SEARCH · Engineering Papers

Results for “UNIFORM FLOW”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 451 records · Page 25

All sources NASA NTRS DOE OSTI

NASA NTRS1994

Aeroacoustic computation of gust-blade interaction

To better understand and address the challenges faced in computing the acoustics of flow fields, test problems must be considered. In the present study, the sound radiated by the interaction of a flat plate with an oncoming gust containing a two component, mean velocity is computed. The gust has a uniform mean flow in x with Mach number M(infinity) equal to 0.5. The gust's mean velocity in y is of smaller amplitude and is given by: v = 0.1 sin(pi/8(x/M(sub infinity) - t)). This problem has been posed for an upcoming ICASE/LaRC workshop on benchmark problems in computational aeroacoustics. A plate with a length of 30 units in x is used. The plate is assumed to be infinitesimally thin and is centered at the origin. All variables are made dimensionless using the scales specified. Acoustic quantities are obtained by numerically integrating the linearized Euler equations. Integration is performed on the computational domain -100.0 less than or equal to x less than or equal to 100.0, -100.0 less than or equal to y less than or equal to 100.0, using unit length grid spacing in x and in y. An integration scheme is sought which will provide accurate solution to the small quantities of interest at a minimal computational expense. Results indicate that with the given discretization a scheme of minimal fourth order accuracy might be adequate to approximate the waves within the given flow. Thus, a variation of the MacCormack scheme with fourth order accuracy in space and second order accuracy in time was chosen. A scheme with sixth order accuracy in space has also been implemented and results compared with those of the fourth order accurate scheme. To ensure no mass flux, zero normal velocity is assigned at the plate. This condition will induce a discontinuity in the pressure across the plate location. Values for the perturbation pressure p' along the surface of the plate are obtained using a one-sided, third order Taylor expansion, such that p'(sub y) = O. In accordance with the Kutta condition, perturbation pressure at the trailing edge is assigned to zero. In the far field, radiation boundary conditions have been implemented. The effectiveness of the far field conditions are validated by computing in a larger computational domain and comparing the results. Early in time, sound waves begin to radiate from the plate. A Doppler effect is observed. After the initial transients disappear, the strongest waves leave the trailing edge at an approximate 45 degree angle. The intensity pattern of pressure fluctuations shows five lobes (of increasing magnitude with increasing downstream direction) emerging from the plate. Undesirable short wave contaminants are observed in the computed pressure distribution along the plate surface. For a more accurate solution at small scales, a more refined discretization will be required.

Martin, James E.↗

NASA NTRS1988

Theoretical derivation and calibration technique of a hemispherical-tipped, five-hole probe

A technique is presented for the calibration of a hemispherical tipped 0.125 inch diameter 5-hole probe. The derivation of equations from the potential flow over a sphere relating the flow angle and velocity to pressure differentials measured by the probe is presented. The technique for acquiring the calibration data and the technique used to calculate the calibration coefficients are presented. The accuracy of the probe in both the uniform calibration flow field and the nonuniform flow field over a 75 degree swept delta wing is discussed.

Kjelgaard, Scott O.↗

NASA NTRS2003

Three-dimensional Aerodynamic Instability in Multi-stage Axial Compressors

Four separate tasks are reported. The first task: A Computational Model for Short Wavelength Stall Inception and Development In Multi-Stage Compressors; the second task: Three-dimensional Rotating Stall Inception and Effects of Rotating Tip Clearance Asymmetry in Axial Compressors; the third task:Development of an Effective Computational Methodology for Body Force Representation of High-speed Rotor 37; and the fourth task:Development of Circumferential Inlet Distortion through a Representative Eleven Stage High-speed axial compressor. The common theme that threaded throughout these four tasks is the conceptual framework that consists of quantifying flow processes at the fadcompressor blade passage level to define the compressor performance characteristics needed for addressing physical phenomena such compressor aerodynamic instability and compressor response to flow distoriton with length scales larger than compressor blade-to-blade spacing at the system level. The results from these two levels can be synthesized to: (1) simulate compressor aerodynamic instability inception local to a blade rotor tip and its development from a local flow event into the nonlinear limit cycle instability that involves the entire compressor as was demonstrated in the first task; (2) determine the conditions under which compressor stability assessment based on two-dimensional model may not be adequate and the effects of self-induced flow distortion on compressor stability limit as in the second task; (3) quantify multistage compressor response to inlet distortion in stagnation pressure as illustrated in the fourth task; and (4) elucidate its potential applicability for compressor map generation under uniform as well as non-uniform inlet flow given three-dimensional Navier-Stokes solution for each individual blade row as was demonstrated in the third task.

Suder, Kenneth↗

DOE OSTI2020

Numerical study of heat transfer enhancement in laminar and turbulent flow in enhanced tubes with uniform and non-uniform thermal boundary conditions

Various types of tube enhancements have been proposed in order to increase heat transfer while minimizing any associated increases in pressure drop. This paper presents a numerical study of two different types of enhanced tubes: the relatively common twisted tape insert tube, and a novel tube with airfoil shaped pins on the inner surface. Each tube is studied in both the laminar and turbulent flow regime. In addition to the typical constant temperature and constant flux boundary conditions, a non-constant heat flux boundary condition is applied on one half of the tube, a situation which may be encountered when absorbing solar radiation. In general, each type of enhanced tube offers a thermal performance factor greater than unity at certain Re. In addition, the non-uniform heat flux boundary condition appears to function similarly as the uniform heat flux boundary condition.

Sandlin, Matthew J.↗

DOE OSTI2024

Numerical study of heat transfer enhancement in laminar and turbulent flow in enhanced tubes with uniform and non-uniform thermal boundary conditions

Sandlin, M↗

NASA NTRS1972

Mass transport contamination study

A theoretical analysis was performed to determine the effects of outgassing and waste dumping on the contamination field around an orbiting spacecraft. The spacecraft was assumed to be spherical in shape with the mass flow emitting uniformly from the spherical surface at a constant rate and in a D'Lambertian spatial distribution. The outflow of gases were assumed to be neutrally charged and of a single species with a molecular weight characteristic of a composite of the actual species involved in the mass flow. The theoretical analysis showed that, for outgassing only, less than 1.5 percent of the outgas products will return to the Skylab spacecraft as a result of intermolecular collisions. When the total mass flow from the spacecraft, including waste dumps and reaction control motor firings, was considered, it was estimated that about 30 percent will return to the spacecraft.

Robertson, S. J.↗

DOE OSTI2020

Nemchinov-Dyson Solutions of the Two-Dimensional Axisymmetric Inviscid Compressible Flow Equations [Review]

We investigate the two-dimensional (2D) inviscid compressible flow equations in axisymmetric coordinates, constrained by an ideal gas equation of state (EOS). Beginning with the assumption that the 2D velocity field is space-time separable and linearly variable in each corresponding spatial coordinate, we proceed to derive an in nite family of elliptic or hyperbolic, uniformly expanding or contracting \gas cloud" solutions. Construction of specific example solutions belonging to this family is dependent on the solution of a system of nonlinear, coupled, second-order ordinary differential equations, and the prescription of an additional physical process of interest (e.g., uniform temperature or uniform entropy flow). The physical and computational implications of these solutions as pertaining to quantitative code verification or model qualification studies are discussed in some detail.

75 CONDENSED MATTER PHYSICS, SUPERCONDUCTIVITY AND↗

NASA NTRS1979

A theory of stationary long waves. I - A simple theory of blocking. II - Resonant Rossby waves in the presence of realistic vertical shears

A theory of stationary long waves in the atmosphere is developed, with attention given to the blocking phenomenon caused by the resonant amplification of large-scale planetary waves in a uniform zonal flow and to resonant Rossby waves in an atmosphere with winds varying with height. A simple model is presented to illustrate the temporal behavior of Rossby waves forced by topography and differential heating of the land and the sea, using a beta-plane approximation and assuming a spatially uniform zonal wind which may vary in time, and quasi-geostrophic disturbances. Results are then extended to the case of resonant Rossby waves in the presence of realistic vertical shears. Numerical experiments in which the wind profiles are varied in a number of physically possible manners reveal the most favorable wind configurations for resonance.

Tung, K. K.↗

NASA NTRS1975

Numeric calculation of unsteady forces over thin pointed wings in sonic flow

A fast and reasonably accurate numerical procedure is proposed for the solution of a simplified unsteady transonic equation. The approach described takes into account many of the effects of the steady flow field. The resulting accuracy is within a few per cent and can be carried out on a computer in less than one minute per case (one frequency and one mode of oscillation). The problem concerns a rigid pointed wing which performs harmonic pitching oscillations of small amplitude in a steady uniform transonic flow. Wake influence is ignored and shocks must be weak. It is shown that the method is more flexible than the transonic box method proposed by Rodemich and Andrew (1965) in that it can easily account for variable local Mach number and rather arbitrary planform so long as the basic assumptions are fulfilled.

Kimble, K. R.↗

NASA NTRS1983

Sound radiation from a vibrating cylinder in flow

An alternative approach to the conventional manner of calculating net power output of a source by integrating the energy flux over a distant control surface is presented. The method gives a means of calculating source radiation by an integral only over the source region and eliminates the need for a possibly difficult flux integration. The radiator chosen here is a finite, open ended, cylindrical shell of negligible thickness in the presence of uniform axial flow. A homogeneous medium without external boundaries is assumed. An integral representation and, in the long wave length approximation, an analytical expression for the power radiated by the vibrating cylinder are obtained. The results show power radiation may be increased or decreased by the presence of flow, depending on the relationship between the wave number of the vibration and the size of the radiator. The result is compared with results for simpler geometries.

Smith, M.↗

DOE OSTI2020

ZND Verification Tests for Reactive Burn Models in FLAG

The ZND theory, named after Zeldovich, von Neumann, and Doering, provides a simple model for one-dimensional ideal steady-state detonation. It assumes that the detonation wave front starts with a shock that is a discontinuous jump and is followed by a finite-length reaction zone. Reactive burn (also called reactive flow) models are based on ZND theory, as they model the shock initiation and detonation process with a finite reaction rate. The ZND wave propagation test is essentially the only available test case where an analytic solution exists for verification of reactive burn models in numerical codes. However, there are extensions and variants of the ZND test that have been devised for verification of multidimensional flows. The objective of this work is to provide verification of the reactive burn models currently implemented in the Lagrangian hydrocode FLAG and investigate the influence of mesh resolution, artificial viscosity models, and the Arbitrary Lagrangian-Eulerian (ALE) Euler relaxer on the simulation results. The burn models of interest are the Wescott-Stewart-Davis (WSD) model, the Scaled Uniform Reactive Flow (SURF) model (specifically with the SURFplus model extension), and the Arrhenius shock temperature state dependent WSD (AWSD) model. Previously, Ralph Menikoff has used ZND tests for verification of the SURF and SURFplus models in the Eulerian hydrocode xRAGE. The ZND tests here are somewhat different than the approach by Menikoff. In particular, we use a piston-driven ZND detonation wave (via a prescribed constant velocity boundary condition) in a Lagrangian framework whereas Menikoff had a ZND wave followed by a invariant rarefaction wave. The xRAGE simulations were carried out on uniform grids and adaptive mesh refinement (AMR) grids. Although AMR was recently implemented in FLAG and now fully functional for 2D simulations, it will not be evaluated with ZND tests at this time. Some work has been done previously for validating the reactive burn models in FLAG. For example, the validation studies of SURF with shock-to-detonation (SDT) tests, cylinder tests, and gap-stick tests. Further validation of the AWSD, WSD, and SURF models is described in for SDT, multi-shock, cylinder, and corner-turning tests. Recently, a large parameter study with approximately twenty SDT tests was performed to validate the AWSD, WSD, SURF, and SURFplus models while also investigating mesh resolution and artificial viscosity settings. To the best of our knowledge, the current work represents the first documented verification of these burn models in FLAG. The reader should note that many of the tables and figures in this report use units of cm/g/µs, which are the default units for FLAG. However, some lengths and velocities are expressed in µm and mm/µs, respectively, which are typical units for detonation analyses.

45 MILITARY TECHNOLOGY, WEAPONRY, AND NATIONAL DEF↗

NASA NTRS1995

Category 5 problem solution using an unstructured finite volume algorithm

For the simulation of flows with complex geometries, unstructured finite volume methods have proven to be very popular, and simulations of a large number of flows have been done with good results using this approach. Since most of the simulations to date were done for steady flows, it is not clear that present unstructured finite volume algorithms can accurately track the unsteady propagation of acoustic waves in a computation. Therefore, there is a need to assess the accuracy of these methods for acoustic calculations. In this paper, we perform the numerical simulation of a very small amplitude acoustic wave incident on the non-uniform steady flow in a quasi- 1 D convergent-divergent nozzle using an unstructured finite volume algorithm with piece-wise linear, least square reconstruction, Roe flux difference splitting, and second-order MacCormack time marching. First, the spatial accuracy of the algorithm is evaluated for the steady flow by running the simulation with a sequence of successively finer meshes. Then the unsteady numerical solution with the acoustic perturbation is presented.

Bui, Trong T.↗

NASA NTRS1976

Acoustic power of a moving point source in a moving medium

The acoustic power output of a moving point-mass source in an acoustic medium which is in uniform motion and infinite in extent is examined. The acoustic medium is considered to be a homogeneous fluid having both zero viscosity and zero thermal conductivity. Two expressions for the acoustic power output are obtained based on a different definition cited in the literature for the average energy-flux vector in an acoustic medium in uniform motion. The acoustic power output of the source is found by integrating the component of acoustic intensity vector in the radial direction over the surface of an infinitely long cylinder which is within the medium and encloses the line of motion of the source. One of the power expressions is found to give unreasonable results even though the flow is uniform.

Cole, J. E., III↗

NASA NTRSDate not supplied

Evaluation of Variable Facesheet Liner Configurations for Broadband Noise Reduction

This paper investigates the broadband noise reduction properties of a distributed (variable) facesheet liner. Candidate uniform samples are selected using the Wave Propagation Model and Two-Parameter Model developed at the NASA Langley Research Center. Three uniform and one distributed facesheet samples are tested in the Normal Incidence Tube (NIT) at NASA Langley. Initial results show that the distributed facesheet sample possesses broadband noise reduction properties. Acoustic impedances educed from these NIT tests are used as inputs into the Convected Helmholtz Equation (CHE) propagation code developed at NASA Langley to synthesize the acoustic pressure field in the Grazing Flow Impedance Tube (GFIT) with no flow. Three uniform, one distributed, as well as one three-zone liner, oriented in two different directions, are investigated. Preliminary results showed the distributed facesheet and three-zone liners exhibit broadband noise reduction properties.

M C Brown↗

NASA NTRS1987

Designing transonic wind-tunnel test sections for flow diagnostics and laser velocimeter applications

The design, fabrication, and validation of the Basic Aerodynamic Facility (BARF), developed for the direct comparison of three-dimensional laser velocimetry (LV) and hot-wire techniques at transonic speeds, are discussed. To accommodate the installation of the three-dimensional orthogonal LV system in the BARF, the transonic slots were redesigned to house windows, and the test section mounting was redesigned to facilitate LV transversing. Two-dimensional and three-dimensional code validation tests of the LV test section showed uniform mean flow through the test section to sonic speeds, and fluctuating pressure measurements at subsonic speeds which are consistent with facilities with good flow quality.

Jones, G. S.↗

NASA NTRS1983

Calculation of potential flow past non-lifting bodies at angle of attack using axial and surface singularity methods

Two different singularity methods have been utilized to calculate the potential flow past a three dimensional non-lifting body. Two separate FORTRAN computer programs have been developed to implement these theoretical models, which will in the future allow inclusion of the fuselage effect in a pair of existing subcritical wing design computer programs. The first method uses higher order axial singularity distributions to model axisymmetric bodies of revolution in an either axial or inclined uniform potential flow. Use of inset of the singularity line away from the body for blunt noses, and cosine-type element distributions have been applied to obtain the optimal results. Excellent agreement to five significant figures with the exact solution pressure coefficient value has been found for a series of ellipsoids at different angles of attack. Solutions obtained for other axisymmetric bodies compare well with available experimental data. The second method utilizes distributions of singularities on the body surface, in the form of a discrete vortex lattice. This program is capable of modeling arbitrary three dimensional non-lifting bodies. Much effort has been devoted to finding the optimal method of calculating the tangential velocity on the body surface, extending techniques previously developed by other workers.

Shu, J. Y.↗

NASA NTRS1979

Evaluation of a new concept for reducing free-stream turbulence in wind tunnels

The feasibility of using a 45 deg-honeycomb for improving the flow quality in wind tunnels was investigated. The results of the experiments indicate that the turbulence levels in a wind tunnel using the 45 deg-honeycomb would be comparable to those obtainable using a conventional honeycomb, i.e., a honeycomb which is positioned normal to the mean flow in the settling chamber. However, this is true only when a 45 deg-screen is mounted immediately downstream of the honeycomb, and when some distance is provided between the trailing edges of the corner turning vanes and the upstream side of the 45 deg-honeycomb. This distance is required for adequate decay of the turning-vane wakes, thereby providing a reasonably uniform mean flow entering the honeycomb. Results demonstrate that this distance should be at least twice the spacing between the turning vanes. Even though the resulting turbulence intensity is the same downstream of both the 45 deg-honeycomb and the conventional honeycomb, there would have to be a significant improvement in the test-section turbulence intensity using the 45 deg-honeycomb to justify the additional expense in fabricating and installing this type of honeycomb.

Wigeland, R. A.↗

NASA NTRS2017

Numerical Experiments in Error Control for Sound Propagation Using a Damping Layer Boundary Treatment

This paper presents results from numerical experiments for controlling the error caused by a damping layer boundary treatment when simulating the propagation of an acoustic signal from a continuous pressure source. The computations are with the 2D Linearized Euler Equations (LEE) for both a uniform mean flow and a steady parallel jet. The numerical experiments are with algorithms that are third, fifth, seventh and ninth order accurate in space and time. The numerical domain is enclosed in a damping layer boundary treatment. The damping is implemented in a time accurate manner, with simple polynomial damping profiles of second, fourth, sixth and eighth power. At the outer boundaries of the damping layer the propagating solution is uniformly set to zero. The complete boundary treatment is remarkably simple and intrinsically independant from the dimension of the spatial domain. The reported results show the relative effect on the error from the boundary treatment by varying the damping layer width, damping profile power, damping amplitude, propagtion time, grid resolution and algorithm order. The issue that is being addressed is not the accuracy of the numerical solution when compared to a mathematical solution, but the effect of the complete boundary treatment on the numerical solution, and to what degree the error in the numerical solution from the complete boundary treatment can be controlled. We report maximum relative absolute errors from just the boundary treatment that range from O[10-2] to O[10-7].

Algorithm↗