ALBERT

Description: In the design of an airframe, the effect of changing the geometry on resulting computations is necessary for design optimization. The geometry is defined in terms of a series of design variables, including design variables to define the wing planform, tail, canard, pylon, and nacelle. Design optimization in this research is based on how these design variable affect the potential flow. The potential flow is computed as a function of the geometry and location of a series of panels describing the airframe, which are in turn a function of the design variables. Multipole accelerated panel methods improve the computational complexity of the problem and thus are an attractive approach. To utilize the methods in design optimization, it was necessary to define the appropriate sensitivity derivatives. The overhead incurred from finding the sensitivity derivatives in conjunction with the original computation should be small. This research developed the background for multipole-accelerated panel methods and the framework for finding sensitivity derivatives in the methods. Potential flow panel codes are commonly used for powered-lift aerodynamic predictions for three dimensional geometries. Given an airframe which has been discretized into a series of panels to define the airframe geometry, potential is computed as a function of the influence of all panels on all other panels. This is a computationally intensive problem for which efficient solutions are desired to improve the computational time and to allow greater resolution by use of more panels. One such solution is the use of hierarchical multipole methods which entail approximations of the effects of far-field terms. Hierarchical multipole methods have become prevalent in molecular dynamics and gravitational physics, and have been introduced into the fields of capacitance calculations, computational fluid dynamics, and electromagnetics. The methods utilize multipole expansions to describe the effect of bodies (i.e. particles, astrophysical bodies, panels, etc.) within a sphere on points distant from the sphere, where the influence diminishes as a function of distance. The expansions are exact with infinite series, however, for practical computations, the series are truncated and accuracy is selected based on the number of terms retained in the expansions. A hierarchical tree structure groups bodies together based on proximity to allow definition of multipole expansions for each group. The multipole expansions are then used to compute the effect of the bodies in a group on distant bodies.

Description: Over the past quarter century, the NASA Langley Research Center (LaRC) and the NASA Dryden Flight Research Center (DFRC) have played major roles in the development, demonstration, and validation of aeroservoelastic modeling, analysis, design, and testing methods. Many of their contributions resulted from their participation in wind-tunnel and flight-test programs aimed at demonstrating advanced active control concepts that interact with and/or exploit the aeroelastic characteristics of flexible structures. Other contributions are a result of their interest in identifying and solving adverse aeroservoelastic interactions that allow unique flight-test demonstrations or flight envelope clearance programs to be successfully completed. This paper provides an overview of some of the more interesting aeroservoelastic investigations conducted in the transonic dynamics tunnel (TDT) at LaRC and in flight at DFRC. Four flight-test projects are reviewed in this paper. These test projects were selected because of their contributions to the state-of-the-art in active controls technology (ACT) or because of the knowledge gained in further understanding the complex mechanisms that cause adverse aeroservoelastic interactions.

Description: The National Aeronautics and Space Administration (NASA) is conducting research to improve airport capacity by reducing the separation distance between aircraft. The limiting factor in reducing separation distances and improving airport capacity is the wake vortex hazard. The ability to accurately model wake vortices and predict the outcome of a vortex encounter is critical in developing a system to safely improve airport capacity. This is the focus of the wake vortex research being done at NASA Langley Research Center (LaRC). This paper will concentrate on two topics. The first topic is the control system developed for the Boeing 737 freeflight model in support of vortex encounter tests to be conducted in the 30- by 60- foot tunnel at NASA Langley Research Center later this year. The second topic discussed is the limited degree of freedom (DOF) trajectory generation study that is being conducted to determine the relative severity of a multitude of paths through a wake vortex.

Description: A thin-layer Navier-Stokes code, CFL3D, was utilized to compute the flow over a high-lift multi-element airfoil. This study was conducted to improve the prediction of high-lift flowfields using various turbulence models and improved glidding techniques. An overset Chimera grid system is used to model the three element airfoil geometry. The effects of wind tunnel wall modeling, changes to the grid density and distribution, and embedded grids are discussed. Computed pressure and lift coefficients using Spalart-Allmaras, Baldwin-Barth, and Menter's kappa-omega - Shear Stress Transport (SST) turbulence models are compared with experimental data. The ability of CFL3D to predict the effects on lift coefficient due to changes in Reynolds number changes is also discussed.

Description: The Federal Aviation Administration (FAA) and NASA have initiated a joint study in the development of reliable means of tracking, detecting, measuring, and predicting trailing wake-vortices of commercial aircraft. Being sought is an accurate model of the wake-vortex hazard, sufficient to increase airport capacity by reducing minimum safe spacings between planes. Several means of measurement are being evaluated for application to wake-vortex detection and tracking, including Doppler RADAR (Radio Detection and Ranging) systems, 2-micron Doppler LIDAR (Light Detection And Ranging) systems, and SODAR (Sound Detection And Ranging) systems. Of specific interest there is the lidar system, which has demonstrated numerous valuable capabilities as a vortex sensor Aerosols entrained in the vortex flow make the wake velocity signature visible to the lidar, (the observable lidar signal is essentially a measurement of the line-of-sight velocity of the aerosols). Measurement of the occurrence of a wake vortex requires effective reception and monitoring of the beat signal which results from the frequency-offset between the transmitted pulse and the backscattered radiation. This paper discusses the mounting, analysis, troubleshooting, and possible use of an analog processing assembly designed for such an application.

Description: The conversion of the Aerodynamic Preliminary Analysis System (APAS) software from a Silicon Graphics UNIX-based platform to a DOS-based IBM PC compatible is discussed. Relevant background information is given, followed by a discussion of the steps taken to accomplish the conversion and a discussion of the type of problems encountered during the conversion. A brief comparison of aerodynamic data obtained using APAS with data from another source is also made.

Description: A computer program that models the Lidar return signal for Wake Vortex experiments conducted by the Aerosol Research Branch was written. The specifications of the program and basic theory behind the calculations are briefly discussed. Results of the research and possible future improvements on it are also discussed.

Description: This report accounts details of two research projects for the Langley Aerospace Research Summer Scholars (LARSS) program. The first project, with the Office of Mission Assurance, involved subjectively predicting the probable success of two aeronautics programs by means of a tool called a Figure of Merit. The figure of merit bases program success on the quality and reliability of the following factors: parts, complexity of research, quality programs, hazards elimination, and single point failures elimination. The second project, for the Office of Safety and Facilities Assurance, required planning, layouts, and source seeking for an addition to the fire house. Forecasted changes in facility layout necessitate this addition which will serve as housing for the fire fighters.

Description: Using a comprehensive flight test database and a parameter identification software program produced at NASA Ames Research Center, a math model of the longitudinal aerodynamics of the Harrier aircraft was formulated. The identification program employed the equation error method using multiple linear regression to estimate the nonlinear parameters. The formulated math model structure adhered closely to aerodynamic and stability/control theory, particularly with regard to compressibility and dynamic manoeuvring. Validation was accomplished by using a three degree-of-freedom nonlinear flight simulator with pilot inputs from flight test data. The simulation models agreed quite well with the measured states. It is important to note that the flight test data used for the validation of the model was not used in the model identification.

Description: A general multi-block three-dimensional volume grid generator is presented which is suitable for Multi-Disciplinary Design Optimization. The code is fast, robust, highly automated, and written in ANSI C for platform independence. Algebraic techniques are used to generate and/or modify block face and volume grids to reflect geometric changes resulting from design optimization. Volume grids are generated/modified in a batch environment and controlled via an ASCII user input deck. This allows the code to be incorporated directly into the design loop. Generated volume grids are presented for a High Speed Civil Transport (HSCT) Wing/Body geometry as well a complex HSCT configuration including horizontal and vertical tails, engine nacelles and pylons, and canard surfaces.

Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the multiaxis thrust-vectoring characteristics of the F-18 High-Alpha Research Vehicle (HARV). A wingtip supported, partially metric, 0.10-scale jet-effects model of an F-18 prototype aircraft was modified with hardware to simulate the thrust-vectoring control system of the HARV. Testing was conducted at free-stream Mach numbers ranging from 0.30 to 0.70, at angles of attack from O' to 70', and at nozzle pressure ratios from 1.0 to approximately 5.0. Results indicate that the thrust-vectoring control system of the HARV can successfully generate multiaxis thrust-vectoring forces and moments. During vectoring, resultant thrust vector angles were always less than the corresponding geometric vane deflection angle and were accompanied by large thrust losses. Significant external flow effects that were dependent on Mach number and angle of attack were noted during vectoring operation. Comparisons of the aerodynamic and propulsive control capabilities of the HARV configuration indicate that substantial gains in controllability are provided by the multiaxis thrust-vectoring control system.

Description: A small scale ground effect test rig was used to study the ground plane flow field generated by a STOVL aircraft in hover. The objective of the research was to support NASA-Ames Research Center planning for the Large Scale Powered Model (LSPM) test for the ARPA-sponsored ASTOVL program. Specifically, small scale oil flow visualization studies were conducted to make a relative assessment of the aerodynamic interference of a proposed strut configuration and a wall configuration on the ground plane stagnation line. A simplified flat plate model representative of a generic jet-powered STOVL aircraft was used to simulate the LSPM. Cold air jets were used to simulate both the lift fan and the twin rear engines. Nozzle Pressure Ratios were used that closely represented those used on the LSPM tests. The flow visualization data clearly identified a shift in the stagnation line location for both the strut and the wall configuration. Considering the experimental uncertainty, it was concluded that either the strut configuration o r the wall configuration caused only a minor aerodynamic interference.

Description: An all-at-once reduced Hessian Successive Quadratic Programming (SQP) scheme has been shown to be efficient for solving aerodynamic design optimization problems with a moderate number of design variables. This paper extends this scheme to allow solution refining. In particular, we introduce a reduced Hessian refining technique that is critical for making a smooth transition of the Hessian information from coarse grids to fine grids. Test results on a nozzle design using quasi-one-dimensional Euler equations show that through solution refining the efficiency and the robustness of the all-at-once reduced Hessian SQP scheme are significantly improved.

Description: This paper introduces a computational scheme for solving a class of aerodynamic design problems that can be posed as nonlinear equality constrained optimizations. The scheme treats the flow and design variables as independent variables, and solves the constrained optimization problem via reduced Hessian successive quadratic programming. It updates the design and flow variables simultaneously at each iteration and allows flow variables to be infeasible before convergence. The solution of an adjoint flow equation is never needed. In addition, a range space basis is chosen so that in a certain sense the 'cross term' ignored in reduced Hessian SQP methods is minimized. Numerical results for a nozzle design using the quasi-one-dimensional Euler equations show that this scheme is computationally efficient and robust. The computational cost of a typical nozzle design is only a fraction more than that of the corresponding analysis flow calculation. Superlinear convergence is also observed, which agrees with the theoretical properties of this scheme. All optimal solutions are obtained by starting far away from the final solution.

Description: A design method has been developed by which an airfoil with a substantial amount of natural laminar flow can be designed, while maintaining other aerodynamic and geometric constraints. After obtaining the initial airfoil's pressure distribution at the design lift coefficient using an Euler solver coupled with an integml turbulent boundary layer method, the calculations from a laminar boundary layer solver are used by a stability analysis code to obtain estimates of the transition location (using N-Factors) for the starting airfoil. A new design method then calculates a target pressure distribution that will increase the larninar flow toward the desired amounl An airfoil design method is then iteratively used to design an airfoil that possesses that target pressure distribution. The new airfoil's boundary layer stability characteristics are determined, and this iterative process continues until an airfoil is designed that meets the laminar flow requirement and as many of the other constraints as possible.

Description: Design for prevention of aeroelastic instability (that is, the critical speeds leading to aeroelastic instability lie outside the operating range) is an integral part of the wing design process. Availability of the sensitivity derivatives of the various critical speeds with respect to shape parameters of the wing could be very useful to a designer in the initial design phase, when several design changes are made and the shape of the final configuration is not yet frozen. These derivatives are also indispensable for a gradient-based optimization with aeroelastic constraints. In this study, flutter characteristic of a typical section in subsonic compressible flow is examined using a state-space unsteady aerodynamic representation. The sensitivity of the flutter speed of the typical section with respect to its mass and stiffness parameters, namely, mass ratio, static unbalance, radius of gyration, bending frequency, and torsional frequency is calculated analytically. A strip theory formulation is newly developed to represent the unsteady aerodynamic forces on a wing. This is coupled with an equivalent plate structural model and solved as an eigenvalue problem to determine the critical speed of the wing. Flutter analysis of the wing is also carried out using a lifting-surface subsonic kernel function aerodynamic theory (FAST) and an equivalent plate structural model. Finite element modeling of the wing is done using NASTRAN so that wing structures made of spars and ribs and top and bottom wing skins could be analyzed. The free vibration modes of the wing obtained from NASTRAN are input into FAST to compute the flutter speed. An equivalent plate model which incorporates first-order shear deformation theory is then examined so it can be used to model thick wings, where shear deformations are important. The sensitivity of natural frequencies to changes in shape parameters is obtained using ADIFOR. A simple optimization effort is made towards obtaining a minimum weight design of the wing, subject to flutter constraints, lift requirement constraints for level flight and side constraints on the planform parameters of the wing using the IMSL subroutine NCONG, which uses successive quadratic programming.

Description: An investigation has been conducted in the Langley 7- by 10-Foot High Speed Wind Tunnel to determine the longitudinal and lateral directional aerodynamic characteristics of a series of personnel launch system concepts. This series of configurations evolved during an effort to improve the subsonic characteristics of a proposed lifting entry vehicle (designated the HL-20). The primary purpose of the overall investigation was to provide a vehicle concept which was inherently stable and trimable from entry to landing while examining methods of improving subsonic aerodynamic performance.

Description: A three-dimensional multiblock Navier-Stokes code, PAB3D, which was developed for propulsion integration and general aerodynamic analysis, has been used extensively by NASA Langley and other organizations to perform both internal (exhaust) and external flow analysis of complex aircraft configurations. This code was designed to solve the simplified Reynolds Averaged Navier-Stokes equations. A two-equation k-epsilon turbulence model has been used with considerable success, especially for attached flows. Accurate predicting of transonic shock wave location and pressure recovery in separated flow regions has been more difficult. Two algebraic Reynolds stress models (ASM) have been recently implemented in the code that greatly improved the code's ability to predict these difficult flow conditions. Good agreement with Direct Numerical Simulation (DNS) for a subsonic flat plate was achieved with ASM's developed by Shih, Zhu, and Lumley and Gatski and Speziale. Good predictions were also achieved at subsonic and transonic Mach numbers for shock location and trailing edge boattail pressure recovery on a single-engine afterbody/nozzle model.

Description: Preliminary measurements have been made of the flow over the tip of an unswept wing flap. To achieve an acceptable Reynolds number based on flap chord, the flap chord was chosen equal to the chord of the main airfoil (c = 19 in. approx. 0.48 m). The model was mounted in a 30 in. x 30 in. wind tunnel running at up to 100 ft/sec. (30 m/s): severe wind-tunnel interference was accepted, and any computations would be done using the tunnel walls as the boundaries of the computational domain. Maximum Reynolds number based on flap chord and tunnel speed was about 1.O x lO(exp 6). The grant ended before a full set of measurements could be made, but the work done so far yields a useful picture of the flow. The vortex originates at about mid-chord on the flap and rises rapidly above the chord line. It has a concentrated core, with total pressure lower than the ambient static pressure, and there is no evidence of large-scale wandering. A simple method of model construction, giving light weight and excellent surface finish, was developed.

Description: An all-at-once reduced Hessian Successive Quadratic Programming (SQP) scheme has been shown to be efficient for solving aerodynamic design optimization problems with a moderate number of design variables. This paper extends this scheme to allow solution refining. In particular, we introduce a reduced Hessian refining technique that is critical for making a smooth transition of the Hessian information from coarse grids to fine grids. Test results on a nozzle design using quasi-one-dimensional Euler equations show that through solution refining the efficiency and the robustness of the all-at-once reduced Hessian SQP scheme are significantly improved.

Description: Wind tunnel tests were made with a scale model of the HL-20 in the Langley Unitary Plan Wind Tunnel. Pitch control was investigated by deflecting the elevon surfaces on the outboard fins and body flaps on the fuselage. Yaw control tests were made with the all movable center fin deflected 5 deg. Almost full negative body flap deflection (-30 deg) was required to trim the HL-20 (moment reference center at 0.54-percent body length from nose) to positive values of life in the Mach number range from 1.6 to 2.5. Elevons were twice as effective as body flaps as a longitudinal trim device. The elevons were effective as a roll control, but because of tip-fin dihedral angle, produced about as much adverse yawing moment as rolling moment. The body flaps were less effective in producing rolling moment, but produced little adverse yawing moment. The yaw effectiveness of the all movable center fin was essentially constant over the angle-of-attack range at each Mach number. The value of yawing moment, however, was small. Center-fin deflection produced almost no rolling moments. The model was directionally unstable over most of the Mach number range with tip-fin dihedral angles less than the baseline value of 50 deg.

Description: 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.

Description: This presentation will examine the key performance aspects of shock tunnels as they relate to their use as aerothermodynamic flow simulation facilities. Assessment of shock tube reservoir conditions and flow contaminants generated in the shock tube will be presented along with their limiting impact on viable test envelopes, Facility nozzle performance as it pertains to test time assessment and nozzle exit flow quality (survey of pressure, temperature, and species) will be addressed. Also included will be a discussion of free stream flow diagnostics, both intrusive and nonintrusive, for measurement of critical flow properties not directly inferred from surface mounted transducers. The use of computational fluid dynamics for purposes of validating experimental measurements as well as predicting performance in regimes where measurements are not feasible or possible will be discussed. The use of CFD for facility research and design will also be presented.

Description: The selection of an airborne platform for the Stratospheric Observatory for Infrared Astronomy (SOFIA) is based not only on economic cost, but technical criteria, as well. Technical issues include aircraft fatigue, resonant characteristics of the cavity-port shear layer, aircraft stability, the drag penalty of the open telescope bay, and telescope performance. Recently, two versions of the Boeing 747 aircraft, viz., the -SP and -200 configurations, were evaluated by computational fluid dynamics (CFD) for their suitability as SOFIA platforms. In each configuration the telescope was mounted behind the wings in an open bay with nearly circular aperture. The geometry of the cavity, cavity aperture, and telescope was identical in both platforms. The aperture was located on the port side of the aircraft and the elevation angle of the telescope, measured with respect to the vertical axis, was 500. The unsteady, viscous, three-dimensional, aerodynamic and acoustic flow fields in the vicinity of SOFIA were simulated by an implicit, finite-difference Navier-Stokes flow solver (OVERFLOW) on a Chimera, overset grid system. The computational domain was discretized by structured grids. Computations were performed at wind-tunnel and flight Reynolds numbers corresponding to one free-stream flow condition (M = 0.85, angle of attack alpha = 2.50, and sideslip angle beta = 0 degrees). The computational domains consisted of twenty-nine(29) overset grids in the wind-tunnel simulations and forty-five(45) grids in the simulations run at cruise flight conditions. The maximum number of grid points in the simulations was approximately 4 x 10(exp 6). Issues considered in the evaluation study included analysis of the unsteady flow field in the cavity, the influence of the cavity on the flow across empennage surfaces, the drag penalty caused by the open telescope bay, and the noise radiating from cavity surfaces and the cavity-port shear layer. Wind-tunnel data were also available to compare to the CFD results; the data permitted an assessment of CFD as a design tool for the SOFIA program.

Description: Advanced hypersonic vehicles, like wave riders, will have sharp leading edges to minimize drag. These designs require accurate finite element modeling (FEM) of the thermal-structural behavior of a diboride ceramic matrix composite sharp leading edge. By coupling the FEM solver to an engineering model of the aerothermodynamic heating environment the impact of non catalytic surfaces, rarefied flow effects, and multidimensional conduction on the performance envelopes of sharp leading edges can be examined.

Description: The present paper shows under which assumptions one-equation models can be derived from two-equation models. Based on that transformation a new one-equation turbulence model is derived that basically behaves like a two-equation model. The new model is compared in detail against existing models

Description: The results of a wind-tunnel test are presented for a two-dimensional NASA 63(sub 2)-215 Mod B airfoil with a 30% chord single-slotted flap. The use of lift-enhancing tabs (similar to Gurney flaps) on the lower surface near the trailing edge of both elements was investigated on four nap configurations. A combination of vortex generators on the flap and lift-enhancing tabs was also investigated. Measurements of surface-pressure distributions and wake profiles were used to determine the aerodynamic performance of each configuration. By reducing flow separation on the flap, a lift-enhancing tab at the main-element trailing edge increased the maximum lift by 10.3% for the 42-deg flap case. The tab had a lesser effect at a moderate flap deflection (32 deg) and adversely affected the performance at the smallest flap deflection (22 deg). A tab located near the flap trailing edge produced an additional lift increment for all flap deflections. The application of vortex generators to the flap eliminated lift-curve hysteresis and reduced flow separation on two configurations with large flap deflections (greater than 40 deg). A maximum-lift coefficient of 3.32 (17% above the optimum baseline) was achieved with the combination of lift-enhancing tabs on both elements and vortex generators on the flap.

Description: The use of flat-plate tabs (similar to Gurney flaps) to enhance the lift of multielement airfoils is extended here by placing them on the pressure side and near the trailing edge of the main element rather than just on the furthest downstream wing element. The tabs studied range in height from 0.125 to 1.25% of the airfoil reference chord. In practice, such tabs would be retracted when the high-lift system is stowed. The effectiveness of the concept was demonstrated experimentally and computationally on a two-dimensional NACA 63(sub 2)-215 Mod B airfoil with a single-slotted, 30%-chord flap. Both the experiments and computations showed that the tabs significantly increase the lift at a given angle of attack and the maximum lift coefficient of the airfoil. The computational results showed that the increased lift was a result of additional turning of the flow by the tab that reduced or eliminated now separation on the flap. The best configuration tested, a 0.5%-chord tab placed 0.5% chord upstream of the trailing edge of the main element, increased the maximum lift coefficient of the airfoil by 12% and the maximum lift-to-drag ratio by 40%.

Description: Flight experiments on NASA Langley's B737-100 (TSRV) airplane have been conducted to document flow characteristics in order to further the understanding of high-lift flow physics, and to correlate and validate computational predictions and wind-tunnel measurements. The project is a cooperative effort involving NASA, industry, and universities. In addition to focusing on in-flight measurements, the project includes extensive application of various computational techniques, and correlation of flight data with computational results and wind-tunnel measurements. Results obtained in the most recent phase of flight experiments are analyzed and presented in this paper. In-flight measurements include surface pressure distributions, measured using flush pressure taps and pressure belts on the slats, main element, and flap elements; surface shear stresses, measured using Preston tubes; off-surface velocity distributions, measured using shear-layer rakes; aeroelastic deformations of the flap elements, measured using an optical positioning system; and boundary-layer transition phenomena, measured using hot-film anemometers and an infrared imaging system. The analysis in this paper primarily focuses on changes in the boundary-layer state that occurred on the slats, main element, and fore flap as a result of changes in flap setting and/or flight condition. Following a detailed description of the experiment, the boundary-layer state phenomenon will be discussed based on data measured during these recent flight experiments.

Description: This progress report, a series of viewgraphs, outlines experiments on the flow physics of confluent boundary layers for high lift systems. The design objective is to design high lift systems with improved C(sub Lmax) for landing approach and improved take-off L/D and simultaneously reduce acquisition and maintenance costs. In effect, achieve improved performance with simpler designs. The research objectives include: establish the role of confluent boundary layer flow physics in high-lift production; contrast confluent boundary layer structure for optimum and non-optimum C(sub L) cases; formation of a high quality, detailed archival data base for CFD/modeling; and examination of the role of relaminarization and streamline curvature.

Description: The objective is to understand supersonic laminar flow stability, transition, and active control. Some prediction techniques will be developed or modified to analyze laminar flow stability. The effects of distributed heating and cooling as an active boundary layer control technique will be studied. The primary tasks of the research apply to the NASA/Ames Proof of Concept (PoC) and Laminar Flow Supersonic Wind Tunnel's (LFSWT's) nozzle design with laminar flow control and are listed as follows: (1) predictions of supersonic laminar boundary layer stability and transition, (2) effects of wall heating and cooling on supersonic laminar flow control, (3) performance evaluation of the PoC and LFSWT nozzle designs with wall heating and cooling applied at different locations and various lengths, and (4) effects of a conducted versus pulse wall temperature distribution for the LFSWT.

Description: This progress report is a series of overviews outlining experiments on the flow physics of confluent boundary layers for high-lift systems. The research objectives include establishing the role of confluent boundary layer flow physics in high-lift production; contrasting confluent boundary layer structures for optimum and non-optimum C(sub L) cases; forming a high quality, detailed archival data base for CFD/modelling; and examining the role of relaminarization and streamline curvature. Goals of this research include completing LDV study of an optimum C(sub L) case; performing detailed LDV confluent boundary layer surveys for multiple non-optimum C(sub L) cases; obtaining skin friction distributions for both optimum and non-optimum C(sub L) cases for scaling purposes; data analysis and inner and outer variable scaling; setting-up and performing relaminarization experiments; and a final report establishing the role of leading edge confluent boundary layer flow physics on high-lift performance.

Description: The advantages and disadvantages of the blocked grid methodology are discussed using results from ENSAERO, CNS, and CNS-FV. The first two codes are based on finite differences and the last on cell-centered finite volume formulation. Techniques that enhance the utility of the blocked (or patched) grid methodology are described. These techniques include mesh discontinuous zonal interfaces, sliding zonal interfaces, fast search procedures, and virtual zones. All of these methods are designed with two goals; namely extend the use of patched grids to unsteady aerodynamics, e.g. oscillating control flaps, and provide the user more flexibility in the grid topologies available for gridding complex aerodynamic configurations. For example, the use of virtual zones allows the user the choice of using one grid topology for surface grids, and another for the volume grids. This additional flexibility has a large impact in the amount of calendar time required to block and grid a complex aerodynamic configuration. Several examples are shown demonstrating the new features. Other issues involving grid generation are also discussed. In particular the existing problems of defining grid quality measures which are relevant are also described.

Description: This paper will summarize recent progress in the numerical simulation of high incidence vortical flow about a generic 65 degree sweep delta wing using the three dimensional, time-dependent, Reynolds averaged, Navier-Stokes (RANS) equations. Computations have been carried out at 15 and 30 degrees angle of attack under subsonic turbulent flow conditions, and compared with experimental data provided by Hanff, Jenkins, and their colleagues. This work has already been published elsewhere and widely disseminated. Computations carried out at 15 degrees angle of attack included static roll angles ranging up to 65 degrees, and a large-amplitude (40 degree), high rate (7 Hz), forced roll motion. There was very good agreement between computed and experimental forces and moments, and static surface pressures. There was a significant hysteresis of the dynamic rolling moment due to the high rate of roll motion. At this angle of attack, no vortex breakdown was observed in the computations or experiment. Computations were also carried out at 30 degrees angle of attack, where vortex breakdown was present in both the computations and experiment. There was overall good agreement in the computed and experimental forces and moments. The static rolling moment varied with roll angle in a highly nonlinear manner, and exhibited three stable trim points and two unstable trim points. This behavior was attributed to the presence of vortex breakdown. Two large-amplitude (30 degrees), high-rate (10 Hz) forced roll motions were computed. The dynamics of the vortex breakdown motion was dramatically visualized by tracking the time-dependent motion of particles released near the delta wing apex. This numerical visualization is analogous to experimental smoke flow techniques. In one of the dynamic cases the breakdown was found to move off the wing, convected downwind of the trailing edge, and later reformed near the trailing edge through an instability of the vortex core. A damped free-to-roll motion was also computed by releasing the wing from rest at 40 degrees of roll. The wing went to the same trim point as in the experiment.

Description: Small radius leading edges and nosetips were utilized to minimize wave drag in early hypersonic vehicle concepts until further analysis demonstrated that extreme aerothermodynamic heating would cause severe ablation or blunting of the available thermal protection system materials. Recent studies indicate that diboride composite materials are shape stable under extreme aerothermodynamic heating at ultra high temperatures. Aerothermal performance envelopes for sharp components made from these materials are presented in this work to demonstrate the effects of convective blocking, surface catalycity, surface emissivity, and rarefied flow effects on steady state operation at altitudes from sea level to 90 km. These components are capable of steady state operation at velocities up to 7.9 km/s at altitudes near 90 km.

Description: Several countries, including the United States. Canada, Germany, England and Russia, are in the process of trying to develop some sort of computer-aided system that will guide controllers at airports on the hazard posed by lift-generated vortices that trail behind subsonic transport aircraft. The emphasis on this particular subject has come about because the hazard posed by wake vortices is currently the only reason why aircraft are spaced at 3 to 6 miles apart during landing and takeoff rather than something like 2 miles. It is well known that under certain weather conditions, aircraft spacings can be safely reduced to as little as the desired 2 miles. In an effort to perhaps capitalize on such a possibility, a combined FAA and NASA program is currently underway in the United States to develop such a system. Needless to say, the problems associated with anticipating the required separation distances when weather conditions are involved is very difficult. Similarly, Canada has a corresponding program to develop a vortex forecast system of their own.

Description: Flow field solutions over the Mars Pathfinder Probe spanning the trajectory through the Martian atmosphere at angles of attack from 0 to 11 degrees are obtained. Aerodynamic coefficients derived from these solutions reveal two regions where the derivative of pitching moment with respect to angle of attack is positive at small angles of attack. The behavior is associated with the transition of the sonic line location between the blunted nose and the windside shoulder of the 70 degree half-angle cone in a gas with a low effective ratio of specific heats. The transition first occurs as the shock layer gas chemistry evolves from highly nonequilibrium to near equilibrium, above approximately 6.5 km/s and 40 km altitude, causing the effective specific heat ratio to decrease. The transition next occurs in an equilibrium flow regime as velocities decrease through 3.5 km/s and the specific heat ratio increases again with decreasing enthalpy. The effects of the expansion over the shoulder into the wake are more strongly felt on the fustrum when the sonic line sits on the shoulder. The transition also produces a counter-intuitive trend in which windside heating levels decrease with increasing angle of attack resulting from an increase in the effective radius of curvature. Six-degree-of-freedom trajectory analyses utilizing the computed aerodynamic coefficients predict a moderate, 3 to 4 degree increase in total angle of attack as the probe, spinning at approximately 2 revolutions per minute, passes through these regions.

Description: For a computational flow simulation tool to be useful in a design environment, it must be very robust and efficient. To develop such a tool for incompressible flow applications, a number of different implicit schemes are compared for several two-dimensional flow problems in the current study. The schemes include Point-Jacobi relaxation, Gauss-Seidel line relaxation, incomplete lower-upper decomposition, and the generalized minimum residual method preconditioned with each of the three other schemes. The efficiency of the schemes is measured in terms of the computing time required to obtain a steady-state solution for the laminar flow over a backward-facing step, the flow over a NACA 4412 airfoil, and the flow over a three-element airfoil using overset grids. The flow solver used in the study is the INS2D code that solves the incompressible Navier-Stokes equations using the method of artificial compressibility and upwind differencing of the convective terms. The results show that the generalized minimum residual method preconditioned with the incomplete lower-upper factorization outperforms all other methods by at least a factor of 2.

Description: Unsteady Navier-Stokes simulations have been performed for vortical flows over an "arrow-wing" configuration of a supersonic transport in the transonic regime. Computed steady pressures and integrated force coefficients with and without control surface deflection at a moderate angle of attack are compared with experiment. For unsteady cases, oscillating trailing-edge control surfaces are modeled by using moving grids. Response characteristics between symmetric and antisymmetric oscillatory motions of the control surfaces on the left and right wings are studied. The antisymmetric case produces higher lift than the steady case with no deflection and the unsteady symmetric case produces higher lift than the antisymmetric case. The detailed analysis of the wake structure revealed a strong interaction between the primary vortex and the wake vortex sheet from the flap region when the flap is deflected up.

Description: An experimental study of laminar horseshoe vortex flows upstream of a cylinder/flat plate juncture has been conducted to verify the existence of saddle-point-of-attachment topologies. In the classical depiction of this flowfield, a saddle point of separation exists on the flat plate upstream of the cylinder, and the boundary layer separates from the surface. Recent computations have indicated that the topology may actually involve a saddle point of attachment on the surface and additional singular points in the flow. Laser light sheet flow visualizations have been performed on the symmetry plane and crossflow planes to identify the saddle-point-of-attachment flowfields. The visualizations reveal that saddle-point-of-attachment topologies occur over a range of Reynolds numbers in both single and multiple vortex regimes. An analysis of the flow topologies is presented that describes the existence and evolution of the singular points in the flowfield.

Description: The generation of significant side forces and yawing moments on an F/A-18 fuselage through tangential slot blowing is analyzed using computational fluid dynamics. The effects of freestream Mach number, jet exit conditions, jet length, and jet location are studied. The effects of over- and underblowing on force and moment production are analyzed. Non-time-accurate solutions are obtained to determine the steady-state side forces, yawing moments, and surface pressure distributions generated by tangential slot blowing. Time-accurate solutions are obtained to study the force onset time lag of tangential slot blowing. Comparison with available experimental data from full-scale wind-tunnel and subscale wind-tunnel tests are made. This computational analysis complements the experimental results and provides a detailed understanding of the effects of tangential slot blowing on the flowfield about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.

Description: A parametric study to predict the extent of laminar flow on the upper surface of a generic swept-back wing (NACA 64A010 airfoil section) at supersonic speeds was conducted. The results were obtained by using surface pressure predictions from an Euler/Navier-Stokes computational fluid dynamics code coupled with a boundary layer code, which predicts detailed boundary layer profiles, and finally with a linear stability code to determine the extent of laminar flow. The parameters addressed are Reynolds number, angle of attack, and leading-edge wing sweep. The results of this study show that an increase in angle of attack, for specific Reynolds numbers, can actually delay transition. Therefore, higher lift capability, caused by the increased angle of attack, as well as a reduction in viscous drag due to the delay in transition is possible for certain flight conditions.

Description: The sizing and efficiency of an aircraft is largely determined by the performance of its high-lift system. Subsonic civil transports most often use deployable multi-element airfoils to achieve the maximum-lift requirements for landing, as well as the high lift-to-drag ratios for take-off. However, these systems produce very complex flow fields which are not fully understood by the scientific community. In order to compete in today's market place, aircraft manufacturers will have to design better high-lift systems. Therefore, a more thorough understanding of the flows associated with these systems is desired. Flight and wind-tunnel experiments have been conducted on NASA Langley's B737-100 research aircraft to obtain detailed full-scale flow measurements on a multi-element high-lift system at various flight conditions. As part of this effort, computational aerodynamic tools are being used to provide preliminary flow-field information for instrumentation development, and to provide additional insight during the data analysis and interpretation process. The purpose of this paper is to demonstrate the ability and usefulness of a three-dimensional low-order potential flow solver, PMARC, by comparing computational results with data obtained from 1/8 scale wind-tunnel tests. Overall, correlation of experimental and computational data reveals that the panel method is able to predict reasonably well the pressures of the aircraft's multi-element wing at several spanwise stations. PMARC's versatility and usefulness is also demonstrated by accurately predicting inviscid three-dimensional flow features for several intricate geometrical regions.

Description: This paper summarizes a method that solves both the three dimensional thin-layer Navier-Stokes equations and the Euler equations using overset structured and solution adaptive unstructured grids with applications to helicopter rotor flowfields. The overset structured grids use an implicit finite-difference method to solve the thin-layer Navier-Stokes/Euler equations while the unstructured grid uses an explicit finite-volume method to solve the Euler equations. Solutions on a helicopter rotor in hover show the ability to accurately convect the rotor wake. However, isotropic subdivision of the tetrahedral mesh rapidly increases the overall problem size.

Description: As a helicopter transitions from hover to forward flight, the main rotor blades experience an asymmetry in flow field around the azimuth, with the blade section tangential velocities increasing on the advancing side and decreasing on the retreating side. To compensate for the reduced dynamic pressure on the retreating side, the blade pitch angles over this part of the rotor disk are increased. Eventually, a high enough forward speed is attained to produce compressibility effects on the advancing side of the rotor disk and stall on the retreating side. The onset of these two phenomena drastically increases the rotor vibratory loads and power requirements, thereby effectively establishing a limit on the maximum achievable forward speed. The alleviation of compressibility and stall (and the associated decrease in vibratory loads and power) would potentially result in an increased maximum forward speed. In the past, several methods have been examined and implemented to reduce the vibratory hub loads. Some of these methods are aimed specifically at alleviating vibration at very high flight speeds and increasing the maximum flight speed, while others focus on vibration reduction within the conventional flight envelope. Among the later are several types passive as well as active schemes. Passive schemes include a variety of vibration absorbers such as mechanical springs, pendulums, and bifilar absorbers. These mechanism are easy to design and maintain, but incur significant weight and drag penalties. Among the popular active control schemes in consideration are Higher Harmonic Control (HHC) and Individual Blade Control (IBC). HHC uses a conventional swash plate to generate a multi-cyclic pitch input to the blade. This requires actuators capable of sufficiently high power and bandwidth, increasing the cost and weight of the aircraft. IBC places actuators in the rotating reference frame, requiring the use of slip rings capable of transferring enough power to the actuators. Both schemes cause an increase in pitch link loads. Trailing Edge Flap (TEF) deployment can also used to generate unsteady aerodynamic forces and moments that counter the original vibratory loads, and thereby reduce rotor vibrations. While the vibrations absorbers, HHC, IBC, and TEF concepts discussed above attempt to reduce the vibratory loads, they do not specifically address the phenomena causing the vibrations at high advance ratios. One passive method that attempts to directly alleviate compressibility and stall, instead of reducing the ensuing vibrations, is the use of advanced tip designs. Taper, sweep, anhedral, and the manipulation of other geometric properties of the blade tips can reduce the severity of stall and compressibility effects , as well as reduce rotor power. A completely different approach to solve these problems is the tiltrotor configuration. As the forward velocity of the aircraft increases, the rotors, in this case, are tilted forward until they are perpendicular to the flow and act as propellers. This eliminates the edgewise flow encountered by conventional rotors and circumvents all the problems associated with flow asymmetry. However, the success involves a tremendous increase in cost and complexity of the aircraft. Another possible approach that has been proposed for the alleviation of vibratory loads at high forward flight speeds involves the use of controlled lead-lag motions to reduce the asymmetry in flow. A correctly phased 1/rev controlled lag motion could be introduced such that it produces a backward velocity on the advancing side and a forward velocity on the retreating side, to delay compressibility effects and stall to a higher advance ratio. Using a large enough lead-lag amplitude, the tip velocities could be reduced to levels encountered in hover. This concept was examined by two groups in the 1950's and early 1960's. In the United States, the Research Labs Division of United Aircraft developed a large lead-lag motion rotor, meant to achieve lag motion amplitudes up to 45 degrees. In order to reduce the required actuation force, the blade hinges were moved to 40% of the blade radius to increase the rotating lag frequency to approximately 1/rev. The blade hinges were redesigned to produce a flap-lag coupling so the large flapwise aerodynamic loads could be exploited to actuate the blades in the lag direction. A wind tunnel test of this rotor concept revealed actuation and blade motion scheduling problems. The project was eventually discontinued due to these problems and high blade stresses. Around the same time, at Boelkow in Germany, a similar lead-lag rotor program was conducted under the leadership of Hans Derschmidt. Here, too, the blade hinges were moved outboard to 34% radius to reduce the actuation loads. The main difference between this and the United Aircraft program was the use of a mechanical actuation scheme with maximum lead-lag motions of 400. This program was also discontinued for unclear reasons. The present study is directed toward conducting a comprehensive analytical examination to evaluate the effectiveness of controlled lead-lag motions in reducing vibratory hub loads and increasing maximum flight speed. Since both previous studies on this subject were purely experimental, only a limited data set and physical understanding of the problem was obtained. With the currently available analytical models and computational resources, the present effort is geared toward developing an in-depth physical understanding of the precise underlying mechanisms by which vibration reduction may be achieved. Additionally, in recognition of the fact that large amplitude lead-lag motions would - (i) be difficult to implement, and (ii) produce very large blade stresses; the present study examines the potential of only moderate-to-small lead-lag motions for reduction of vibratory hub loads. Using such an approach, the emphasis is not on eliminating the periodic variations in tangential velocity at the blade tip, but at best reducing these variations slightly so that compressibility and stall are delayed to slightly higher advance ratios. This study was conducted in two steps. In the first step, a hingeless helicopter rotor was modeled using rigid blades undergoing flap-lag-torsion rotations about spring restrained hinges and bearings. This model was then modified by separating the lead-lag degree of freedom into two components, a free and a prescribed motion. Using this model, a parametric study of the effect of phase and amplitude of a prescribed lead-lag motion on hub vibration was conducted. The data gathered was analyzed to obtain an understanding of the basic physics of the problem and show the capability of this method to reduce vibration and expand the flight envelope. In the second half of the study, the similar analysis was conducted using an elastic blade model to confirm the effects predicted by the simpler model.

Description: Results of a numerical study are presented for hypersonic low-density flow about a 70-deg blunt cone using direct simulation Monte Carlo (DSMC) and Navier-Stokes calculations. Particular emphasis is given to the effects of chemistry on the near-wake structure and on the surface quantities and the comparison of the DSMC results with the Navier-Stokes calculations. The flow conditions simulated are those experienced by a space vehicle at an altitude of 85 km and a velocity of 7 km/s during Earth entry. A steady vortex forms in the near wake for these freestream conditions for both chemically reactive and nonreactive air gas models. The size (axial length) of the vortex for the reactive air calculations is 25% larger than that of the nonreactive air calculations. The forebody surface quantities are less sensitive to the chemistry than the base surface quantities. The presence of the afterbody has no effect on the forebody flow structure or the surface quantities. The comparisons of DSMC and Navier-Stokes calculations show good agreement for the wake structure and the forebody surface quantities.

Description: The geometry of the velocity field in a numerically simulated incompressible turbulent boundary layer over a flat plate at Re theta=670 has been studied using the invariants of the velocity gradient tensor. These invariants are computed at every grid point in the flow and used to form the discriminant. Of primary interest are those regions in the flow where the discriminant is positive; regions where, according to the characteristic equation, the eigenvalues of the velocity gradient tensor are complex. An observer moving with a frame of reference which is attached to a fluid particle lying within such a region would see a local flow pattern of the type stable-focus-stretching or unstable-focus-compressing. When the flow is visualized this way, continuous, connected, large-scale structures are revealed that extend from the point just below the buffer layer out to the beginning of the wake region. These structures are aligned with the mean shear close to the wall and arch in the cross-stream direction away from the wall. In some cases the structures observed are very similar to to the hairpin eddy vision of boundary layer structure proposed by Theodorsen. That the structure of the flow is revealed more effectively by the discriminant rather than by the vorticity is important and adds support to recent observations of the discriminant in a channel flow simulation. Of particular importance is the fact that the procedure does not require the use of an arbitrary threshold in the discriminant. Further analysis using computer flow visualization shows a high degree of spatial correlation between regions of positive discriminant, extreme negative pressure fluctuations and large instantaneous values of Reynolds shear stress.

Description: Experimental longitudinal and lateral-directional aerodynamic characteristics were obtained for the Pegasus and Pegasus XL configurations over a Mach number range from 1.6 to 6 and angles of attack from -4 to +24 degrees. Angle of sideslip was varied from -6 to +6 degrees, and control surfaces were deflected to obtain elevon, aileron, and rudder effectiveness. Experimental data for the Pegasus configuration are compared with engineering code predictions performed by Nielsen Engineering & Research, Inc. (NEAR) in the aerodynamic design of the Pegasus vehicle, and with results from the Aerodynamic Preliminary Analysis System (APAS) code. Comparisons of experimental results are also made with longitudinal flight data from Flight #2 of the Pegasus vehicle. Results show that the longitudinal aerodynamic characteristics of the Pegasus and Pegasus XL configurations are similar, having the same lift-curve slope and drag levels across the Mach number range. Both configurations are longitudinally stable, with stability decreasing towards neutral levels as Mach number increases. Directional stability is negative at moderate to high angles of attack due to separated flow over the vertical tail. Dihedral effect is positive for both configurations, but is reduced 30-50 percent for the Pegasus XL configuration because of the horizontal tail anhedral. Predicted longitudinal characteristics and both longitudinal and lateral-directional control effectiveness are generally in good agreement with experiment. Due to the complex leeside flowfield, lateral-directional characteristics are not as well predicted by the engineering codes. Experiment and flight data are in good agreement across the Mach number range.

Description: Full Navier-Stokes equations were conducted to determine the feasibility of automating the control of wave instabilities within a flat plate boundary layer with sensors, actuators, and a spectral controller. The results indicate that a measure of wave cancellation can be obtained for small and large amplitude instabilities without feedback; however, feedback is required to optimize the control amplitude and phase for exact wave cancellation.

Description: A definitive measurement of the low-speed flight characteristics of waverider-based aircraft is required to augment the overall design database for this important class of vehicles which have great potential for efficient high-speed flight. Two separate waverider-derived vehicles were tested; one in the 14- by 22-Foot Tunnel and the other in the 12-Foot Low Speed Tunnel at Langley Research Center. These tests provided measurements of moments and forces about all three axes, control effectiveness, flow field characteristics and the effects of configuration changes. The results of these tunnel tests are summarized and the subsonic aerodynamic characteristics of the two configurations are shown.

Description: Unsteady flowfields of a two-dimensional oscillating airfoil are calculated using an implicit, finite-difference, Navier Stokes numerical scheme. Five widely used turbulence models are used with the numerical scheme to assess the accuracy and suitability of the models for simulating the retreating blade stall of helicopter rotor in forward flight. Three unsteady flow conditions corresponding to an essentially attached flow, light-stall, and deep-stall cases of an oscillating NACA 0015 wing experiment were chosen as test cases for computations. Results of unsteady airloads hysteresis curves, harmonics of unsteady pressures, and instantaneous flowfield patterns are presented. Some effects of grid density, time-step size, and numerical dissipation on the unsteady solutions relevant to the evaluation of turbulence models are examined. Comparison of unsteady airloads with experimental data show that all models tested are deficient in some sense and no single model predicts airloads consistently and in agreement with experiment for the three flow regimes. The chief findings are that the simple algebraic model based on the renormalization group theory (RNG) offers some improvement over the Baldwin Lomax model in all flow regimes with nearly same computational cost. The one-equation models provide significant improvement over the algebraic and the half-equation models but have their own limitations. The Baldwin-Barth model overpredicts separation and underpredicts reattachment. In contrast, the Spalart-Allmaras model underpredicts separation and overpredicts reattachment.

Description: The effect of tangential slot blowing on the flowfield about a generic chined forebody at high angles of attack is investigated numerically using solutions of the thin-layer, Reynolds-averaged, Navier-Stokes equations. The effects of jet mass now ratios, angle of attack, and blowing slot location in the axial and circumferential directions are studied. The computed results compare well with available wind-tunnel experimental data. Computational results show that for a given mass now rate, the yawing moments generated by slot blowing increase as the body angle of attack increases. It is observed that greater changes in the yawing moments are produced by a slot located closest to the lip of the nose. Also, computational solutions show that inboard blowing across the top surface is more effective at generating yawing moments than blowing outboard from the bottom surface.

Description: An important part of building mathematical models based on measured data is calculating the accuracy associated with statistical estimates of the model parameters. Indeed, without some idea of this accuracy, the parameter estimates themselves have limited value. In this work, an expression for computing quantitatively correct parameter accuracy measures for maximum likelihood parameter estimates with colored residuals is developed and validated. This result is important because experience in analyzing flight test data reveals that the output residuals from maximum likelihood estimation are almost always colored. The calculations involved can be appended to conventional maximum likelihood estimation algorithms. Monte Carlo simulation runs were used to show that parameter accuracy measures from the new technique accurately reflect the quality of the parameter estimates from maximum likelihood estimation without the need for correction factors or frequency domain analysis of the output residuals. The technique was applied to flight test data from repeated maneuvers flown on the F-18 High Alpha Research Vehicle (HARV). As in the simulated cases, parameter accuracy measures from the new technique were in agreement with the scatter in the parameter estimates from repeated maneuvers, while conventional parameter accuracy measures were optimistic.

Description: A two-dimensional computer code to solve the Burnett equations has been developed which computes the flow interaction between an exhausted plume and hypersonic external flow near the afterbody of a flight vehicle. This Burnett-2D code extends the capability of Navier-Stokes solver (RPLUS2D code) to include high-order Burnett source terms and slip-wall conditions for velocity and temperature. Higher-order Burnett viscous stress and heat flux terms are discretized using central-differencing and treated as source terms. Blocking logic is adopted in order to overcome the difficulty of grid generation. The computation of exhaust plume flow field is divided into two steps. In the first step, the thruster nozzle exit conditions are computed which generates inflow conditions in the base area near the afterbody. Results demonstrated that at high altitudes, the computations of nozzle exit conditions must include the effects of base flow since significant expansion exists in the base region. In the second step, Burnett equations were solved for exhaust plume flow field near the afterbody. The free stream conditions are set at an altitude equal to 80km and the Mach number is equal to 5.0. The preliminary results show that the plume expansion, as altitude increases, will eventually cause upstream flow separation.

Description: The paper will describe the Development of a general three-dimensional multiple grid zone Navier-Stokes flowfield simulation program (ENS3D-MPP) designed for efficient execution on the Intel Paragon Massively Parallel Processor (MPP) supercomputer, and the subsequent application of this method to the prediction of the viscous flowfield about the V-22 Osprey tiltrotor vehicle. The flowfield simulation code solves the thin Layer or full Navier-Stoke's equation - for viscous flow modeling, or the Euler equations for inviscid flow modeling on a structured multi-zone mesh. In the present paper only viscous simulations will be shown. The governing difference equations are solved using a time marching implicit approximate factorization method with either TVD upwind or central differencing used for the convective terms and central differencing used for the viscous diffusion terms. Steady state or Lime accurate solutions can be calculated. The present paper will focus on steady state applications, although time accurate solution analysis is the ultimate goal of this effort. Laminar viscosity is calculated using Sutherland's law and the Baldwin-Lomax two layer algebraic turbulence model is used to compute the eddy viscosity. The Simulation method uses an arbitrary block, curvilinear grid topology. An automatic grid adaption scheme is incorporated which concentrates grid points in high density gradient regions. A variety of user-specified boundary conditions are available. This paper will present the application of the scalable and superscalable versions to the steady state viscous flow analysis of the V-22 Osprey using a multiple zone global mesh. The mesh consists of a series of sheared cartesian grid blocks with polar grids embedded within to better simulate the wing tip mounted nacelle. MPP solutions will be shown in comparison to equivalent Cray C-90 results and also in comparison to experimental data. Discussions on meshing considerations, wall clock execution time, load balancing, and scalability will be provided.

Description: A former Martin Marietta Manned Space Systems engineer, Robert T. Thurman went from analyzing airloads on the Space Shuttle External Tank to analyzing airloads on golf balls for Wilson Sporting Goods Company. Using his NASA know-how, Thurman designed the Ultra 500 golf ball, which has three different-sized dimples in 60 triangular faces (instead of the usual 20) formed by a series of intersecting "parting" lines. This balances the asymmetry caused by the molding line in all golf balls. According to Wilson, the ball sustains initial velocity longer and produces the most stable ball flight for "unmatched" accuracy and distance.

Description: The spreading characteristics of jets from several asymmetric nozzles are studied in comparison to those of an axisymmetric jet, over the Mach number (M(sub J)) range of 0.3 to 1.96. The effect of tabs in two cases, the axisymmetric nozzle fitted with four tabs and a rectangular nozzle fitted with two large tabs, is also included in the comparison. Compared to the axisymmetric jet, the asymmetric jets spread only slightly faster at subsonic conditions, while at supersonic conditions, when screech occurs, they spread much faster. Screech profoundly increases the spreading of all jets. The effect varies in the different stages of screech, and the corresponding unsteady flowfield characteristics are documented via phase-averaged measurement of the fluctuating total pressure. An organization and intensification of the azimuthal vortical structures under the screeching condition is believed to be responsible for the increased spreading. Curiously, the jet from a 'lobed mixer' nozzle spreads much less at supersonic conditions compared to all other cases. This is due to the absence of screech with this nozzle. Jet spreading for the two tab configurations, on the other hand, is significantly more than any of the no-tab cases. This is true in the subsonic regime, as well as in the supersonic regime in spite of the fact that screech is essentially eliminated by the tabs. The dynamics of the streamwise vortex pairs produced by the tabs cause the most efficient jet spreading thus far observed in the study.