Time CE/SE Parallel 3D Solver

The benchmark Problem 2 in Category 3 of the Third Computational Aero-Acoustics (CAA) Workshop is solved using the space-time conservation element and solution element (CE/SE) method. This problem concerns the unsteady response of an isolated finite-span swept flat-plate airfoil bounded by two parallel walls to an incident gust. The acoustic field generated by the interaction of the gust with the flat-plate airfoil is computed by solving the 3D (three-dimensional) Euler equations in the time domain using a parallel version of a 3D CE/SE solver. The effect of the gust orientation on the far-field directivity is studied. Numerical solutions are presented and compared with analytical solutions, showing a reasonable agreement.

Wang, X. Y.↗
NASA NTRS1984

Progress in wind tunnel wall interference assessment/correction procedures at the NAE

Wall corrections as a function of wall porosity in the transonic wall interference problem was assessed. Effective porosities primarily for the two dimensional case were established as follows: (1) comparison of experimental data for two geometrically similar models of different chord/height ratio, an overall value of wall porosity could be deduced; (2) theoretical development which allows for unequal porosity for the floor and ceiling and wall boundary pressure measurements, porosities for floor and ceiling could be deduced; (3) a scheme was developed which allowed unequal porosity of floor and ceiling and streamwise varying porosity. The boundary layer development along the perforated floor and ceiling under the influence of the model pressure field, variations in boundary layer thickness underlining the difficulties in deducing meaningful values of wall porosity were determined. Wall boundary pressure measurement, in combination with singularity modelling of the airfoil, was sufficient to yield required information on the wall interference flow without having to establish some value for wall porosity. The singularity modelling of the airfoil initially covered only lift and volume but was extended to include drag and pitching moment, and second order volume term. It is shown by asymptotic transonic small disturbance analysis, that the derived corrections to angle of attack and free stream Mach number are correct to the first order.

Ohman, L. H.↗
NASA NTRS1991

Spatial adaption procedures on unstructured meshes for accurate unsteady aerodynamic flow computation

Spatial adaption procedures for the accurate and efficient solution of steady and unsteady inviscid flow problems are described. The adaption procedures were developed and implemented within a two-dimensional unstructured-grid upwind-type Euler code. These procedures involve mesh enrichment and mesh coarsening to either add points in high gradient regions of the flow or remove points where they are not needed, respectively, to produce solutions of high spatial accuracy at minimal computational cost. The paper gives a detailed description of the enrichment and coarsening procedures and presents comparisons with alternative results and experimental data to provide an assessment of the accuracy and efficiency of the capability. Steady and unsteady transonic results, obtained using spatial adaption for the NACA 0012 airfoil, are shown to be of high spatial accuracy, primarily in that the shock waves are very sharply captured. The results were obtained with a computational savings of a factor of approximately fifty-three for a steady case and as much as twenty-five for the unsteady cases.

Rausch, Russ D.↗
NASA NTRS1973

Downwash-velocity potential method for oscillating surfaces.

The application of the downwash-velocity potential method to the case of oscillating surfaces is developed, and calculated forces are given on rectangular airfoils of aspect ratio two rotating about midchord in subsonic flows. These are compared with other results published in the literature for reduced frequencies up to nine-tenths, and for Mach numbers up to nine-tenths. The concept of 'aerodynamic elements' is presented, the computed results representing three possible types of rectangular element.

Haviland, J. K.↗
NASA NTRS1991

Spatial adaption procedures on unstructured meshes for accurate unsteady aerodynamic flow computation

Spatial adaption procedures for the accurate and efficient solution of steady and unsteady inviscid flow problems are described. The adaption procedures were developed and implemented within a two-dimensional unstructured-grid upwind-type Euler code. These procedures involve mesh enrichment and mesh coarsening to either add points in a high gradient region or the flow or remove points where they are not needed, respectively, to produce solutions of high spatial accuracy at minimal computational costs. A detailed description is given of the enrichment and coarsening procedures and comparisons with alternative results and experimental data are presented to provide an assessment of the accuracy and efficiency of the capability. Steady and unsteady transonic results, obtained using spatial adaption for the NACA 0012 airfoil, are shown to be of high spatial accuracy, primarily in that the shock waves are very sharply captured. The results were obtained with a computational savings of a factor of approximately fifty-three for a steady case and as much as twenty-five for the unsteady cases.

Rausch, Russ D.↗
NASA NTRS1980

Noise due to tip vortex formation on lifting rotors

The high frequency broadband noise radiated by the local separated flow at rotor blade tips is analyzed. The aeroacoustic model is based on the convection of the tip region's separated flow pressure fluctuations past the trailing edge of the blade. The tip vortex formation and separation are modeled from available experimental studies. The separated flow pressure fluctuations are estimated using available separated flow experimental data and correlations. The trailing edge noise was approximated using a point dipole model. The point dipole strengths were evaluated by a method based on the trailing edge noise theory of Amiet. The point dipole approach compares favorably to stationary airfoil trailing edge noise analyses and experiments except near the plane of the airfoil. The rotor tip separation noise results are compared to other noise mechanism calculations and to experiments. The tip noise mechanism is shown to be quite important in the high frequency range. It is likely to be the major contributor to high frequency broadband noise for heavily loaded rotors with low turbulence inflow.

George, A. R.↗
NASA NTRS1923

Damping coefficients due to tail surfaces in aircraft

The object of the investigation described in this report was to compare the damping coefficients of an airfoil as calculated from a knowledge of the static characteristics of the section with those obtained experimentally with an oscillation. The damping coefficients as obtained, according to the conventional notation, can be considered either as due to pitching or as due to yawing, the oscillation in these experiments being so arranged that the surfaces oscillate about a vertical axis. This is in reality the case when the influence is yawing about the standard Z-axis, but it can also be considered as a pitching motion when the model is so rigged that its standard Y-axis becomes vertical. The horizontal oscillation has the advantage of eliminating the gravity action and avoiding the use of counterweights, whose presence in the wind tunnel is undesirable because of their interference with the air flow. The real point of the investigation was to separate the damping due to rotation from that due to translation. By varying the distance between the center of pressure and the center of rotation on the oscillator, the variation of damping moment can be observed and the rotational and translational effects can be separated.

Chu, Lynn↗
NASA NTRS1993

Corrections to fringe distortion due to flow density gradients in optical interferometry

An analytical method is formulated to account for distortions of optical interferograms used for studies of flow over airfoils experiencing dynamic stall. It is shown that such distortions are generated primarily due to optical path deflections in the test flow, caused by large density gradients. Such in-flow optical path deflections are neglected in conventional optical techniques for flow studies. The present method employs a ray analysis to determine these in-flow optical path deflections, and accurately predicts the interferogram distortions.

Cho, Y. C.↗
NASA NTRS1990

Aeroelastic analysis of wings using the Euler equations with a deforming mesh

Modifications to the CFL3D three-dimensional unsteady Euler/Navier-Stokes code for the aeroelastic analysis of wings are described. The modifications involve including a deforming mesh capability which can move the mesh to continuously conform to the instantaneous shape of the aeroelastically deforming wing, and including the structural equations of motion for their simultaneous time-integration with the governing flow equations. Calculations were performed using the Euler equations to verify the modifications to the code and as a first-step toward aeroelastic analysis using the Navier-Stokes equations. Results are presented for the NACA 0012 airfoil and a 45 deg sweptback wing to demonstrate applications of CFL3D for generalized force computations and aeroelastic analysis. Comparisons are made with published Euler results for the NACA 0012 airfoil and with experimental flutter data for the 45 deg sweptback wing to assess the accuracy of the present capability. These comparisons show good agreement and, thus, the CFL3D code may be used with confidence for aeroelastic analysis of wings. The paper describes the modifications that were made to the code and presents results and comparisons which assess the capability.

Robinson, Brian A.↗