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At least 523 records · Page 29

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NASA NTRS2014

Analysis of Low Speed Stall Aerodynamics of a Swept Wing with Laminar Flow Glove

Reynolds-Averaged Navier-Stokes (RANS) computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a GIII aircraft's swept wing modified with a laminar-flow wing glove. The stall aerodynamics of the gloved wing were analyzed and compared with the unmodified wing for the flight speed of 120 knots and altitude of 2300 ft above mean sea level (MSL). The Star-CCM+ polyhedral unstructured CFD code was first validated for wing stall predictions using the wing-body geometry from the First American Institute of Aeronautics and Astronautics (AIAA) CFD High-Lift Prediction Workshop. It was found that the Star-CCM+ CFD code can produce results that are within the scattering of other CFD codes considered at the workshop. In particular, the Star-CCM+ CFD code was able to predict wing stall for the AIAA wing-body geometry to within 1 degree of angle of attack as compared to benchmark wind-tunnel test data. Current results show that the addition of the laminar-flow wing glove causes the gloved wing to stall much earlier than the unmodified wing. Furthermore, the gloved wing has a different stall characteristic than the clean wing, with no sharp lift drop-off at stall for the gloved wing.

computational fluid dynamics↗

NASA NTRS1989

A three-dimensional upwind parabolized Navier-Stokes code for real gas flows

A real gas, upwind, parabolized Navier-Stokes (PNS) code has been developed to compute the three-dimensional hypersonic flow of equilibrium air around various body shapes. The new code is an extension of the upwind (perfect gas) PNS code of Lawrence et al. (1986). The upwind algorithm is based on Roe's (1981) flux-difference splitting scheme which has been modified to account for real gas effects using the nearly exact approach of Vinokur and Liu (1988). Simplified curve fits are employed to obtain the thermodynamic and transport properties of equilibrium air. The new code has been validated by computing the M-infinity = 25 laminar flow of air over cones at various angles of attack. The results of these computations are compared with the results from a conventional centrally-differenced, real gas PNS code and the previous axisymmetric, upwind, real gas code. The agreement is excellent in all cases.

Tannehill, John C.↗

NASA NTRS1988

An upwind parabolized Navier-Stokes code for real gas flows

A real gas, upwind, parabolized Navier-Stokes (PNS) code has been developed to compute the two-dimensional/axisymmetric hypersonic flow of equilibrium air around various body shapes. The new code is an extension of the upwind (perfect gas) PNS code of Lawrence, Tannehill and Chaussee. The upwind algorithm is based on Roe's flux-difference splitting scheme which has been modified to account for real gas effects. Simplified curve fits are used to obtain the thermodynamic and transport properties of equilibrium air. The new code has been validated by computing the hypersonic laminar flow of air over a flat plate, a wedge, a ramp, and a cone. The results of these computations are compared with the results from a conventional centrally-differenced, real gas, PNS code and the agreement is excellent, except in the vicinity of shock waves where the present code exhibits far superior shock capturing capabilities.

Tannehill, John C.↗

NASA NTRS1988

An upwind parabolized Navier-Stokes code for chemically reacting flows

A new upwind, parabolized Navier-Stokes (PNS) code has been developed to compute the hypersonic, viscous, chemically reacting flow around two-dimensional or axisymmetric bodies. The new code is an extension of the upwind (perfect gas) PNS code of Lawrence et al. (1986). The upwind algorithm is based on Roe's flux-difference splitting scheme which has been modified to account for real gas effects. The algorithm solves the gas dynamic and species continuity equations in a 'loosely' coupled manner. The new code has been validated by computing the laminar flow (at free stream Mach number 25) of chemically reacting air over a wedge and a cone. The results of these computations are compared with the results from a centrally-differenced, fully coupled, nonequilibrium PNS code. The agreement is excellent, except in the vicinity of the shock wave where the present code exhibits superior shock capturing capabilities.

Tannehill, John C.↗

NASA NTRS1998

Prediction of Unsteady Transitional Layers in Turbomachinery Using Navier Stokes Equations

The objective of the research reported in this presentation is to develop computational techniques for the prediction of unsteady transitional flows associated with the rotor stator interaction in turbomachinery. Three low-Reynolds number turbulence models are incorporated in two unsteady Navier-Stokes codes (one is pressure based and the other is time marching with Runge-Kutta time stepping) and evaluated for accuracy in predicting the onset and the end of unsteady transitional patches due to wake passing. The best model is then used for modification and improvement for the leading edge effect. An existing steady Navier-Stokes code was modified to include pseudo-time stepping, which provided acceleration from 5 to 25 times that of the original code. A systematic validation procedure was implemented to assess the effects of the grid, artificial dissipation, physical, and the pseudo-time step for an accurate prediction of transitional flows resulting from the rotor-stator interaction. The ability of the Navier-Stokes code to predict the unsteady transitional flow on a turbomachinery blade is demonstrated. The unsteady pressure and velocity fields are in good agreement with the experimental data and the prediction from the Euler/boundary layer approach. The numerical solver was able to capture all zones (wake induced transitional strip, wake induced turbulent strip, calmed region, etc.) associated with wake induced transition in a compressor cascade. Another significant step is the assessment of k-epsilon turbulence models, including the leading edge modifications. Best results were obtained from the FLB model. The LB model predicted earlier inception of the transition and shorter transition length. Modification of the k-epsilon model was found to be essential for an accurate prediction of the unsteady transitional flow in a compressor cascade. The CH model failed to predict the unsteady transitional flow. Predicted boundary layer was turbulent from the leading edge, even with the modification of the k-epsilon model near the stagnation point. A comparison of the instantaneous shape factor, the skin friction coefficient, and the momentum thickness indicates that the Navier-Stokes predictions are reasonably good. Interaction between the upstream wake and stator wake results in shedding of unsteady vortices from the trailing edge and increased dissipation in the stator wake and, as a consequence, increased rate of decay of the stator wake. The procedure developed at Penn State should aid the designers of turbomachinery in allowing for the unsteady transitional flows due to rotor-stator interaction.

Lakshminarayana, B.↗

NASA NTRS1981

A parametric study of staged fuel injector configurations for scramjet applications

A parametric study of staged (multiple) perpendicular fuel injector configurations has been conducted using a computer code which solves the two-dimensional elliptic Navier-Stokes equations. The program computes the turbulent mixing and reaction of hydrogen fuel and air and allows the study of separated regions of the flow immediately preceding and following the injectors. The validity of the code is demonstrated in a cold flow helium injection study with a single injector. Results are presented that describe the flow field near opposing staged injectors over a range of parameters. Parameters that are varied include injector size, fuel split, and distance between injectors. Comparisons of the configurations are made to assess their mixing and potential flame-holding qualities.

Weidner, E. H.↗

NASA NTRS1990

A validation study of the Spark Navier Stokes code for nonreacting scramjet combustor flowfields

A detailed study of Mach 2 flow over a rearward facing step and staged normal injection behind a rearward facing step into a Mach 2 mainflow has been performed using the Spark three-dimensional Navier-Stokes code. Very good overall agreement was observed for the flowfield without normal injection. The notable exceptions were the velocity for which the measured values showed more scatter and the temperature at the wall for which the value measured was below the adiabatic wall temperature. Faithful representation of the incoming boundary layer was found to be important in achieving good agreement. For the staged jet case, the calculated and measured flow quantities showed reasonably good agreement, with two significant discrepancies: the calculated jet trajectory is inclined more to the vertical then the measured one, and the measured temperatures downstream of the jets are lower than the computed ones. Possible reasons for the discrepancies are suggested.

Eklund, Dean R.↗

NASA NTRS1989

Computations of three-dimensional steady and unsteady viscous incompressible flows

The INS3D family of computational fluid dynamics computer codes is presented. These codes are used to as tools in developing and assessing algorithms for solving the incompressible Navier-Stokes equations for steady-state and unsteady flow problems. This work involves applying the codes to real-world problems involving complex three-dimensional geometries. The algorithms utilized include the method of pseudocompressibility including both central and upwind differencing, several types of artificial dissipation schemes, approximate factorization, and an implicit line-relaxation scheme. These codes have been validated using a wide range of problems including flow over a backward-facing step, driven cavity flow, flow through various types of ducts, and steady and unsteady flow over a circular cylinder. Many diverse flow applications have been solved using these codes including parts of the Space Shuttle Main Engine, problems in naval hydrodynamics, low-speed aerodynamics, and biomedical fluid flows. The presentation details several of these, including the flow through a Space Shuttle Main Engine inducer, vortex shedding behind a circular cylinder, and flow through an artificial heart.

Kwak, Dochan↗

NASA NTRS2012

Development of a Fluid Structures Interaction Test Technique for Fabrics

Application of fluid structures interaction (FSI) computational techniques to configurations of interest to the entry, descent and landing (EDL) community is limited by two factors - limited characterization of the material properties for fabrics of interest and insufficient experimental data to validate the FSI codes. Recently ILC Dover Inc. performed standard tests to characterize the static stress-strain response of four candidate fabrics for use in EDL applications. The objective of the tests described here is to address the need for a FSI dataset for CFD validation purposes. To reach this objective, the structural response of fabrics was measured in a very simple aerodynamic environment with well controlled boundary conditions. Two test series were undertaken. The first series covered a range of tunnel conditions and the second focused on conditions that resulted in fabric panel buckling.

Zilliac, Gregory G.↗

NASA NTRS1995

A computational fluid dynamics simulation of the hypersonic flight of the Pegasus(TM) vehicle using an artificial viscosity model and a nonlinear filtering method

The computational fluid dynamics code, PARC3D, is tested to see if its use of non-physical artificial dissipation affects the accuracy of its results. This is accomplished by simulating a shock-laminar boundary layer interaction and several hypersonic flight conditions of the Pegasus(TM) launch vehicle using full artificial dissipation, low artificial dissipation, and the Engquist filter. Before the filter is applied to the PARC3D code, it is validated in one-dimensional and two-dimensional form in a MacCormack scheme against the Riemann and convergent duct problem. For this explicit scheme, the filter shows great improvements in accuracy and computational time as opposed to the nonfiltered solutions. However, for the implicit PARC3D code it is found that the best estimate of the Pegasus experimental heat fluxes and surface pressures is the simulation utilizing low artificial dissipation and no filter. The filter does improve accuracy over the artificially dissipative case but at a computational expense greater than that achieved by the low artificial dissipation case which has no computational time penalty and shows better results. For the shock-boundary layer simulation, the filter does well in terms of accuracy for a strong impingement shock but not as well for weaker shock strengths. Furthermore, for the latter problem the filter reduces the required computational time to convergence by 18.7 percent.

Mendoza, John Cadiz↗

NASA NTRS1983

TAS: A Transonic Aircraft/Store flow field prediction code

A numerical procedure has been developed that has the capability to predict the transonic flow field around an aircraft with an arbitrarily located, separated store. The TAS code, the product of a joint General Dynamics/NASA ARC/AFWAL research and development program, will serve as the basis for a comprehensive predictive method for aircraft with arbitrary store loadings. This report described the numerical procedures employed to simulate the flow field around a configuration of this type. The validity of TAS code predictions is established by comparison with existing experimental data. In addition, future areas of development of the code are outlined. A brief description of code utilization is also given in the Appendix. The aircraft/store configuration is simulated using a mesh embedding approach. The computational domain is discretized by three meshes: (1) a planform-oriented wing/body fine mesh, (2) a cylindrical store mesh, and (3) a global Cartesian crude mesh. This embedded mesh scheme enables simulation of stores with fins of arbitrary angular orientation.

Thompson, D. S.↗

NASA NTRS2001

Unsteady Analysis of Separated Aerodynamic Flows Using an Unstructured Multigrid Algorithm

An implicit method for the computation of unsteady flows on unstructured grids is presented. The resulting nonlinear system of equations is solved at each time step using an agglomeration multigrid procedure. The method allows for arbitrarily large time steps and is efficient in terms of computational effort and storage. Validation of the code using a one-equation turbulence model is performed for the well-known case of flow over a cylinder. A Detached Eddy Simulation model is also implemented and its performance compared to the one equation Spalart-Allmaras Reynolds Averaged Navier-Stokes (RANS) turbulence model. Validation cases using DES and RANS include flow over a sphere and flow over a NACA 0012 wing including massive stall regimes. The project was driven by the ultimate goal of computing separated flows of aerodynamic interest, such as massive stall or flows over complex non-streamlined geometries.

Pelaez, Juan↗

NASA NTRS1991

A parametric study of the leading edge attachment line for the F-16XL

A three-dimensional Navier-Stokes code is used to computationally simulate the flow about a modified F-16XL. Transition mechanisms (e.g. attachment line location and crossflow instability) near the swept wing leading edge are analyzed in detail. Flow visualization is used to study the influence of angle-of-attack on the aforementioned transition mechanisms. Validation of the code is accomplished by comparison of numerically generated surface pressures with that obtained by in-flight experiments.

Flores, Jolen↗

NASA NTRS1987

Computational studies in low speed rotor aerodynamics

Recently, curved vortex elements have been developed for use in helicopter free wake calculations. As previously documented, these elements allow more efficient and accurate calculation of the wake-on-wake interactions than do traditional straight vortex elements. Curved elements have been applied here to the prediction of the rotor wake geometry and its associated inflow distribution in low speed forward flight. By coupling the wake model to a blade dynamic analysis, a simulation capable of predicting blade motion and hub moments in forward flight has been developed. The resulting code has been validated by successful predictions of the lateral flapping of articulated rotors in low speed level flight, and additional simulators of climb and descent have been undertaken. Successful correlation of other experimental data, including moments due to cyclic on hingeless rotors and transient thrust response in hover, has given encouraging evidence of the capabilities of this code as a tool for analysis of the low speed regime.

Quackenbush, Todd R.↗

NASA NTRS1999

Exploration of Uncertainty in Glacier Modelling

There are procedures and methods for verification of coding algebra and for validations of models and calculations that are in use in the aerospace computational fluid dynamics (CFD) community. These methods would be efficacious if used by the glacier dynamics modelling community. This paper is a presentation of some of those methods, and how they might be applied to uncertainty management supporting code verification and model validation for glacier dynamics. The similarities and differences between their use in CFD analysis and the proposed application of these methods to glacier modelling are discussed. After establishing sources of uncertainty and methods for code verification, the paper looks at a representative sampling of verification and validation efforts that are underway in the glacier modelling community, and establishes a context for these within overall solution quality assessment. Finally, an information architecture and interactive interface is introduced and advocated. This Integrated Cryospheric Exploration (ICE) Environment is proposed for exploring and managing sources of uncertainty in glacier modelling codes and methods, and for supporting scientific numerical exploration and verification. The details and functionality of this Environment are described based on modifications of a system already developed for CFD modelling and analysis.

Thompson, David E.↗

NASA NTRS1986

Unsteady transonic flow calculations for wing-fuselage configurations

The development of the XTRAN3S wing fuselage capability for predicting transonic unsteady aerodynamic loads for aeroplastic applications is described. The modifications of the XTRAN3S code, and XTRAN3S fuselage modeling based on an alternating-direction implicit finite-difference algorithm for solving a modified transonic small-disturbance equation, are analyzed. The XTRAN3S transonic code is applied to the steady and unsteady calculations for three wing-fuselage configurations and a transport-type wing-fuselage model. The effects of fuselage aerodynamic interference on the transonic unsteady air load and flutter characteristics of wings are investigated. Good correlation is detected for the calculated and experimental pressure distributions of the models confirming the validity of the code.

Batina, J. T.↗

NASA NTRS2001

Thermal Analysis for Orbiter and ISS Plume Impingement on International Space Station

The NASA Reaction Control System (RCS) Plume Model (RPM) is an exhaust plume flow field and impingement heating code that has been updated and applied to components of the International Space Station (ISS). The objective of this study was to use this code to determine if plume environments from either Orbiter PRCS jets or ISS reboost and Attitude Control System (ACS) jets cause thermal issues on ISS component surfaces. This impingement analysis becomes increasingly important as the ISS is being assembled with its first permanent crew scheduled to arrive by the end of fall 2000. By early summer 2001 , the ISS will have a number of major components installed such as the Unity (Node 1), Destiny (Lab Module), Zarya (Functional Cargo Block), and Zvezda (Service Module) along with the P6 solar arrays and radiators and the Z-1 truss. Plume heating to these components has been analyzed with the RPM code as well as additional components for missions beyond Flight 6A such as the Propulsion Module (PM), Mobile Servicing System, Space Station Remote Manipulator System, Node 2, and the Cupola. For the past several years NASA/JSC has been developing the methodology to predict plume heating on ISS components. The RPM code is a modified source flow code with capabilities for scarfed nozzles and intersecting plumes that was developed for the 44 Orbiter RCS jets. This code has been validated by comparison with Shuttle Plume Impingement Flight Experiment (SPIFEX) heat flux and pressure data and with CFD and Method of Characteristics solutions. Previous analyses of plume heating predictions to the ISS using RPM have been reported, but did not consider thermal analysis for the components nor jet-firing histories as the Orbiter approaches the ISS docking ports. The RPM code has since been modified to analyze surface temperatures with a lumped mass approach and also uses jet-firing histories to produce pulsed heating rates. In addition, RPM was modified to include plume heating from ISS jets to ISS components where the jet coordinates are specified, together with the engine cant angle. These latter studies have been focused on the PM with plumes from its reboost and ACS jets impinging on various ISS components and also focused on the Japanese H2 Transfer Vehicle (HTV) with the plumes from its reboost engines impinging on the Cupola window. This paper will present plume heating and surface temperature results on a number of ISS components with and without jet-firing histories, evaluate post-flight data, and describe any potential thermal issues

Rochelle, William C.↗

NASA NTRS1992

Design of a wing shape for study of hypersonic crossflow transition in flight

Computational fluid dynamics methods were used in the design of a wing shape for study of hypersonic crossflow transition in flight. The flight experiment is to be performed on the delta wing of the first stage of a Pegasus launch vehicle as a piggy-back experiment to support boundary-layer stability code development and validation. The design goal is to obtain crossflow-induced transition at 20-40 percent of the chord for a flight Mach number of approximately six. The present paper describes the design and analysis process utilized to obtain desired glove shape. A variety of schemes were used in the design, ranging from simple empirical crossflow correlations to three-dimensional Navier-Stokes codes in conjunction with linear stability/N-factor computations. The sensitivity to various parameters, such as trajectory variations, allowable wing thickness, leading-edge radius and surface temperature, is also discussed.

Godil, A.↗