Engineering topics
Publications and source records attributed to Flores, Jolen.
An overview describing the current organizational structure, center mission, and technology focus areas of the Ames Research Center will be presented. The presentation will highlight research of the Aeronautics Directorate with some emphasis on activities related to fighter aircraft technology. The presentation will be limited to viewgraph material that is publicly available.
The viewgraphs and discussion on code validation for the simulation of supersonic viscous flow about the F-16XL are provided. Because of the large potential gains related to laminar flow on the swept wings of supersonic aircraft, interest in the applications of laminar flow control (LFC) techniques in the supersonic regime has increased. A supersonic laminar flow control (SLFC) technology program is currently underway within NASA. The objective of this program is to develop the data base and design methods that are critical to the development of laminar flow control technology for application to supersonic transport aircraft design. Towards this end, the program integrates computational investigations underway at NASA Ames-Moffett and NASA Langley with flight-test investigations being conducted on the F-16XL at the NASA Ames-Dryden Research Facility in cooperation with Rockwell International. The computational goal at NASA Ames-Moffett is to integrate a thin-layer Reynolds averaged Navier-Stokes flow solver with a stability analysis code. The flow solver would provide boundary layer profiles to the stability analysis code which in turn would predict transition on the F-16XL wing. To utilize the stability analysis codes, reliable boundary layer data is necessary at off-design cases. Previously, much of the prediction of boundary layer transition has been accomplished through the coupling of boundary layer codes with stability theory. However, boundary layer codes may have difficulties at high Reynolds numbers, of the order of 100 million, and with the current complex geometry in question. Therefore, a reliable code which solves the thin-layer Reynolds averaged Navier-Stokes equations is needed. Two objectives are discussed, the first in greater depth. The first objective is method verification, via comparisons of computations with experiment, of the reliability and robustness of the code. To successfully implement LFC techniques to the F-16XL wing, the flow about the leading edge must be maintained as laminar flow. Therefore, the second objective is to focus on a series of numerical simulations with different values of angle of attack, alpha, and Reynolds numbers. The purpose of the simulations is to study their effects on the two main factors which precipitate transition to turbulence at leading edges of highly swept wings (e.g., 'spanwise contamination' and 'crossflow instability').
Report describes numerical simulations of transonic flow of air about isolated wing and about wing and fuselage (not tail assembly) of F-16A aircraft. Fast, diagonalized Beam-Warming numerical-integration algorithm for solution of Euler and Navier-Stokes equations applied via Transonic Navier-Stokes computer program, using computational grids divided into zones of coarse and finer meshes.
Report describes computational study of transonic flows about airplane having shape similar to F-16A fighter. Navier-Stokes calculations simulate complicated flow fields and interactions.
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.
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Explore the source record for details and available documents.
Explore the source record for details and available documents.
Transonic Navier-Stokes (TNS) code solutions gathered from the literature for three-dimensional geometries, including two different wings and a complete F-16A aircraft, are presently discussed. The TNS codes use a zonal grid approach whose number of zones vary from four to 54. The Euler equations are solved in zones away from no-slip surfaces, and the thin-layer TNS equations are solved in all zones immediately adjacent to no-slip surfaces. In the case of 'corner' zones possessing no-slip boundary conditions on two different surfaces, a thin-layer formulation along both directions is employed. Employing these features, a zonal construction with the requisite set of boundary conditions can be devised for almost any application.
The thin layer, Reynolds-averaged, Navier-Stokes equations are used to simulate the transonic viscous flow about the complete F-16A fighter aircraft. These computations demonstrate how computational fluid dynamics can be used to simulate turbulent viscous flow about realistic aircraft geometries. A zonal grid approach is used to provide adequate viscous grid clustering on all aircraft surfaces. Zonal grids extend inside the F-16A inlet and up to the compressor face while power on conditions are modeled by employing a zonal grid extending from the exhaust nozzle to the far field. Computations are compared with existing experimental data and are in fair agreement. Computations for the F-16A in side slip are also presented.
CNS, a new computational fluid dynamics procedure, has been developed to aid in hypersonic vehicle design. The code can be used to model the entire external flow around hypersonic vehicle shapes, from the captured shock at the nose to the beginning of the wake. Unlike space-marching codes, the technique allows axially separated flow regions to be modeled. Validation trials using sphere-cone data reveal good solution accuracy for the surface pressure and flowfield temperature.
Computational fluid dynamics (CFD) research for hypersonic flows presents new problems in code validation because of the added complexity of the physical models. This paper surveys code validation procedures applicable to hypersonic flow models that include real gas effects. The current status of hypersonic CFD flow analysis is assessed with the Compressible Navier-Stokes (CNS) code as a case study. The methods of code validation discussed to beyond comparison with experimental data to include comparisons with other codes and formulations, component analyses, and estimation of numerical errors. Current results indicate that predicting hypersonic flows of perfect gases and equilibrium air are well in hand. Pressure, shock location, and integrated quantities are relatively easy to predict accurately, while surface quantities such as heat transfer are more sensitive to the solution procedure. Modeling transition to turbulence needs refinement, though preliminary results are promising.
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Problems encountered in numerical simulations of transonic wind-tunnel experiments with low-aspect-ratio wings are surveyed and illustrated. The focus is on the zonal Euler/Navier-Stokes program developed by Holst et al. (1985) and its application to shock-induced separation. The physical basis and numerical implementation of the method are reviewed, and results are presented from studies of the effects of artificial dissipation, boundary conditions, grid refinement, the turbulence model, and geometry representation on the simulation accuracy. Extensive graphs and diagrams and typical flow visualizations are provided.
The development of a Navier-Stokes code for the simulation of hypersonic flow will be described. The code uses a zonal philosophy which helps alleviate the grid generation problem for complex configurations and allows an efficient solution procedure. The code has been used to simulate equilibrium flow past a simple generic hypersonic configuration. Validation of the code on a blunt cone has also been conducted.
A numerical approach for simulating the viscous transonic flow about the complete F-16A fighter aircraft is presented using the Navier-Stokes equations. This finite difference approach utilizes a body conforming zonal grid system to provide appropriate viscous clustering near all body surfaces. A comparison between computational and experimental pressure coefficients is good, and integrated quantities such as lift and drag are within 2.6 percent and 1.6 percent, respectively. The versatility of the method is demonstrated by further modeling the flow inside the inlet up to the compressor face and the exhaust nozzle plume. Results for the F-16A in sideslip are also presented and indicate the proper trends.
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