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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Navier-Stokes calculation of transonic flow past the NTF 65 deg delta wing

Models of four delta wings were built and tested in the 8 x 8 ft Transonic Wind Tunnel at LaRC. The wings are identical in planform shape with a swept-back angle of 65 degrees, but bear different leading edge profiles. The models were tested under pressurized and cryogenic conditions to simulate true flight Reynolds numbers. Data on the aerodynamic forces and pressure distributions at various locations on the wings were taken at various flight Mach numbers and angles of attack. Effects of high Reynolds number and leading edge radius on the aerodynamic characteristics of the wings are being accessed. To thoroughly understand the turbulent, vortical flow around the wings, an effort to perform computational aerodynamic analysis is being made. The objective of the analysis is to supplement and validate the experimental data and explain the high Reynolds number and leading edge effects. GRIDGEN, a software developed by General Dynamics, is being used to generate the grid topology for the flow field around the wings. The flow solver to be used is CFL3D, a computation fluid dynamics code developed at LaRC. Based on the geometric description of the wings, a FORTRAN program called WINGSURF was written to generate the databases defining the surface geometry of the wings. To match the true geometry of the models in the wind tunnel for realistic comparison with experimental results, databases defining the sting support for the wing models were also created by two other FORTRAN programs, STINJOIN and STINREAR. Listing of the programs are attached and the geometry of a typical wing model with the sting support is shown. Other aspects of the investigation are discussed.

Wu, Chivey C.↗
NASA NTRS1999

Uncertainty Analysis of Instrument Calibration and Application

Experimental aerodynamic researchers require estimated precision and bias uncertainties of measured physical quantities, typically at 95 percent confidence levels. Uncertainties of final computed aerodynamic parameters are obtained by propagation of individual measurement uncertainties through the defining functional expressions. In this paper, rigorous mathematical techniques are extended to determine precision and bias uncertainties of any instrument-sensor system. Through this analysis, instrument uncertainties determined through calibration are now expressed as functions of the corresponding measurement for linear and nonlinear univariate and multivariate processes. Treatment of correlated measurement precision error is developed. During laboratory calibration, calibration standard uncertainties are assumed to be an order of magnitude less than those of the instrument being calibrated. Often calibration standards do not satisfy this assumption. This paper applies rigorous statistical methods for inclusion of calibration standard uncertainty and covariance due to the order of their application. The effects of mathematical modeling error on calibration bias uncertainty are quantified. The effects of experimental design on uncertainty are analyzed. The importance of replication is emphasized, techniques for estimation of both bias and precision uncertainties using replication are developed. Statistical tests for stationarity of calibration parameters over time are obtained.

Tripp, John S.↗
NASA NTRS1993

Flow resolution and domain of influence in rarefied hypersonic blunt-body flows

The study assesses the effects of the upstream domain size and grid resolution upon flow properties and body aerodynamics computed for rarefied flows over cold blunt bodies with a direct simulation Monte Carlo (DSMC) particle method. Empirical correlations are suggested for aerodynamic coefficients for two-dimensional flows past a perpendicular flat plate. Free-stream parameters which were varied in the study include the Mach number, Knudsen number, surface temperature, and intermolecular potential. Insufficient grid resolution leads to overprediction of aerodynamic heating and forces in the DSMC method. Solution accuracy correlates well with the Reynolds number defined at the wall temperature and the stagnation mean free path relative to the cell dimension. Insufficient upstream domain size in the DSMC method leads to overprediction of heating and drag. Errors in aerodynamic coefficients correlate well with the distance ahead of the body where flow temperature reaches half of its peak value. Simulation of a hard-sphere gas is more sensitive to grid resolution, while simulation of a Maxwell gas is more sensitive to upstream domain size.

Haas, Brian L.↗
NASA NTRS2008

Transient Three-Dimensional Startup Side Load Analysis of a Regeneratively Cooled Nozzle

The objective of this effort is to develop a computational methodology to capture the startup side load physics and to anchor the computed aerodynamic side loads with the available data from a regeneratively cooled, high-aspect-ratio nozzle, hot-fired at sea level. The computational methodology is based on an unstructured-grid, pressure-based, reacting flow computational fluid dynamics and heat transfer formulation, a transient 5 s inlet history based on an engine system simulation, and a wall temperature distribution to reflect the effect of regenerative cooling. To understand the effect of regenerative wall cooling, two transient computations were performed using the boundary conditions of adiabatic and cooled walls, respectively. The results show that three types of shock evolution are responsible for side loads: generation of combustion wave; transitions among free-shock separation, restricted-shock separation, and simultaneous free-shock and restricted shock separations; along with the pulsation of shocks across the lip, although the combustion wave is commonly eliminated with the sparklers during actual test. The test measured two side load events: a secondary and lower side load, followed by a primary and peak side load. Results from both wall boundary conditions captured the free-shock separation to restricted-shock separation transition with computed side loads matching the measured secondary side load. For the primary side load, the cooled wall transient produced restricted-shock pulsation across the nozzle lip with peak side load matching that of the test, while the adiabatic wall transient captured shock transitions and free-shock pulsation across the lip with computed peak side load 50% lower than that of the measurement. The computed dominant pulsation frequency of the cooled wall nozzle agrees with that of a separate test, while that of the adiabatic wall nozzle is more than 50% lower than that of the measurement. The computed teepee-like formation and the tangential motion of the shocks during lip pulsation also qualitatively agree with those of test observations. Moreover, a third transient computation was performed with a proportionately shortened 1 s sequence, and lower side loads were obtained with the higher ramp rate.

Wang, Ten-See↗
NASA NTRSDate not supplied

Fan Noise Predictions of the NASA Source Diagnostic Test using Unsteady Simulations with LAVA Part II - Tonal and Broadband Noise Assessment

The NASA Source Diagnostic Test (SDT) campaign experimental data is used for validation of a sliding mesh technique recently implemented within the Launch, Ascent, and Vehicle Aerodynamics computational framework for time-accurate simulation of rotating fans. The far-field acoustics are analyzed in this work, building upon the aerodynamic validation studies previously published in Part I. Two modeling approaches are explored: the unsteady Reynolds-averaged Navier Stokes (uRANS) with the negative Spalart-Allmaras (SA-neg) turbulence model closure, and a hybrid Reynolds-averaged Navier Stokes/large-eddy simulation (RANS/LES) paradigm employing a zonal detached-eddy simulation (ZDES) closure with enhanced shielding protection. Two convective flux scheme approaches with different dissipation properties are also explored with ZDES. The Ffowcs-Williams and Hawkings (FW-H) permeable surface approach is used for propagation of the near-field acoustics to the far-field microphone locations. This work analyzes the low-speed (approach) condition, characterized by a fan rotation speed of 7808 rotations-per-minute (RPM). Three different grid levels ranging between 200 million and 1.1 billion grid points are considered. Results show good prediction of the broadband noise levels at sideline angles ranging between 70° and 110°. The forward arc observers show an under-prediction of the overall sound pressure levels (OASPL) even at the fine grid level. A breakdown of the inlet and exhaust contributions reveals a steep drop-off in the broadband noise levels past a blade-passing frequency (BPF) of 1.5, potentially caused by a lack of resolved small-scale turbulent fluctuations in the interstage region. Past 110° the OASPL are over-predicted by up to 10 dB due to an over-prediction of the low-frequency broadband noise levels in the aft arc. The detuning of BPF2 caused by small deviations in the blade stagger angle around the wheel is captured, and a corresponding decrease in the sound power level for this cut-on tone is observed.

AATT↗
NASA NTRS1993

Nonlinear aerodynamics and the design of wing tips

This contract began in April 1990 with studies on the effect of wingtip shape on induced drag. That work considered problems of determining induced drag from computational aerodynamic methods and examined effects of wake roll up on the induced drag of elliptical and crescent wings. The research contract was continued in April 1991 and again in April 1992 with scope augmented to include considerations other than induced drag. The idea was to develop methods for comparing wing tip shapes on the basis of their effects on total drag, structural weight, and high lift performance. To accomplish this, work was done to improve methods for multidisciplinary analysis and optimization of wings. This report describes results of research conducted from April 1992 through March 1993. The general objective was to improve an existing wing optimization method and apply the method to specific problems of interest. The method, while a valuable tool for wing tip design studies, can be applied to more general problems, and has been applied to some of these other problems during its development. Specific goals that were accomplished are listed below and are explained in more detail in attached reports: analysis methods were improved in the area of structural modeling; small improvements to the high lift model were made, and ideas were collected on how clean wing c(sub l(sub max)) should vary with sweep - these ideas should provide a guide for future examination of high lift with experiment or CFD; and a calculation for drag due to the loss of leading edge suction was added to the wing analysis.

Kroo, Ilan↗
NASA NTRS1989

Computer Simulation of Aircraft Aerodynamics

The role of Ames Research Center in conducting basic aerodynamics research through computer simulations is described. The computer facilities, including supercomputers and peripheral equipment that represent the state of the art, are described. The methodology of computational fluid dynamics is explained briefly. Fundamental studies of turbulence and transition are being pursued to understand these phenomena and to develop models that can be used in the solution of the Reynolds-averaged Navier-Stokes equations. Four applications of computer simulations for aerodynamics problems are described: subsonic flow around a fuselage at high angle of attack, subsonic flow through a turbine stator-rotor stage, transonic flow around a flexible swept wing, and transonic flow around a wing-body configuration that includes an inlet and a tail.

Inouye, Mamoru↗
NASA NTRS1991

Computation of supersonic jet mixing noise for an axisymmetric CD nozzle using k-epsilon turbulence model

The turbulent mixing noise of a supersonic jet is calculated for a round convergent-divergent nozzle at the design pressure ratio. Aerodynamic computations are performed using the PARC code with a k-epsilon turbulence model. Lighthill's acoustic analogy combined with Ribner's assumption is adopted. The acoustics solution is based upon the methodology followed by GE in the MGB code. The source correlation function is expressed as a linear combination of second-order tensors. Assuming separable second-order correlations and incorporating Batchelor's isotropic turbulence model, the source term was calculated from the kinetic energy of turbulence. A Gaussian distribution for the time-delay of correlation was introduced. The computational fluid dynamics (CFD) solution was used to obtain the source strength as well as the characteristic time-delay of correlation. The effect of sound/flow interaction was incorporated using the high frequency asymptotic solution to Lilley's equation for axisymmetric geometries. Acoustic results include sound pressure level directivity and spectra at different polar angles. The aerodynamic and acoustic results demonstrate favorable agreement with experimental data.

Khavaran, Abbas↗
NASA NTRS1992

Computation of supersonic jet mixing noise for an axisymmetric CD nozzle using k-epsilon turbulence model

The turbulent mixing noise of a supersonic jet is calculated for a round convergent-divergent nozzle at the design pressure ratio. Aerodynamic computations are performed using the PARC code with a k-epsilon turbulence model. Lighthill's acoustic analogy combined with Ribner's assumption is adopted. The acoustics solution is based upon the methodology followed by GE in the MGB code. The source correlation function is expressed as a linear combination of second-order tensors. Assuming separable second-order correlations and incorporating Batchelor's isotropic turbulence model, the source term was calculated from the kinetic energy of turbulence. A Gaussian distribution for the time-delay of correlation was introduced. The computational fluid dynamics (CFD) solution was used to obtain the source strength as well as the characteristic time-delay of correlation. The effect of sound/flow interaction was incorporated using the high frequency asymptotic solution to Lilley's equation for axisymmetric geometries. Acoustic results include sound pressure level directivity and spectra at different polar angles. The aerodynamic and acoustic results demonstrate favorable agreement with experimental data.

Khavaran, A.↗
NASA NTRS2011

Analysis of Asymmetric Aircraft Aerodynamics Due to an Experimental Wing Glove

Aerodynamic computational fluid dynamics analysis of a wing glove attached to one wing of a business jet is presented and discussed. A wing glove placed on only one wing will produce asymmetric aerodynamic effects that will result in overall changes in the forces and moments acting on the aircraft. These changes, referred to as deltas, need to be determined and quantified to ensure that the wing glove does not have a significant effect on the aircraft flight characteristics. TRANAIR (Calmar Research Corporation, Cato, New York), a nonlinear full potential solver, and Star-CCM+ (CD-adapco, Melville, New York), a finite volume full Reynolds-averaged Navier-Stokes computational fluid dynamics solver, are used to analyze a full aircraft with and without the glove at a variety of flight conditions, aircraft configurations, and angles of attack and sideslip. Changes in the aircraft lift, drag, and side force along with roll, pitch, and yaw are presented. Span lift and moment distributions are also presented for a more detailed look at the effects of the glove on the aircraft. Aerodynamic flow phenomena due to the addition of the glove are discussed. Results show that the glove produces only small changes in the aerodynamic forces and moments acting on the aircraft, most of which are insignificant.

Hartshorn, Fletcher↗
NASA NTRS1996

Aerodynamic Influence Coefficient Computations Using Euler/Navier-Stokes Equations on Parallel Computers

Modem design requirements for an aircraft push current technologies used in the design process to their limit or sometimes require more advanced technologies to meet the requirement. New design requirements always demand to improve the operational performance. Accurate prediction of aerodynamic coefficients is essential to improve the performance. For example, in the design of an advanced subsonic civil transport, since the fluid flow at transonic regime shows strong nonlinearities, high fidelity equations, such as the Euler or Navier-Stokes equations predict flow characteristics more accurately than the linear aerodynamics, which are widely used in the current design process However, high fidelity flow equations are computationally expensive and require an order of magnitude longer time to obtain aerodynamic coefficients required in the design. Parallel computing is one possibility to cut down the computational turn-around time in using high fidelity equations so that high fidelity equations would be incorporated into the design process. By doing so, high fidelity equations would be used in the routine design process. This work will demonstrate the feasibility of using high fidelity flow equations in a design process by computing aerodynamic influence coefficients of a wing-body-empennage configuration on a multiple-instruction, multiple-data parallel computer.

Byun, Chansup↗
NASA NTRS1986

Analysis of aerodynamic coefficients using gradient data: Spanwise turbulence effects on airplane response

The influence of spanwise turbulence on airplane dynamic behavior is determined. Calculations are based on data collected from test flights with the NASA B-57 research aircraft. The approach is to first compute aerodynamic forces and moments due to a spanwise distribution of angle of attack and airspeed. Secondly, these quantities are incorporated into the equations of motion. Simulation of flights done with the effects of spanwise turbulence included are compared to simulations without any spanwise turbulence. The findings of the study are that the moments developed by turbulence along the span are significant and that more realistic flight simulation can be achieved by including the spanwise turbulence terms.

Ringnes, E. A.↗
NASA NTRS1986

An Euler aerodynamic method for leading-edge vortex flow simulation

The current capabilities and the future plans for a three dimensional Euler Aerodynamic Method are described. The basic solution algorithm is based on the finite volume, Runge-Kutta pseudo-time-stepping scheme of FLO-57. Several modifications to improve accuracy and computational efficiency were incorporated and others are being investigated. The computer code is used to analyze a cropped delta wing at 0.6 Mach number and an arrow wing at 0.85 Mach number. Computed aerodynamic parameters are compared with experimental data. In all cases, the configuration is impulsively started and no Kutta condition is applied at sharp edges. The results indicate that with additional development and validation, the present method will be a useful tool for engineering analysis of high speed aircraft.

Raj, P.↗
NASA NTRS1990

Assured Crew Return Vehicle flowfield and aerodynamic characteristics

A lifting body has been proposed as a candidate for the Assured Crew Return Vehicle which will serve as crew rescue vehicle for the Space Station. The focus of this work is on body surface definition, surface and volume grid definition, and the computation of inviscid flowfields about the vehicle at wind-tunnel conditions. Very good agreement is shown between the computed aerodynamic characteristics of the vehicle at a freestream Mach number of 10 and those measured in wind-tunnel tests.

Weilmuenster, K. James↗
NASA NTRS1989

The design of a low Reynolds number RPV

This paper covers the ongoing design process for a long endurance Remotely Piloted Vehicle (RPV). The 45-pound RPV must operate in the 25-50 knot speed range at low altitudes. Specially designed low-Reynolds number airfoils were used for efficient aerodynamics. A wing of AR = 23 was selected which will give an estimated L/D(max) of 27. Wing tip feathers will be used to reduce the induced drag. A comparison between the computed aerodynamic predictions and wind tunnel results is given. The predicted handling quality and performance results are given.

Siddiqi, S.↗
NASA NTRS1989

Euler solvers for transonic applications

The 1980s may well be called the Euler era of applied aerodynamics. Computer codes based on discrete approximations of the Euler equations are now routinely used to obtain solutions of transonic flow problems in which the effects of entropy and vorticity production are significant. Such codes can even predict separation from a sharp edge, owing to the inclusion of artificial dissipation, intended to lend numerical stability to the calculation but at the same time enforcing the Kutta condition. One effect not correctly predictable by Euler codes is the separation from a smooth surface, and neither is viscous drag; for these some form of the Navier-Stokes equation is needed. It, therefore, comes as no surprise to observe that the Navier-Stokes has already begun before Euler solutions were fully exploited. Moreover, most numerical developments for the Euler equations are now constrained by the requirement that the techniques introduced, notably artificial dissipation, must not interfere with the new physics added when going from an Euler to a full Navier-Stokes approximation. In order to appreciate the contributions of Euler solvers to the understanding of transonic aerodynamics, it is useful to review the components of these computational tools. Space discretization, time- or pseudo-time marching and boundary procedures, the essential constituents are discussed. The subject of grid generation and grid adaptation to the solution are touched upon only where relevant. A list of unanswered questions and an outlook for the future are covered.

Vanleer, Bram↗
NASA NTRS1984

Status and prospects of computational fluid dynamics for unsteady transonic viscous flows

Applications of computational aerodynamics to aeronautical research, design, and analysis have increased rapidly over the past decade, and these applications offer significant benefits to aeroelasticians. The past developments are traced by means of a number of specific examples, and the trends are projected over the next several years. The crucial factors that limit the present capabilities for unsteady analyses are identified; they include computer speed and memory, algorithm and solution methods, grid generation, turbulence modeling, vortex modeling, data processing, and coupling of the aerodynamic and structural dynamic analyses. The prospects for overcoming these limitations are presented, and many improvements appear to be readily attainable. If so, a complete and reliable numerical simulation of the unsteady, transonic viscous flow around a realistic fighter aircraft configuration could become possible within the next decade. The possibilities of using artificial intelligence concepts to hasten the achievement of this goal are also discussed.

Mccroskey, W. J.↗
NASA NTRS1985

Status and prospects of computational fluid dynamics for unsteady transonic flow

Applications of computational aerodynamics to aeronautical research, design, and analysis have increased rapidly over the past decade, and these applications offer significant benefits to aeroelasticians. The past developments are traced by means of a number of specific examples, and the trends are projected over the next several years. The crucial factors that limit the present capabilities for unsteady analyses are identified; they include computer speed and memory, algorithm and solution methods, grid generation, turbulence modeling, vortex modeling, data processing, and coupling of the aerodynamic and structural dynamic analyses. The prospects for overcoming these limitations are presented, and many improvements appear to be readily attainable. If so, a complete and reliable numerical simulation of the unsteady, transonic viscous flow around a realistic fighter aircraft configuration could become possible within the next decade. The possibilities of using artificial intelligence concepts to hasten the achievement of this goal are also discussed.

Mccroskey, W. J.↗