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At least 1,279 records · Page 71
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

NASA Wingtip Code Validation Grid

I have been contacted by Alan Celic to request the computational fluid dynamics (CFD) grid file for the NASA Wingtip CFD validation case. Alan is currently a Ph.D. student at the University of Stuttgard in Germany and spent a year here at Ames Research Center (ARC) as an Ames Associate. This case is a standard validation case studied by many within the U.S. The case is of the flow around a rectangular wing with a rounded wing tip. The airfoil is a NACA 0012. There is nothing special or unusual about the geometry or the grid file. The grid file is a single-zone grid with 2.5 million points. This geometry is generic and is not similar to any currently flying vehicle. All of the results published by NASA, as part of this study, (completed in 1995) are currently in the public domain.

Zilliac, Greg↗
NASA NTRS2000

Test Cases for a Rectangular Supercritical Wing Undergoing Pitching Oscillations

Steady and unsteady measured pressures for a Rectangular Supercritical Wing (RSW) undergoing pitching oscillations have been presented. From the several hundred compiled data points, 27 static and 36 pitching oscillation cases have been proposed for computational Test Cases to illustrate the trends with Mach number, reduced frequency, and angle of attack. The wing was designed to be a simple configuration for Computational Fluid Dynamics (CFD) comparisons. The wing had an unswept rectangular planform plus a tip of revolution, a panel aspect ratio of 2.0, a twelve per cent thick supercritical airfoil section, and no twist. The model was tested over a wide range of Mach numbers, from 0.27 to 0.90, corresponding to low subsonic flows up to strong transonic flows. The higher Mach numbers are well beyond the design Mach number such as might be required for flutter verification beyond cruise conditions. The pitching oscillations covered a broad range of reduced frequencies. Some early calculations for this wing are given for lifting pressure as calculated from a linear lifting surface program and from a transonic small perturbation program. The unsteady results were given primarily for a mild transonic condition at M = 0.70. For these cases the agreement with the data was only fair, possibly resulting from the omission of viscous effects. Supercritical airfoil sections are known to be sensitive to viscous effects (for example, one case cited). Calculations using a higher level code with the full potential equations have been presented for one of the same cases, and with the Euler equations. The agreement around the leading edge was improved, but overall the agreement was not completely satisfactory. Typically for low-aspect-ratio rectangular wings, transonic shock waves on the wing tend to sweep forward from root to tip such that there are strong three-dimensional effects. It might also be noted that for most of the test, the model was tested with free transition, but a few points were taken with an added transition strip for comparison. Some unpublished results of a rigid wing of the same airfoil and planform that was tested on the pitch and plunge apparatus mount system (PAPA) showed effects of the lower surface transition Strip on flutter at the lower subsonic Mach numbers. Significant effects of a transition strip were also obtained on a wing with a thicker supercritical section on the PAPA mount system. Both of these flutter tests on the PAPA resulted in very low reduced frequencies that may be a factor in this influence of the transition strip. However, these results indicate that correlation studies for RSW may require some attention to the estimation of transition location to accurately treat viscous effects. In this report several Test Cases are selected to illustrate trends for a variety of different conditions with emphasis on transonic flow effects. An overview of the model and tests is given and the standard formulary for these data is listed. Sample data points are presented in both tabular and graphical form. A complete tabulation and plotting of all the Test Cases is given. Only the static pressures and the real and imaginary parts of the first harmonic of the unsteady pressures are available. All the data for the test are available in electronic file form. The Test Cases are also available as separate electronic files.

Bennett, Robert M.↗
NASA NTRS2000

Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

In this project, we continue to develop the previous joint research between the Fluid Mechanics and Acoustics Laboratory (FM&AL) at Hampton University (HU) and the Jet Noise Team (JNT) at the NASA Langley Research Center (NASA LaRC). The FM&AL was established at Hampton University in June of 1996 and has conducted research under two NASA grants: NAG-1-1835 (1996-99), and NAG-1-1936 (1997-00). In addition, the FM&AL has jointly conducted research with the Central AeroHydrodynamics Institute (TsAGI, Moscow) in Russia under a Civilian Research and Development Foundation (CRDF) grant #RE2-136 (1996-99). The goals of the FM&AL programs are twofold: (1) to improve the working efficiency of the FM&AUs team in generating new innovative ideas and in conducting research in the field of fluid dynamics and acoustics, basically for improvement of supersonic and subsonic aircraft engines, and (2) to attract promising minority students to this research and training and, in cooperation with other HU departments, to teach them basic knowledge in Aerodynamics, Gas Dynamics, and Theoretical and Experimental Methods in Aeroacoustics and Computational Fluid Dynamics (CFD). The research at the HU FM&AL supports reduction schemes associated with the emission of engine pollutants for commercial aircraft and concepts for reduction of observables for military aircraft. These research endeavors relate to the goals of the NASA Strategic Enterprise in Aeronautics concerning the development of environmentally acceptable aircraft. It is in this precise area, where the US aircraft industry, academia, and Government are in great need of trained professionals and which is a high priority goal of the Minority University Research and Education (MUREP) Program, that the HU FM&AL can make its most important contribution. The main achievements for the reporting period in the development of concepts for noise reduction and improvement in efficiency for jet exhaust nozzles and inlets for aircraft engines are as follows: (1) Publications- The AIAA Paper #99-1924 has been presented at the 5th AIAA/CEAS Aeroacoustics Conference, May 10-12, 1999, Seattle, WA; the AIAA Paper #00-3315 has been accepted for the 36th AIAA/ASME/ SAE/ASEE Joint Propulsion Conference, 17-19 July, 2000, Huntsville, AL; and another paper has been accepted for the International Environmental Congress, 14-16 June, 2000, St.-Petersburg, Russia. (2) Two patents were granted on July 20, 1999, and January 12, 2000. (3) Three reports/presentations at the NASA LaRC and GRC (06/22199, 09/26/ 99, and 06/25/00). (4) Grants and Proposals: Four proposals were submitted to the NASA and CRDF; a NASA Faculty Award was granted on January, 2000. A CRDF Young Investigator Program Award was granted for a 3 months visit of the Russian scientist to the HU FM&AL (03/99-05/99). (5) Theory and Numerical Simulations- Analytical theory, numerical simulation, comparison of theoretical with experimental results, and modification of theoretical approaches, models, grids etc. have been conducted for several complicated 2D and 3D nozzle and inlet designs using NASA codes based on full Euler and Navier-Stokes solvers: CFL3D, CRAFT, GODUNOV, and others. New approach for environmental monitoring via infrasound. (6) Experimental Tests: Experimental acoustic tests at the TsAGI, Moscow, with nozzles having Screwdriver or Axisymmetric Plug and Permeable Shells. A small scale working model of the NASA Low Speed Wind Tunnel (LSWT) has been installed in the Experimental Hall of the HU FM&AL (June, 1999). Preliminary preparations for experimental tests were made. (7) Students Research Activity: Involvement of the two graduate students as research assistants in the current research project.

Gilinsky, Mikhail↗
NASA NTRS2000

On the Coupling of CDISC Design Method with FPX Rotor Code

A rotor section aerodynamics design package is developed by coupling Constrained Direct Iterative Surface Curvature (CDISC) design method with the FPX rotor code. The coupling between the CDISC design and the FPX flow analysis is fully automated. The CDISC design method employs a predictor-corrector procedure iteratively to determine a surface geometry which produces a target pressure distribution, where the target pressure distributions is either pre-defined or automatically generated through flow and geometry constraints. The FPX code is an eXtended Full-Potential rotor Computational Fluid Dynamics (CFD) code, which solves the three-dimensional unsteady full-potential equation in a strong conservative form using an implicit approximate factorization finite-difference scheme with entropy and viscosity corrections. Application of the CDISC design method coupled with the FPX rotor code is made for rotor blades in hovering motions. Several design examples are presented to demonstrate the capability of the new package in rotor section design.

Hu, Hong↗
NASA NTRS2000

Computational Analysis for Rocket-Based Combined-Cycle Systems During Rocket-Only Operation

A series of Reynolds-averaged Navier-Stokes calculations were employed to study the performance of rocket-based combined-cycle systems operating in an all-rocket mode. This parametric series of calculations were executed within a statistical framework, commonly known as design of experiments. The parametric design space included four geometric and two flowfield variables set at three levels each, for a total of 729 possible combinations. A D-optimal design strategy was selected. It required that only 36 separate computational fluid dynamics (CFD) solutions be performed to develop a full response surface model, which quantified the linear, bilinear, and curvilinear effects of the six experimental variables. The axisymmetric, Reynolds-averaged Navier-Stokes simulations were executed with the NPARC v3.0 code. The response used in the statistical analysis was created from Isp efficiency data integrated from the 36 CFD simulations. The influence of turbulence modeling was analyzed by using both one- and two-equation models. Careful attention was also given to quantify the influence of mesh dependence, iterative convergence, and artificial viscosity upon the resulting statistical model. Thirteen statistically significant effects were observed to have an influence on rocket-based combined-cycle nozzle performance. It was apparent that the free-expansion process, directly downstream of the rocket nozzle, can influence the Isp efficiency. Numerical schlieren images and particle traces have been used to further understand the physical phenomena behind several of the statistically significant results.

Steffen, C. J., Jr.↗
NASA NTRS2001

Aeroelastic Stability & Response of Rotating Structures

A summary of the work performed under NASA grant NCC3-605 is presented. More details can be found in the cited references. This grant led to the development of relatively faster aeroelastic analyses methods for predicting flutter and forced response in fans, compressors, and turbines using computational fluid dynamic (CFD) methods.

Keith, Theo G., Jr.↗
NASA NTRS2001

Evaluation of Turbulence-Model Performance in Jet Flows

The importance of reducing jet noise in both commercial and military aircraft applications has made jet acoustics a significant area of research. A technique for jet noise prediction commonly employed in practice is the MGB approach, based on the Lighthill acoustic analogy. This technique requires as aerodynamic input mean flow quantities and turbulence quantities like the kinetic energy and the dissipation. The purpose of the present paper is to assess existing capabilities for predicting these aerodynamic inputs. Two modern Navier-Stokes flow solvers, coupled with several modern turbulence models, are evaluated by comparison with experiment for their ability to predict mean flow properties in a supersonic jet plume. Potential weaknesses are identified for further investigation. Another comparison with similar intent is discussed by Barber et al. The ultimate goal of this research is to develop a reliable flow solver applicable to the low-noise, propulsion-efficient, nozzle exhaust systems being developed in NASA focused programs. These programs address a broad range of complex nozzle geometries operating in high temperature, compressible, flows. Seiner et al. previously discussed the jet configuration examined here. This convergent-divergent nozzle with an exit diameter of 3.6 inches was designed for an exhaust Mach number of 2.0 and a total temperature of 1680 F. The acoustic and aerodynamic data reported by Seiner et al. covered a range of jet total temperatures from 104 F to 2200 F at the fully-expanded nozzle pressure ratio. The aerodynamic data included centerline mean velocity and total temperature profiles. Computations were performed independently with two computational fluid dynamics (CFD) codes, ISAAC and PAB3D. Turbulence models employed include the k-epsilon model, the Gatski-Speziale algebraic-stress model and the Girimaji model, with and without the Sarkar compressibility correction. Centerline values of mean velocity and mean temperature are compared with experimental data.

Woodruff, S. L.↗
NASA NTRS1994

Hypersonic Nozzle/Afterbody Experiment: Flow Visualization and Boundary Layer Experiments

This study was conducted to experimentally characterize the flow field created by the interaction of a single-expansion-ramp-nozzle (SERN) flow with a hypersonic external stream Data were obtained from a generic nozzle/afterbody model in the 3.5-Foot Hypersonic Wind Tunnel of the NASA Ames Research Center in a cooperative experimental program involving Ames and the McDonnell Douglas Aerospace. The model design and test planning were performed in close cooperation with members of the Ames computational fluid dynamics (CFD) team for the National Aero-Space Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadowgraph flow-visualization photographs, afterbody surface-pressure distributions, boundary-layer rake measurements, and Preston-tube skin-friction measurements.

Keener, Earl R.↗
NASA NTRS2001

Supersonic Wind Tunnel Tests of a Half-axisymmetric 12 Deg-spike Inlet to a Rocket-based Combined-cycle Propulsion System

Results of an isolated inlet test for NASA's GTX air-breathing launch vehicle concept are presented. The GTX is a Vertical Take-off/ Horizontal Landing reusable single-stage-to-orbit system powered by a rocket-based combined-cycle propulsion system. Tests were conducted in the NASA Glenn 1- by 1-Foot Supersonic Wind Tunnel during two entries in October 1998 and February 1999. Tests were run from Mach 2.8 to 6. Integrated performance parameters and static pressure distributions are reported. The maximum contraction ratios achieved in the tests were lower than predicted by axisymmetric Reynolds-averaged Navier-Stokes computational fluid dynamics (CFD). At Mach 6, the maximum contraction ratio was roughly one-half of the CFD value of 16. The addition of either boundary-layer trip strips or vortex generators had a negligible effect on the maximum contraction ratio. A shock boundary-layer interaction was also evident on the end-walls that terminate the annular flowpath cross section. Cut-back end-walls, designed to reduce the boundary-layer growth upstream of the shock and minimize the interaction, also had negligible effect on the maximum contraction ratio. Both the excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls were identified as possible reasons for the discrepancy between the CFD predictions and the experiment. It is recommended that the centerbody spike and throat angles be reduced in order to lessen the induced pressure rise. The addition of a step on the cowl surface, and planar end-walls more closely approximating a plane of symmetry are also recommended. Provisions for end-wall boundary-layer bleed should be incorporated.

DeBonis, J. R.↗
NASA NTRS2001

ICEG2D (v2.0) - An Integrated Software Package for Automated Prediction of Flow Fields for Single-Element Airfoils With Ice Accretion

An integrated geometry/grid/simulation software package, ICEG2D, is being developed to automate computational fluid dynamics (CFD) simulations for single- and multi-element airfoils with ice accretions. The current version, ICEG213 (v2.0), was designed to automatically perform four primary functions: (1) generate a grid-ready surface definition based on the geometrical characteristics of the iced airfoil surface, (2) generate high-quality structured and generalized grids starting from a defined surface definition, (3) generate the input and restart files needed to run the structured grid CFD solver NPARC or the generalized grid CFD solver HYBFL2D, and (4) using the flow solutions, generate solution-adaptive grids. ICEG2D (v2.0) can be operated in either a batch mode using a script file or in an interactive mode by entering directives from a command line within a Unix shell. This report summarizes activities completed in the first two years of a three-year research and development program to address automation issues related to CFD simulations for airfoils with ice accretions. As well as describing the technology employed in the software, this document serves as a users manual providing installation and operating instructions. An evaluation of the software is also presented.

Thompson David S.↗
NASA NTRS2000

Aerothermodynamic Analysis of the Project FIRE II Afterbody Flow

35 years later, the Project FIRE II ballistic reentry to Earth at a nominal velocity of 11.4 km/s remains one of the best sources of heating data for the design of sample return capsules. The data from this flight experiment encompass both the thermochemical non-equilibrium and equilibrium flow regimes and include measurements of both radiative and total heating on the forebody and afterbody. Because of this, a number of researchers have performed computational fluid dynamics (CFD) simulations of the forebody of the FIRE II entry vehicle, with generally good results. In particular, Olynick et. al. coupled a Navier-Stokes solver (GIANTS) with a radiation code (NOVAR) and showed excellent agreement in surface heat transfer over the FIRE II trajectory between 1634 and 1651 seconds (77 km to 37 km). However, in most cases the primary motivation of the previous work was to understand and model the coupling between shock layer radiation and aerothermodynamics, and thus the simulations concentrated on the forebody flow only. To our knowledge there have been no prior published attempts to reproduce the afterbody heating data. However, an understanding of this data is critical to our efforts to design the next generation of Earth and planetary entry vehicles and to assess our need for additional flight data.

Wright, Micheal J.↗
NASA NTRS2001

Development of Advanced Casing Treatments for Flow Control

Under the Base R&T and Ultra Efficient Engine Technology programs, the NASA-Goddard Space Flight Center Compressor Branch is investigating flow control strategies required to increase the loading and efficiency of core compressors while maintaining current levels of operability. Flow-control strategies being studied include advanced casing treatments, wall jet injection, and blade-tip injection for compressor stability enhancement, directed jets for surface boundary layer control, and vortex-generating devices. The use of computational fluid dynamics (CFD) simulations to assess the effectiveness of flow-control devices and to guide their design is a key element in this research. CFD simulations serve to screen potential flow-control concepts at a lower cost than executing physical experiments in turbomachinery facilities. CFD simulations also provide guidance in designing physical experiments for those flow control concepts, which appear promising.

Keith, Theo G., Jr.↗
NASA NTRS2001

Acoustical and Flowfield Characterization of a Scaled Tabletop Rocket

An analysis of the acoustical and flowfield environment for the scaled 1-pound-force (lbf) thrust tabletop motor was performed. This tabletop motor from NASA Stennis Space Center Is composed of Plexiglas burning In gaseous oxygen with a graphite insert for the nozzle portion. The nozzle has a throat diameter of 0.2 inch and an exit diameter of 0.38 Inch. With a chamber pressure at 55 pounds per square Inch absolute (psia), a normal shock is formed immediately downstream of the nozzle exit plane as the combustion products exhaust into the ambient at atmospheric pressure. The jet characterization Is based on computational fluid dynamics (CFD) in conjunction with Kirchhoff surface integral formulation and compared with correlations developed for measured rocket noise and a pressure fluctuation scaling (PFS) method. Predictions and comparisons are made for the overall sound pressure levels (OASPL's) and spectral dependence of sound pressure level (SPL). The overall objective of this effort is to develop methods for scaling the acoustic and flowfield environment of rockets with a wide range of thrust (1 lbf to 1 million lbf).

Kandula, Max↗
NASA NTRS2001

Generic Wing-Body Aerodynamics Data Base

The wing-body aerodynamics data base consists of a series of CFD (Computational Fluid Dynamics) simulations about a generic wing body configuration consisting of a ogive-circular-cylinder fuselage and a simple symmetric wing mid-mounted on the fuselage. Solutions have been obtained for Nonlinear Potential (P), Euler (E) and Navier-Stokes (N) solvers over a range of subsonic and transonic Mach numbers and angles of attack. In addition, each solution has been computed on a series of grids, coarse, medium and fine to permit an assessment of grid refinement errors.

Holst, Terry L.↗
NASA NTRS1994

FAST User Guide

The Flow Analysis Software Toolkit, FAST, is a software environment for visualizing data. FAST is a collection of separate programs (modules) that run simultaneously and allow the user to examine the results of numerical and experimental simulations. The user can load data files, perform calculations on the data, visualize the results of these calculations, construct scenes of 3D graphical objects, and plot, animate and record the scenes. Computational Fluid Dynamics (CFD) visualization is the primary intended use of FAST, but FAST can also assist in the analysis of other types of data. FAST combines the capabilities of such programs as PLOT3D, RIP, SURF, and GAS into one environment with modules that share data. Sharing data between modules eliminates the drudgery of transferring data between programs. All the modules in the FAST environment have a consistent, highly interactive graphical user interface. Most commands are entered by pointing and'clicking. The modular construction of FAST makes it flexible and extensible. The environment can be custom configured and new modules can be developed and added as needed. The following modules have been developed for FAST: VIEWER, FILE IO, CALCULATOR, SURFER, TOPOLOGY, PLOTTER, TITLER, TRACER, ARCGRAPH, GQ, SURFERU, SHOTET, and ISOLEVU. A utility is also included to make the inclusion of user defined modules in the FAST environment easy. The VIEWER module is the central control for the FAST environment. From VIEWER, the user can-change object attributes, interactively position objects in three-dimensional space, define and save scenes, create animations, spawn new FAST modules, add additional view windows, and save and execute command scripts. The FAST User Guide uses text and FAST MAPS (graphical representations of the entire user interface) to guide the user through the use of FAST. Chapters include: Maps, Overview, Tips, Getting Started Tutorial, a separate chapter for each module, file formats, and system administration.

Walatka, Pamela P.↗
NASA NTRS1995

Optimization on a Network-based Parallel Computer System for Supersonic Laminar Wing Design

A set of Computational Fluid Dynamics (CFD) routines and flow transition prediction tools are integrated into a network based parallel numerical optimization routine. Through this optimization routine, the design of a 2-D airfoil and an infinitely swept wing will be studied in order to advance the design cycle capability of supersonic laminar flow wings. The goal of advancing supersonic laminar flow wing design is achieved by wisely choosing the design variables used in the optimization routine. The design variables are represented by the theory of Fourier series and potential theory. These theories, combined with the parallel CFD flow routines and flow transition prediction tools, provide a design space for a global optimal point to be searched. Finally, the parallel optimization routine enables gradient evaluations to be performed in a fast and parallel fashion.

Garcia, Joseph A.↗
NASA NTRS2001

Viscous Incompressible Flow Computations for 3-D Steady and Unsteady Flows

This viewgraph presentation gives an overview of viscous incompressible flow computations for three-dimensional steady and unsteady flows. Details are given on the use of computational fluid dynamics (CFD) as an engineering tool, solution methods for incompressible Navier-Stokes equations, numerical and physical characteristics of the primitive variable approach, and the role of CFD in the past and in current engineering and research applications.

Kwak, Dochan↗
NASA NTRS2002

Aeroheating Predictions for the X-43 Cowl-Closed Configuration at Mach 7 and 10

A pre-flight analysis was conducted for the Mach 7 and Mach 10 X-43 cowl-closed configurations to evaluate three-dimensional flow-field effects on localized heating in the vicinity of the closed cowl flap and sidewall. First, engineering-level analyses based on documented laminar and turbulent surface heating database were used to estimate localized heating amplification factors to account for turbulent flow effects, corner flow effects and the presence of forebody boundary layer trip devices. These estimated factors were applied to 2-D computational predictions to develop surface heat loads. Second, thin-film gauge surface heating data were obtained for a Mach 10 cowl-closed configuration in a reflected shock tunnel at simulated flight conditions. Third, computational fluid dynamics (CFD) flow-field predictions were obtained at Mach 10 ground test conditions to provide comparisons with the engineering database predictions and experimental measurements. Similarly, a CFD solution was also obtained at Mach 7 flight conditions to provide qualitative comparisons with engineering predictions. The Mach 10 CFD predictions compare favorably with the experimental measurements and the engineering estimates for surface heating amplification on the closed cowl. However, experimental measurements were not of sufficient quantity on the sidewall surface to capture the highest localized heating rates, and also the Mach 10 CFD predictions suggest higher peak heating on the sidewall compared to the engineering estimates. Also, the Mach 7 CFD predictions indicate qualitative agreement with the database estimates in terms of flow-field structure and peak heating locations on the closed-cowl flap and sidewall surfaces.

Cockrell, Charles E., Jr.↗