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At least 1,099 records · Page 61
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Transonic CFD applications at Boeing

The use of computational methods for three dimensional transonic flow design and analysis at the Boeing Company is presented. A range of computational tools consisting of production tools for every day use by project engineers, expert user tools for special applications by computational researchers, and an emerging tool which may see considerable use in the near future are described. These methods include full potential and Euler solvers, some coupled to three dimensional boundary layer analysis methods, for transonic flow analysis about nacelle, wing-body, wing-body-strut-nacelle, and complete aircraft configurations. As the examples presented show, such a toolbox of codes is necessary for the variety of applications typical of an industrial environment. Such a toolbox of codes makes possible aerodynamic advances not previously achievable in a timely manner, if at all.

Tinoco, E. N.↗
NASA NTRS1989

Mach 5 inlet CFD and experimental results

An experimental research program was conducted in the NASA Lewis Research Center 10 x 10 ft supersonic wind tunnel. The 2-D inlet model was designed to study the Mach 3.0 to 5.0 speed range for an over-under turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both analytical code validation data as well as inlet performance information. Support studies for the program include flow field predictions with both 3-D parabolized Navier-Stokes (PNS) and 3-D full Navier-Stokes (FNS) analytical codes. Analytical predictions and experimental results are compared.

Weir, Lois J.↗
NASA NTRS1988

Application of advanced computational technology to propulsion CFD

The Internal Fluid Mechanics Division of the NASA Lewis Research Center is combining the key elements of computational fluid dynamics, aerothermodynamic experiments, and advanced computational technology to bring internal computational fluid dynamics (ICFM) to a state of practical application for aerospace propulsion system design. This paper presents an overview of efforts underway at NASA Lewis to advance and apply computational technology to ICFM. These efforts include the use of modern, software engineering principles for code development, the development of an AI-based user-interface for large codes, the establishment of a high-performance, data communications network to link ICFM researchers and facilities, and the application of parallel processing to speed up computationally intensive and/or time-critical ICFM problems. A multistage compressor flow physics program is cited as an example of efforts to use advanced computational technology to enhance a current NASA Lewis ICFM research program.

Szuch, John R.↗
NASA NTRS1989

CFD simulation of square cross-section, contoured nozzle flows - Comparison with data

Computational analyses have been made of the flow in NASA Langley's Arc-Heated Scramjet Test Facility's Mach 4.7 and Mach 6 square cross-section contoured nozzles, for comparison with experimental results. The analyses, which were performed using a three-dimensional RANS computer code assuming a single species gas with constant specific heats, were intended to provide insight into the nature of the flow development in this type of nozzle. The computational results showed the exit flow distribution to be affected by counter-rotating vortices along the centerline of each nozzle sidewall. Calculated flow properties show general, but not complete, agreement with experimental measurements in both nozzles.

Ostrander, Mark J.↗
NASA NTRS1989

Studies of vortex flow aerodynamics using CFD flow visualizations

Navier-Stokes computations of three-dimensional vortical flows over a round-edged double-delta wing and a tangent-ogive forebody are performed using an implicit upwind finite-difference scheme. Computed particle traces are compared with experimental oil-flow streaks.

Hsu, C.-H.↗
NASA NTRS1989

Mach 5 inlet CFD and experimental results

An experimental research program was conducted in the NASA Lewis Research Center 10 ft. by 10 ft. supersonic wind tunnel. The two-dimensional inlet model was designed to study the Mach 3.0 to 5.0 speed range for an 'over-under' turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both analytical code validation data as well as inlet performance information. Support studies for the program include flow field predictions with both three-dimensional parabolized Navier-Stokes (PNS) and three-dimensional full Navier-Stokes (FNS) analytical codes. Analytical predictions and experimental results are compared.

Weir, Lois J.↗
NASA NTRS1988

Implicit methods in CFD

A class of implicit approximate factorization schemes is examined for stability and convergence characteristics. These schemes include Newton's method, factorization, and flux-vector splitting. Examples are used to show that all practical methods suffer from some limited stability or asymptotic convergence restriction. A three-dimensional factored scheme is shown which suffers from unconditional instability which can only be ameliorated by added artificial dissipation. An F3D + or - flux split scheme is described which avoids unconditional instability, but in the end has similar convergence characteristics.

Pulliam, Thomas H.↗
NASA NTRS1990

Swept shock/boundary layer interaction experiments in support of CFD code validation

Research on the topic of shock wave/turbulent boundary layer interaction was carried out. Skin friction and surface pressure measurements in fin-induced, swept interactions were conducted, and heat transfer measurements in the same flows are planned. The skin friction data for a strong interaction case (Mach 4, fin-angles equal 16 and 20 degrees) were obtained, and their comparison with computational results was published. Surface pressure data for weak-to-strong fin interactions were also obtained.

Settles, G. S.↗
NASA NTRS1990

Unstructured finite element mesh generation and adaptive procedures for CFD

A method is described for generating unstructured meshes of triangles or tetrahedra for computational domains of complex geometrical shape. To illustrate the power of the approach, it is applied to the solution of flows past several complete aircraft configurations. The advocated approach allows for the natural incorporation of mesh adaptivity and this is demonstrated for both inviscid and viscous computations in two and three dimensions.

Peraire, J.↗
NASA NTRS1990

Laser skin friction measurements and CFD comparison of weak-to-strong swept shock/boundary-layer interactions

A joint experimental and computational study of skin friction in weak-to-strong swept shock wave/turbulent boundary-layer interactions has been carried out. A planar shock wave is generated by a sharp fin at angles of attack alpha = 10 deg and 16 deg at M(infinity) = 3 and 16 and 20 deg at M(infinity) = 4. Measurements are made using the Laser Interferometer Skin Friction meter, which optically detects the rate of thinning of an oil film applied to the test surface. The results show a systematic rise in the peak c(f) at the rear part of the interaction, where the separated flow atttaches. For the stronget case studied, this peak is an order of magnitude higher than the incoming freestream c(f)level.

Kim, K.-S.↗
NASA NTRS1988

CFD and transonic helicopter sound

A computational method which predicts far-field impulsive noise from a transonic motor blade is demonstrated. This method couples near-field results from a full-potential finite-difference method with a new Kirchhoff integral formulation to extend the finite-difference results into the acoustic far-field. This Kirchhoff formula is written in a blade-fixed coordinate system. It requires initial data from the potential code on a plane at the sonic radius. A recent hovering rotor experiment is described where accurate pressure measurements were recorded on the sonic cylinder and at 2 and 3 radii. The potential code prediction of sonic cylinder pressures is excellent. Acoustic far-field pressure predictions show good agreement with hover experimental data over the range of speeds from 0.85 to 0.92 tip Mach number, the latter of which have delocalized transonic flow. These results are some of the first successful predictions for peak pressure amplitudes using a computational code.

Purcell, Timothy W.↗
NASA NTRS1990

Angle-of-attack validation of a new zonal CFD method for airfoil simulations

The angle-of-attack validation of a new concept suggested by Summa (1990) for coupling potential and viscous flow methods has been investigated for two-dimensional airfoil simulations. The fully coupled potential/Navier-Stokes code, ZAP2D (Zonal Aerodynamics Program 2D), has been used to compute the flow field around an NACA 0012 airfoil for a range of angles of attack up to stall at a Mach number of 0.3 and a Reynolds number of 3 million. ZAP2D calculation for various domain sizes from 25 to 0.12 chord lengths are compared with the ARC2D large domain solution as well as with experimental data.

Yoo, Sungyul↗
NASA NTRS1990

Grid generation for computational CFD of full aerodynamic configurations

It is now possible to compute flow about very complex three-dimensional aerodynamic configurations. Grid generation is an essential part of such computations and a basic capability has been demonstrated. The present problems are: (1) choosing among the competing approaches (structured, unstructured or rectangular); (2) automating and packaging software for grid generation; (3) assessing and assuring grid quality; (4) solid boundary-surface representation and grid generation on solid boundaries; (5) automated domain decomposition (structured grids); (6) grid adaption during the flow solution process. This paper discusses some of the capabilities in the computation of complex three-dimensional grids and addresses some of he current challenges.

Smith, Robert E.↗
NASA NTRS1990

Application of CFD to sonic boom near and mid flow-field prediction

A 3-D parabolized Navier-Stokes (PNS) code was used to calculate the supersonic overpressures from three different geometries at near- and mid-flow fields. Wind tunnel data is used for code validation. Comparison of the computed results with different grid refinements is shown. It is observed that a large number of grid points is needed to resolve the tail shock/expansion fan interaction. Therefore, an adaptive grid approach is employed to calculate the flow field. The agreement between the numerical results and the wind tunnel data confirms that computational fluid dynamics can be applied to the problem of sonic boom prediction.

Cheung, Samson H.↗
NASA NTRS1989

CFD for applications to aircraft aeroelasticity

Strong interactions of structures and fluids are common in many engineering environments. Such interactions can give rise to physically important phenomena such as those occurring for aircraft due to aeroelasticity. Aeroelasticity can significantly influence the safe performance of aircraft. At present exact methods are available for making aeroelastic computations when flows are in either the linear subsonic or supersonic range. However, for complex flows containing shock waves, vortices and flow separations, computational methods are still under development. Several phenomena that can be dangerous and limit the performance of an aircraft occur due to the interaction of these complex flows with flexible aircraft components such as wings. For example, aircraft with highly swept wings experience vortex induced aeroelastic oscillations. Correct understanding of these complex aeroelastic phenomena requires direct coupling of fluids and structural equations. Here, a summary is presented of the development of such coupled methods and applications to aeroelasticity since about 1978 to present. The successful use of the transonic small perturbation theory (TSP) coupled with structures is discussed. This served as a major stepping stone for the current stage of aeroelasticity using computational fluid dynamics. The need for the use of more exact Euler/Navier-Stokes (ENS) equations for aeroelastic problems is explained. The current development of unsteady aerodynamic and aeroelastic procedures based on the ENS equations are discussed. Aeroelastic results computed using both TSP and ENS equations are discussed.

Guruswamy, Guru P.↗
NASA NTRS1989

CFD analysis for high speed inlets

The increased national interest in high speed flight has increased research for high speed propulsion components. The highly 3-D flows present in supersonic/hypersonic inlets are currently being studied at NASA-Lewis both experimentally and computationally using a family of steady Parabolized Navier-Stokes (PNS) and Navier-Stokes (NS) solvers and unsteady NS solvers. Some of the results of these efforts are presented with an emphasis on the comparison of the computational and experimental results. The flow in high speed inlets typically involves the interaction of compression shock waves and boundary layers on the internal surfaces. The fundamentals of these interactions have been studied experimentally for many years, while more recently, computations have been used to study these complex 3-D flow fields. Attempts to control the flow through boundary layer bleed are being investigated computationally prior to wind tunnel experiments. The ultimate goal is the higher performing inlets required for high speed flight.

Benson, Tom↗
NASA NTRS1991

A multidimensional finite element method for CFD

A finite element method is used to solve the equations of motion for 2- and 3-D fluid flow. The time-dependent equations are solved explicitly using quadrilateral (2-D) and hexahedral (3-D) elements, mass lumping, and reduced integration. A Petrov-Galerkin technique is applied to the advection terms. The method requires a minimum of computational storage, executes quickly, and is scalable for execution on computer systems ranging from PCs to supercomputers.

Pepper, Darrell W.↗
NASA NTRS1991

A CFD Study of Jet Mixing in Reduced Flow Areas for Lower Combustor Emissions

The Rich-burn/Quick-mix/Lean-burn (RQL) combustor has the potential of significantly reducing NO(x) emissions in combustion chambers of High Speed Civil Transport aircraft. Previous work on RQL combustors for industrial applications suggested the benefit of necking down the mixing section. A 3-D numerical investigation was performed to study the effects of neckdown on NO(x) emissions and to develop a correlation for optimum mixing designs in terms of neckdown area ratio. The results of the study showed that jet mixing in reduced flow areas does not enhance mixing, but does decrease residence time at high flame temperatures, thus reducing NO(x) formation. By necking down the mixing flow area by 4, a potential NO(x) reduction of 16:1 is possible for annual combustors. However, there is a penalty that accompanies the mixing neckdown: reduced pressure drop across the combustor swirler. At conventional combustor loading parameters, the pressure drop penalty does not appear to be excessive.

Smith, C. E.↗