ALBERT

Description: A NASA report detailing a wind tunnel investigation of a variable camber and twist could effectively reduce drag, thus improving performance. The resulting VooDoo fin is made of composite materials, has a rigid internal spar and a flexible polymer exterior coating. It is computer-designed and exceptionally durable.

Description: A dual-frequency acoustic levitator containing water was developed for studying bubble and drop dynamics in low gravity. It was flown on USML-1 where it was used in the Glovebox facility. High frequency (21 or 63 kHz) ultrasonic waves were modulated by low frequencies to excite shape oscillations on bubbles and oil drops ultrasonically trapped in the water. Bubble diameters were typically close to 1 cm or larger. When such large bubbles are acoustically trapped on the Earth, the acoustic radiation pressure needed to overcome buoyancy tends to shift the natural frequency for quadrupole (n = 2) oscillations above the prediction of Lamb's equation. In low gravity, a much weaker trapping force was used and measurements of n = 2 and 3 mode frequencies were closer to the ideal case. Other video observations in low gravity include: (i) the transient reappearance of a bulge where a small bubble has coalesced with a large one, (ii) observations of the dynamics of bubbles coated by oil indicating that shape oscillations can shift a coated bubble away from the oil-water interface of the coating giving a centering of the core, and (iii) the agglomeration of bubbles induced by the sound field.

Description: Any aircraft preliminary design study requires a structural model of the proposed configuration. The model must be capable of estimating the structural weight of a given configuration, and of predicting the deflections which will result from foreseen flight and ground loads. The present work develops such a model for the proposed Oblique All Wing airplane. The model is based on preliminary structural work done by Jack Williams and Peter Rudolph at Mdng, and is encoded in a FORTRAN program. As a stand-alone application, the program can calculate the weight CG location, and several types of structural deflections; used in conjunction with an aerodynamics model, the program can be used for mission analysis or sizing studies.

Description: The purpose of this investigation is to provide a comprehensive data base for the validation of numerical simulations. The objective of the present paper is to provide a tabulation of the experimental data. The data were obtained in the two-dimensional, transonic flowfield surrounding a supercritical airfoil. A variety of flows were studied in which the boundary layer at the trailing edge of the model was either attached or separated. Unsteady flows were avoided by controlling the Mach number and angle of attack. Surface pressures were measured on both the model and wind tunnel walls, and the flowfield surrounding the model was documented using a laser Doppler velocimeter (LDV). Although wall interference could not be completely eliminated, its effect was minimized by employing the following techniques. Sidewall boundary layers were reduced by aspiration, and upper and lower walls were contoured to accommodate the flow around the model and the boundary-layer growth on the tunnel walls. A data base with minimal interference from a tunnel with solid walls provides an ideal basis for evaluating the development of codes for the transonic speed range because the codes can include the wall boundary conditions more precisely than interference connections can be made to the data sets.

Description: The spatial evolution of cross flow-vortex packets in a laminar boundary layer on a swept wing are computed by the direct numerical simulation of the incompressible Navier- Stokes equations. A wall-normal velocity distribution of steady suction and blowing at the wing surface is used to generate a strip of equally spaced and periodic disturbances along the span. Three simulations are conducted to study the effect of initial amplitude on the disturbance evolution, to determine the role of traveling cross ow modes in transition, and to devise a correlation function to guide theories of transition prediction. In each simulation, the vortex packets first enter a chordwise region of linear independent growth, then, the individual packets coalesce downstream and interact with adjacent packets, and, finally, the vortex packets nonlinearly interact to generate inflectional velocity profiles. As the initial amplitude of the disturbance is increased, the length of the evolution to breakdown decreases. For this pressure gradient, stationary modes dominate the disturbance evolution. A two-coeffcient function was devised to correlate the simulation results. The coefficients, combined with a single simulation result, provide sufficient information to generate the evolution pattern for disturbances of any initial amplitude.

Description: Hybrid grids, composed of structured and unstructured grids, combines the best features of both. The chimera method is a major stepstone toward a hybrid grid from which the present approach is evolved. The chimera grid composes a set of overlapped structured grids which are independently generated and body-fitted, yielding a high quality grid readily accessible for efficient solution schemes. The chimera method has been shown to be efficient to generate a grid about complex geometries and has been demonstrated to deliver accurate aerodynamic prediction of complex flows. While its geometrical flexibility is attractive, interpolation of data in the overlapped regions - which in today's practice in 3D is done in a nonconservative fashion, is not. In the present paper we propose a hybrid grid scheme that maximizes the advantages of the chimera scheme and adapts the strengths of the unstructured grid while at the same time keeps its weaknesses minimal. Like the chimera method, we first divide up the physical domain by a set of structured body-fitted grids which are separately generated and overlaid throughout a complex configuration. To eliminate any pure data manipulation which does not necessarily follow governing equations, we use non-structured grids only to directly replace the region of the arbitrarily overlapped grids. This new adaptation to the chimera thinking is coined the DRAGON grid. The nonstructured grid region sandwiched between the structured grids is limited in size, resulting in only a small increase in memory and computational effort. The DRAGON method has three important advantages: (1) preserving strengths of the chimera grid; (2) eliminating difficulties sometimes encountered in the chimera scheme, such as the orphan points and bad quality of interpolation stencils; and (3) making grid communication in a fully conservative and consistent manner insofar as the governing equations are concerned. To demonstrate its use, the governing equations are discretized using the newly proposed flux scheme, AUSM+, which will be briefly described herein. Numerical tests on representative 2D inviscid flows are given for demonstration. Finally, extension to 3D is underway, only paced by the availability of the 3D unstructured grid generator.

Description: A new nonintrusive flow diagnostics instrumentation system, Doppler global velocimetry, is presented. The system is capable of making simultaneous, three-component velocity measurements within a selected measurement plane at video camera rates. These velocity images can provide the researcher with spatial and temporal information about the flow field in a global sense. The investigation of a vortical flow above a 75-degree delta wing comparing standard three-component, fringe-type laser velocimetry measurements with Doppler global velocimetry measurements is presented.

Description: An implicit, Navier-Stokes solution algorithm is presented for the computation of turbulent flow on unstructured grids. The inviscid fluxes are computed using an upwind algorithm and the solution is advanced in time using a backward-Euler time-stepping scheme. At each time step, the linear system of equations is approximately solved with a point-implicit relaxation scheme. This methodology provides a viable and robust algorithm for computing turbulent flows on unstructured meshes. Results are shown for subsonic flow over a NACA 0012 airfoil and for transonic flow over a RAE 2822 airfoil exhibiting a strong upper-surface shock. In addition, results are shown for 3 element and 4 element airfoil configurations. For the calculations, two one equation turbulence models are utilized. For the NACA 0012 airfoil, a pressure distribution and force data are compared with other computational results as well as with experiment. Comparisons of computed pressure distributions and velocity profiles with experimental data are shown for the RAE airfoil and for the 3 element configuration. For the 4 element case, comparisons of surface pressure distributions with experiment are made. In general, the agreement between the computations and the experiment is good.

Description: A technique is presented for triangulation of NURBS surfaces. This technique is built upon an advancing front technique combined with grid point projection. This combined approach has been successfully implemented for structured and unstructured grids.

Description: A three-dimensional computational fluid dynamics code, RPLUS3D, which was developed for the reactive propulsive flows of ramjets and scramjets, was validated for glancing shock wave-boundary layer interactions. Both laminar and turbulent flows were studied. A supersonic flow over a wedge mounted on a flat plate was numerically simulated. For the laminar case, the static pressure distribution, velocity vectors, and particle traces on the flat plate were obtained. For turbulent flow, both the Baldwin-Lomax and Chien two-equation turbulent models were used. The static pressure distributions, pitot pressure, and yaw angle profiles were computed. In addition, the velocity vectors and particle traces on the flat plate were also obtained from the computed solution. Overall, the computed results for both laminar and turbulent cases compared very well with the experimentally obtained data.

Description: An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Description: The NASA Langley 8-Foot Transonic Pressure Tunnel is a continuous-flow, variable-pressure wind tunnel with control capability to independently vary Mach number, stagnation pressure, stagnation temperature, and humidity. The top and bottom walls of the test section are axially slotted to permit continuous variation of the test section Mach number from 0.2 to 1.2, the slot-width contour provides a gradient-free test section 50 in. long for Mach numbers equal to or greater than 1.0 and 100 in. long for Mach numbers less than 1.0. The stagnation pressure may be varied from 0.25 to 2.0 atm. The tunnel test section has been recalibrated to determine the relationship between the free-stream Mach number and the test chamber reference Mach number. The hardware was the same as that of an earlier calibration in 1972 but the pressure measurement instrumentation available for the recalibration was about an order of magnitude more precise. The principal result of the recalibration was a slightly different schedule of reentry flap settings for Mach numbers from 0.80 to 1.05 than that determined during the 1972 calibration. Detailed tunnel contraction geometry, test section geometry, and limited test section wall boundary layer data are presented.

Description: An experiment was performed on oscillatory thermocapillary flow in the Glovebox aboard the USML-1 Spacelab which was launched in July, 1992. Cylindrical containers of 1 and 3 em in diameter were used. Silicone oils of 2 and 5 cSt viscosity were the test fluids. The fluid was heated by a cylindrical heater placed along the centerline of the container. The diameter of the heater was 10% of the container diameter. The fluid motion was studied by flow visualization. Although oscillations were observed briefly, bubbles generated in the fluid during the experiment disturbed the flow substantially so that the critical temperature differences could not be determined.

Description: The field of Computational Fluid Dynamics (CFD) has advanced to the point where it can now be used for many applications in fluid mechanics research and aerospace vehicle design. A few applications being explored at NASA Ames Research Center will be presented and discussed. The examples presented will range in speed from hypersonic to low speed incompressible flow applications. Most of the results will be from numerical solutions of the Navier-Stokes or Euler equations in three space dimensions for general geometry applications. Computational results will be used to highlight the presentation as appropriate. Advances in computational facilities including those associated with NASA's CAS (Computational Aerosciences) Project of the Federal HPCC (High Performance Computing and Communications) Program will be discussed. Finally, opportunities for future research will be presented and discussed. All material will be taken from non-sensitive, previously-published and widely-disseminated work.

Description: Line Integral Convolution (LIC) is a powerful technique for imaging and animating vector fields. We extend the LIC paradigm in three ways: (1) The existing technique is limited to vector fields over a regular Cartesian grid. We extend it to vector fields over parametric surfaces, such as those found in curvilinear grids, used in computational fluid dynamics simulations; (2) Periodic motion filters can be used to animate the flow visualization. When the flow lies on a parametric surface, however, the motion appears misleading. We explain why this problem arises and show how to adjust the LIC algorithm to handle it; (3) We introduce a technique to visualize vector magnitudes as well as vector direction. Cabral and Leedom have suggested a method for variable-speed animation, which is based on varying the frequency of the filter function. We develop a different technique based on kernel phase shifts which we have found to show substantially better results. Our implementation of these algorithms utilizes texture-mapping hardware to run in real time, which allows them to be included in interactive applications.

Description: The proposed paper will present a numerical investigation of the flow characteristics and boundary layer development in the nozzles of high enthalpy shock tunnel facilities used for hypersonic propulsion testing. The computed flow will be validated against existing experimental data. Pitot pressure data obtained at the entrance of the test cabin will be used to validate the numerical simulations. It is necessary to accurately model the facility nozzles in order to characterize the test article flow conditions. Initially the axisymmetric nozzle flow will be computed using a Navier Stokes solver for a range of reservoir conditions. The calculated solutions will be compared and calibrated against available experimental data from the DLR HEG piston-driven shock tunnel and the 16-inch shock tunnel at NASA Ames Research Center. The Reynolds number is assumed to be high enough at the throat that the boundary layer flow is assumed turbulent at this point downstream. The real gas affects will be examined. In high Mach number facilities the boundary layer is thick. Attempts will be made to correlate the boundary layer displacement thickness. The displacement thickness correlation will be used to calibrate the quasi-1D codes NENZF and LSENS in order to provide fast and efficient tools of characterizing the facility nozzles. The calibrated quasi-1D codes will be implemented to study the effects of chemistry and the flow condition variations at the test section due to small variations in the driver gas conditions.

Description: In a numerical flow simulation. it is common to generate several thousand time steps of unsteady (time-dependent) flow data. Each time step may require tens to hundreds of megabytes for disk storage, and the total disk requirement for storing the unsteady flow data may be hundreds of gigabytes. Interactive visualization of unsteady flow data of this magnitude is presently impossible with the current hardware technology. This chapter describes the current approaches for unsteady flow visualization. An effective particle tracing technique for unsteady flow is also described. First, the life cycle of a typical numerical flow simulation is outlined. Several unsteady flow data sets from real-world problems are then given. The current approaches for visualizing unsteady flow are then described. There are many existing systems for flow visualization, and some of them are discussed. Streaklines depict time-varying phenomena that are sometimes difficult or impossible to see with other visualization techniques. The algorithms for computing streaklines are described. Several unsteady flow data sets have been visualized using streaklines, and the results are presented. Finally, some current issues in unsteady flow visualization are discussed.

Description: We have implemented a three-dimensional compressible Navier-Stokes code on the Connection Machine CM-5. The code is set up for implicit time-stepping on single or multiple structured grids. For multiple grids and geometrically complex problems, we follow the 'chimera' approach, where flow data on one zone is interpolated onto another in the region of overlap. We will describe our design philosophy and give some timing results for the current code. A parallel machine like the CM-5 is well-suited for finite-difference methods on structured grids. The regular pattern of connections of a structured mesh maps well onto the architecture of the machine. So the first design choice, finite differences on a structured mesh, is natural. We use centered differences in space, with added artificial dissipation terms. When numerically solving the Navier-Stokes equations, there are liable to be some mesh cells near a solid body that are small in at least one direction. This mesh cell geometry can impose a very severe CFL (Courant-Friedrichs-Lewy) condition on the time step for explicit time-stepping methods. Thus, though explicit time-stepping is well-suited to the architecture of the machine, we have adopted implicit time-stepping. We have further taken the approximate factorization approach. This creates the need to solve large banded linear systems and creates the first possible barrier to an efficient algorithm. To overcome this first possible barrier we have considered two options. The first is just to solve the banded linear systems with data spread over the whole machine, using whatever fast method is available. This option is adequate for solving scalar tridiagonal systems, but for scalar pentadiagonal or block tridiagonal systems it is somewhat slower than desired. The second option is to 'transpose' the flow and geometry variables as part of the time-stepping process: Start with x-lines of data in-processor. Form explicit terms in x, then transpose so y-lines of data are in-processor. Form explicit terms in y, then transpose so z-lines are in processor. Form explicit terms in z, then solve linear systems in the z-direction. Transpose to the y-direction, then solve linear systems in the y-direction. Finally transpose to the x direction and solve linear systems in the x-direction. This strategy avoids inter-processor communication when differencing and solving linear systems, but requires a large amount of communication when doing the transposes. The transpose method is more efficient than the non-transpose strategy when dealing with scalar pentadiagonal or block tridiagonal systems. For handling geometrically complex problems the chimera strategy was adopted. For multiple zone cases we compute on each zone sequentially (using the whole parallel machine), then send the chimera interpolation data to a distributed data structure (array) laid out over the whole machine. This information transfer implies an irregular communication pattern, and is the second possible barrier to an efficient algorithm. We have implemented these ideas on the CM-5 using CMF (Connection Machine Fortran), a data parallel language which combines elements of Fortran 90 and certain extensions, and which bears a strong similarity to High Performance Fortran. We make use of the Connection Machine Scientific Software Library (CMSSL) for the linear solver and array transpose operations.

Description: Steady and unsteady flows for propulsion systems are efficiently simulated by solving the incompressible Navier-Stokes equations. The solution method is based on the pseudo compressibility approach and uses an implicit-upwind differencing scheme together with the Gauss-Seidel line relaxation method. Current computations use one equation Baldwin-Barth turbulence model which is derived from a simplified form of the standard kappa - epsilon model equations. The resulting computer code is applied to the flow analysis inside an advanced rocket pump impeller in steadily rotating reference frames. Numerical results are compared with experimental measurements. The effects of exit and shroud cavities with the leak-age flow are investigated. Time-accurate incompressible Navier-Stokes formulation with the overlapped grid scheme capability was evaluated by using MIT flapping foil experiment. The grid dependency, turbulence model effects, and the effect of order of differencing were investigated. Numerical results were compared against experimental data. The resulting procedure were applied to unsteady flapping foil calculations. Two upstream NACA 0025 foils perform high frequency synchronized motion and generate unsteady flow conditions to the downstream larger stationary foil. Comparison between unsteady experimental data and numerical results from two different moving boundary procedures will be presented.

Description: A numerical investigation is carried out to determine the magnitude of wake radiation for a proposed Venus composition probe. One of the scientific goals of the mission is to determine the atmospheric composition of Venus by examining the intensity of scattered sunlight through the wake of the vehicle during planetary entry. In the wake of the vehicle, excited particles generated in the bow shock and boundary layers absorb and emit radiation. Thus, the purpose of this study is to determine if the radiation sensor will be able to sense the incoming solar radiative flux relative to the radiative flux generated in the wake. During portions of the entry trajectory the incident surface heat flux will be high enough to produce significant ablation. Ablation products such as CN are known to be strong radiators. Also, the ablation will be driven by strong radiation emanating from the bow shock. Thus, radiation and ablation will be coupled into the Navier-Stokes flow solutions.

Description: The diffusion controlled binary reaction between initially segregated reactants in a two-dimensional low Mach number mixing layers is studied via numerical simulation. The stoichiometric ratio of the reactants is chosen to be much larger than one, as is typical of hydrocarbon flames in air. This results is a flame that is offset from the main vortical region of the mixing layer. In agreement with experimental observations, the flame remains surprisingly uncontorted during the flow evolution and is not entrained into the mixing layer. The effect of the heat release of the flame on the evolution of the mixing layer is thus similar to the effect of a difference in free-stream density between the two sides of the layer. The resulting baroclinic torque inhibits the familiar rolup and pairing of mixing layer vortices common in constant density flows. This also contributes to the layers inability to entrain the flame. The increase in viscosity caused by the heating of the flame reduces the effective Reynolds number of the flow. But, contrary to what has commonly been suggested, this is not the major reason for the inhibition of the usual large-scale mixing layer structures.

Description: Computational fluid dynamic (CFD) analysis is performed on the Lockheed Lifting Body Single-Stage-to-Orbit vehicle to determine the heat transfer to the vehicle during its descent trajectory. Seven species, chemical nonequilibriurn computations using the GASP code will be completed at several trajectory points to assess the thermal protection requirements of the vehicle. Sophisticated surface boundary conditions including in-depth conduction, catalycity, and a variable temperature wall have been incorporated into the flow solver.

Description: MIT flapping foil experiment was used as a validation case to evaluate the current incompressible Navier-Stokes approach with overlapped grid schemes. Steady-state calculations were carried out for overlapped and patched grids. The grid dependency, turbulence model effects, and the effect of order of differencing were investigated. Numerical results were compared against experimental data. The resulting procedure were applied to unsteady flapping foil calculations. Two upstream NACA 0025 foils perform high-frequency synchronized motion and generate unsteady flow conditions to the downstream larger stationary foil. Comparison between unsteady experimental data and numerical results from two different moving boundary procedures will be presented.

Description: An adiabatic demagnetization refrigerator (ADR) is under development at NASA-Ames Research Center that will operate between 2 K and 10 K and will provide 50 mW of cool ng at 2 K. Gadolinium Gallium Garnet (GGG) is selected as the refrigerant for the ADR, To minimize temperature gradients in the GGG, thick slices of GGG are sandwiched together with strips of high-purity copper in between them. The copper strips are used to exchange heat between the GGG and the 2 K and the 10 K heat switches. The heat transfer across the Cu-GGG interfaces is improved by placing thin foils of' high-purity indium at the interfaces. The heat switches employed in the ADR have no moving parts. The 10 K heat switch is a helium gas-gap heat switch; while, the 2 K heat switch is a He ll-gap heat switch. A switch is on when its gap Is filled with helium and is off' when the gap is emptied. This is accomplished with an activated carbon pump (ACP). The ACP adsorbs helium when cooled and desorbs it when heated. A superconducting magnet capable of providing 9 T at 2 K is used for the ADR cycle. A prototype of this refrigerator has been built and is currently under test. A detailed design of the ADR and preliminary test results performed on the prototype ADR will be presented.

Description: Fluid dynamics of turbomachines are complicated by inherently three dimensional structures such as endwall boundary layers, hub corner separation bubbles and tip-leakage flows. In addition, the relative motion between rotors and stators causes unsteady aerodynamic interactions to occur between blade rows. It is necessary to understand the aerodynamics associated with these interactions in order to design turbomachines that are both light and compact as well as reliable and efficient. An unsteady, three-dimensional, thin-layer, Navier-Stokes zonal algorithm is used to investigate the unsteady aerodynamics of multi-stage turbines and compressors. Relative motion between rotors and stators is made possible by the use of systems of patched and overlaid grids. Time-averaged pressures and pressure envelopes have been computed for several two- and three-dimensional single- and multi-stage configurations. Flow visualizations and computed results are in good agreement with experimental data.

Description: Based on the geometry of Mars Environment Survey (MESUR) Pathfinder aeroshell and an estimated Mars entry trajectory, two-dimensional axisymmetric time dependent calculations have been obtained using GIANTS (Gauss-Siedel Implicit Aerothermodynamic Navier-Stokes code with Thermochemical Surface Conditions) code and CMA (Charring Material Thermal Response and Ablation) Program for heating analysis and heat shield material sizing. These two codes are interfaced using a loosely coupled technique. The flowfield and convective heat transfer coefficients are computed by the GIANTS code with a species balance condition for an ablating surface, and the time dependent in-depth conduction with surface blowing is simulated by the CMA code with a complete surface energy balance condition. In this study, SLA-561V has been selected as heat shield material. The solutions, including the minimum heat shield thicknesses over aeroshell forebody, pyrolysis gas blowing rates, surface heat fluxes and temperature distributions, flowfield, and in-depth temperature history of SLA-561V, are presented and discussed in detail.

Description: Up to today, preconditioning methods on massively parallel systems have faced a major difficulty. The most successful preconditioning methods in terms of accelerating the convergence of the iterative solver such as incomplete LU factorizations are notoriously difficult to implement on parallel machines for two reasons: (1) the actual computation of the preconditioner is not very floating-point intensive, but requires a large amount of unstructured communication, and (2) the application of the preconditioning matrix in the iteration phase (i.e. triangular solves) are difficult to parallelize because of the recursive nature of the computation. Here we present a new approach to preconditioning for very large, sparse, unsymmetric, linear systems, which avoids both difficulties. We explicitly compute an approximate inverse to our original matrix. This new preconditioning matrix can be applied most efficiently for iterative methods on massively parallel machines, since the preconditioning phase involves only a matrix-vector multiplication, with possibly a dense matrix. Furthermore the actual computation of the preconditioning matrix has natural parallelism. For a problem of size n, the preconditioning matrix can be computed by solving n independent small least squares problems. The algorithm and its implementation on the Connection Machine CM-5 are discussed in detail and supported by extensive timings obtained from real problem data.

Description: The ability to control the extent of laminar flow on swept wings at supersonic speeds may be a critical element in developing the enabling technology for a High Speed Civil Transport (HSCT). Laminar boundary layers are less resistive to forward flight than their turbulent counterparts, thus the farther downstream that transition from laminar to turbulent flow in the wing boundary layer is extended can be of significant economic impact. Due to the complex processes involved experimental studies of boundary layer stability and transition are needed, and these are performed in "quiet" wind tunnels capable of simulating the low-disturbance environment of free flight. At Ames, a wind tunnel has been built to operate at flow conditions which match those of the HSCT laminar flow flight demonstration 'aircraft, the F-16XL, i.e. at a Mach number of 1.6 and a Reynolds number range of 1 to 3 million per foot. This will allow detailed studies of the attachment line and crossflow on the leading edge area of the highly swept wing. Also, use of suction as a means of control of transition due to crossflow and attachment line instabilities can be studied. Topics covered include: test operating conditions required; design requirements to efficiently make use of the existing infrastructure; development of an injector drive system using a small pilot facility; plenum chamber design; use of computational tools for tunnel and model design; and early operational results.

Description: A nonequilibrium, axisymmetric, Navier-Stokes flow solver with coupled radiation has been developed to use in the design of thermal protection systems for vehicles where radiation effects are important. The present method has been compared with an existing flow and radiation solver and with the Project Fire II experimental data. Very good agreement has been obtained over the entire Fire II trajectory with the experimentally determined values of the stagnation radiation intensity in the .2 to 6.2 eV range and with the total stagnation heating. The agreement was significantly better than previous numerical predictions. The effects of a number of flow models are examined to determine which combination of physical models produces the best agreement with the experimental data. These models include radiation coupling, multi-temperature thermal models, finite-rate chemistry, and a quasi-steady-state or Boltzmann assumption for the calculation of the excited electronic states. Finally, the computational efficiency of the present model is evaluated. The radiation properties model developed for this study is shown to offer significant computational savings compared to existing codes.

Description: NASA Ames Research Center is pursuing the development of SOFIA, the Stratospheric Observatory For Infrared Astronomy. SOFIA will consist of a 2.5 meter telescope mounted aft of the wing of a Boeing 747 aircraft. Since a large portion of the infrared spectrum is not visible at ground level due to absorption by water vapor in the atmosphere below 40,000 feet, it is highly desirable to make observations above this altitude. SOFIA will provide the opportunity for astronomers to conduct high-altitude research for extended periods of time. Current study is focused on wind tunnel testing for the open cavity. If not controlled, air would create resonance and damage the telescope. For this reason, SOFIA will design a boundary layer control device to achieve laminar flow over the cavity. This also provides a clearer flow for seeing, thus improving resolution on infrared sources. Other effects being tested in the wind tunnel are aerodynamic torque loads on the telescope, and flutter loads on the tail.

Description: Rotational temperatures have been measured in rarefied, nonequilibrium, heated freejet expansions of nitrogen using the electron beam fluorescence technique at the University of California at Berkeley Low Density Wind Tunnel facility. Spectroscopic measurements of the (0,0) band of the first negative system of nitrogen reveal the nonequilibrium behavior in the flowfield upstream of, and through the Mach disk, which forms as the freejet expands into a region of finite back pressure. Results compare well with previous freejet expansion data and computations regarding location of the Mach disk and terminal rotational temperature in the expansion. Measurements are also presented for shock thickness based on the rotational temperature changes in the flow. Thickening shock layers, departures of rotational temperature from equilibrium in the expansion region, and downstream rotational temperature recovery much below that of an isentropic normal shock provide indications of the rarefied, nonequilibrium flow behavior. The data are analyzed to infer constant values of the rotational-relaxation collision number from 2.2 to 6.5 for the various flow conditions. Collision numbers are also calculated in a consistent manner for data from other investigations for which is seen a qualitative increase with increasing temperature. Rotational-relaxation collision numbers are seen as not fully descriptive of the rarefied freejet flows. This may be due to the high degree of nonequilibrium in the flowfields, and/or to the use of a temperature-insensitive rotational-relaxation collision number model in the data analyses.

Description: An earlier proposed constitutive relation for normal stresses originated by random particle fluctuations is used to describe a joint effect of thermal and shear-induced fluctuations on concentrational distributions in suspension flow. Averaged products of components of the fluctuation velocity are evaluated on a basis of the rational mechanics approach combined with a simple kinematic consideration. The equation of momentum conservation of the dispersed phase of a suspension closed with this constitutive relation is applied to unidirectional shear flow in the gravity field and to rotational Couette flow. Coupling of the thermal and shear-induced fluctuations results in that the ability of shear flow to suspend particles has a minimum at a certain particle size, all other things being equal. The developed model provides also for a reasonable explanation of particle distributions observed in Couette flow. The approach based on the consideration of momentum balance for the dispersed phase is proved to lead to an effective equation of convective diffusion of the suspended particles. Coefficients of mutual diffusion due to both thermal and shear-induced fluctuations are drastically different from corresponding self-diffusivities as regards both their scaling and their concentrational dependence.

Description: The dynamic regime of gas injection through a circular plate orifice into an ideally wetting liquid is considered, when successively detached bubbles may be regarded as separate identities. In normal gravity and at relatively low gas flow rates, a growing bubble is modeled as a spherical segment touching the orifice perimeter during the whole time of its evolution. If the flow rate exceeds a certain threshold value, another stage of the detachment process takes place in which an almost spherical gas envelope is connected with the orifice by a nearly cylindrical stem that lengthens as the bubble rises above the plate. The bubble shape resembles then that of a mushroom and the upper envelope continues to grow until the gas supply through the stem is completely cut off. Such a stage is always present under conditions of sufficiently low gravity, irrespective of the flow rate. Two major reasons make for bubble detachment: the buoyancy force and the force due to the momentum inflow into the bubble with the injected gas. The former force dominates the process at normal gravity whereas the second one plays a key role under negligible gravity conditions. It is precisely this fundamental factor that conditions the drastic influence on bubble growth and detachment that changes in gravity are able to cause. The frequency of bubble formation is proportional to and the volume of detached bubbles is independent of the gas flow rate in sufficiently low gravity, while at normal and moderately reduced gravity conditions the first variable slightly decreases and the second one almost linearly increases as the flow rate grows. Effects of other parameters, such as the orifice radius, gas and liquid densities, and surface tension are discussed.

Description: Experimental results for a two-dimensional separated turbulent boundary layer behind a backward facing step for five different Reynolds numbers are reported. Results are presented in the form of tables, graphs and a floppy disk for an easy access of the data. Reynolds number based on the step height was varied by changing the reference velocity upstream of the step, U(sub o), and the step height, h. Hot-wire measurement techniques were used to measure three Reynolds stresses and four triple-velocity correlations. In addition, surface pressure and skin friction coefficients were measured. All hot-wire measurements were acquired in a measuring domain which excluded recirculating flow region due to the directional insensitivity of hot-wires. The downstream extent of the domain from the step was 51 h for the largest and I 14h for the smallest step height. This significant downstream length permitted extensive study of the flow recovery. Prediction of perturbed flows and their recovery is particularly attractive for popular turbulence models since variations of turbulence length and time scales and flow interactions in different regions are generally inadequately predicted. The data indicate that the flow in the free shear layer region behaves like the plane mixing layer up to about 2/3 of the mean reattachment length when the flow interaction with the wall commences the flow recovery to that of an ordinary turbulent boundary layer structure. These changes of the flow do not occur abruptly with the change of boundary conditions. A reattachment region represents a transitional region where the flow undergoes the most dramatic adjustments to the new boundary conditions. Large eddies, created in the upstream free-shear layer region, are being torn, recirculated, reentrained back into the main stream interacting with the incoming flow structure. It is foreseeable that it is quite difficult to describe the physics of this region in a rational and quantitative manner other than statistical. Downstream of the reattachment point the flow recovers at different rates near the wall, in the newly developing internal boundary layer, and in the outer part of the flow. It appears that Reynolds stresses do not fully recover up to the longest recovery length of 114 h.

Description: Tail buffet studies were conducted on a full-scale, production, F/A-18 fighter aircraft in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center at Moffett Field, California. Tail buffet data were acquired over an angle-of-attack range of +20 deg to +40 deg, a side-slip range of -16 deg to + 16 deg, and at wind speeds up to 100 knots. The maximum speed corresponds to a Reynolds number of l2.3 x l0(exp 6) based on mean aerodynamic chord and a Mach number of 0. 15. The port, vertical tail fin was instrumented with ninety-six surface-pressure transducers, arranged in six by eight arrays, on each side of the fin. ne aircraft was also equipped with a removable Leading-Edge Extension (LEX) fence whose purpose is to reduce tail-buffet loads. Current analysis methods for the unsteady aerodynamic pressures and loads are described. Only results for the zero side-slip condition are to be presented, both with and without the LEX fence. Results of the time-averaged, power-spectral analysis are presented for the tail fin bending moments which are derived from the integrated pressure field. Local wave velocities on the tail surfaces are calculated from pressure correlations. It was found that the LEX fence significantly reduces the magnitude of the root-mean-square pressures and bending moments. Scaling and repeatability issues are addressed by comparing the present full scale results for pressures at the 60%-span and 45%-chord location with previous full-scale F/A-18 tail-buffet test in the 80- by 120- Foot Wind Tunnel, and with several small-scale tests. The comparisons show that the tail buffet frequency scales very well with tail chord and free-stream velocity, and that there is good agreement with the previous full-scale test. Root-mean-square pressures and power spectra do not scale as well as the frequency results. Addition of a LEX fence caused tail-buffet loads to be reduced at all model scales.

Description: The occurrence of large-scale coherent structures in turbulent free shear flows (especially the planar mixing layer) has been recognized for some time. Indeed, the observation of such structures in mixing layers did much to promote interest in the study of coherent structures in turbulence. It has been widely assumed that the large-scale structures in these flows are responsible for the entrainment of free-stream fluid and the overall growth of the layer, while the small-scale structures provide mixing and dissipation. A model of scalar mixing based on these ideas was proposed for these flows. However, recent experimental and computational evidence suggests that the dominance of the large-scale structures in turbulent mixing layers is not universal. In addition, there is a substantial variation among experiments in several statistical measures of self-similar mixing layers, for example growth rate and velocity variances. To investigate the importance of large-scale structures, several free shear flows (mixing layers and wakes) have been simulated via direct numerical simulation. The simulations are designed to mimic experimental mixing layers in which the splitter plate boundary layers are turbulent. Different levels of two-dimensional forcing are included resulting in large-scale structures of differing strength and importance. These simulations are used to investigate the role of large-scale coherent structures in free shear layers and the effect of these structures on relevant turbulence statistics and scalar mixing. It is found that the statistics and structures in several experiments involving turbulent mixing layers are in better agreement with simulations that do not exhibit dominant large-scale structures than those in which the common mixing layer structures do dominate. It is also found that the level of forcing can have a profound effect on the qualitative and quantitative features of these shear layer, even when they are nominally self-similar.

Description: Computational fluid dynamics (CFD) is beginning to play a major role in the aircraft industry of the United States because of the realization that CFD can be a new and effective design tool and thus could provide a company with a competitive advantage. It is also playing a significant role in research institutions, both governmental and academic, as a tool for researching new fluid physics, as well as supplementing and complementing experimental testing. In this presentation, some of the progress made to date in CFD at NASA Ames will be reviewed. The presentation addresses the status of CFD in terms of methods, examples of CFD solutions, and computer technology. In addition, the role CFD will play in supporting the revolutionary goals set forth by the Aeronautical Policy Review Committee established by the Office of Science and Technology Policy is noted. The need for validated CFD tools is also briefly discussed.

Description: Large Dewars often use aluminum radiation shields and stainless steel vent lines. A simple, low cost method of making thermal contact between the shield and the line is to deform the shield around the line. A knowledge of the thermal conductance of such a joint is needed to thermally analyze the system. The thermal conductance of pressed metal contacts consisting of one aluminum and one stainless steel contact has been measured at 77 K, with applied forces from 8.9 N to 267 N. Both 5052 or 5083 aluminum were used as the upper contact. The lower contact was 304L stainless steel. The thermal conductance was found to be linear in temperature over the narrow temperature range of measurement. As the force was increased, the thermal conductance ranged from roughly 9 to 21 mW/K within a range of errors from 3% to 8%. Within the range of error no difference could be found between the using either of the aluminum alloys as the upper contact. Extrapolating the data to zero applied force does not result in zero thermal conductance. Possible causes of this anomalous effect are discussed.

Description: The recent resurgence of interest in utilizing laminar flow on aircraft surfaces for reduction in skin friction drag has generated a considerable amount of research in natural laminar flow (NLF) and hybrid laminar flow control (HLFC) on transonic aircraft wings. This research has focused primarily on airfoil design and understanding transition behavior with little concern for the surface imperfections and manufacturing variations inherent to most production aircraft. In order for laminar flow to find wide-spread use on production aircraft, techniques for constructing the wings must be found such that the large surface imperfections present in the leading edge region of current aircraft do not occur. Toward this end, a modification to existing leading edge construction techniques was devised such that the resulting surface did not contain large gaps and steps as are common on current production aircraft of this class. A lowspeed experiment was first conducted on a simulation of the surface that would result from this construction technique. Preston tube measurements of the boundary layer downstream of the simulated joint and flow visualization using sublimation chemicals validated the literature on the effects of steps on a laminar boundary layer. These results also indicated that the construction technique was indeed compatible with laminar flow. In order to fully validate the compatibility of this construction technique with laminar flow, thus proving that it is possible to build wings that are smooth enough to be used on business jets and light transports in a manner compatible with laminar flow, a flight experiment is being conducted. In this experiment Mach number and Reynolds number will be matched in a real flight environment. The experiment is being conducted using the NASA Dryden F-104 Flight Test Fixture (FTF). The FTF is a low aspect ratio ventral fin mounted beneath an F-104G research aircraft. A new nose shape was designed and constructed for this experiment. This nose shape provides an accelerating pressure gradient in the leading edge region. By flying the aircraft at appropriate Mach numbers and altitudes, this nose shape simulates the leading edge region of a laminar flow wing for a business jet or light transport. Manufactured into the nose shape is a spanwise slot located approximately four inches downstream of the leading edge. The slot, which is an inch wide and one-eighth of an inch deep allows the simulation of surface imperfections, such as gaps and steps at skin joints, which will occur on aircraft using this new construction technique. By placing strips of aluminum of various sizes and shapes in the slot, the effect on the boundary layer of different sizes and shapes of steps and gaps will be examined. It is planned to use five different configurations, differing primarily in the size and number of gaps. Downstream of the slot, the state of the boundary layer is determined using hot film gages and Stanton gages. Agreement between these two very different techniques of measuring boundary layer properties is considered important to being able to state with confidence the effects on the boundary layer of the simulated manufacturing imperfections. To date, the aircraft has not flown. First flights of the aircraft are on schedule to begin October 4, 1993. Low-speed, preliminary experiments at matching Reynolds numbers have been completed.

Description: Compressibility effects on turbulent transport of a passive scalar are studied within homogeneous turbulence using a kinematic decomposition of the velocity field into solenoidal and dilatational parts. It is found that the dilatational velocity does not produce a passive scalar flux, and that all of the passive scalar flux is due to the solenoidal velocity.

Description: Available redundancy among aircraft control surfaces allows for effective wing camber modifications. As shown in the past, this fact can be used to improve aircraft performance. To date, however, algorithm developments for in-flight camber optimization have been limited. This paper presents a perturbational approach for cruise optimization through in-flight camber adaptation. The method uses, as a performance index, an indirect measurement of the instantaneous net thrust. As such, the actual performance improvement comes from the integrated effects of airframe and engine. The algorithm, whose design and robustness properties are discussed, is demonstrated on the NASA Dryden B-720 flight simulator.

Description: This lecture attempts to illustrate the basic ideas of how the recent advances in nonlinear dynamical systems theory (dynamics) can provide new insights into the understanding of numerical algorithms used in solving nonlinear differential equations (DEs). Examples will be given of the use of dynamics to explain unusual phenomena that occur in numerics. The inadequacy of the use of linearized analysis for the understanding of long time behavior of nonlinear problems will be illustrated, and the role of dynamics in studying the nonlinear stability, accuracy, convergence property and efficiency of using time- dependent approaches to obtaining steady-state numerical solutions in computational fluid dynamics (CFD) will briefly be explained.

Description: A highly-instrumented UH-60A aircraft was tested at NASA-Ames Research Center from August 1993 to February 1994 obtaining an extensive data base for level flight, maneuvers, acoustics (both with respect to ground microphone arrays and inflight microphones), and flight dynamics. A majority of the data obtained are now in an electronic data base, however, only a small fraction of the data have been examined. The proposed paper will examine the issue of hovering steadiness in more detail. In particular, a single set of data obtained during ground acoustic testing may provide considerable insight as the wind speeds were measured at a hover height of 250 feet and the aircraft was positioned in 15 deg. steps in heading from 0 to 180 deg. Also, hover housekeeping data were obtained for many of the 31 flights and these will also allow a characterization of the unsteadiness. The variation in section lift will be examined in terms of the induced flow angle variation and this will be related to possible physical explanations.

Description: A collaborative team of researchers from fields of Computational Fluid Dynamics (CFD), fluid physics, computer architectures, and computer science and knowledge engineering have begun work on a prototype system that addresses several of industry's concerns in using NASA-developed CFD codes as part of the design cycle. A major problem exists in the application of CFD technologies within the aeronautics design cycle due primarily to misunderstandings in the ranges of applicability of the various solver codes or turbulence models. Features that arise during the CFD solution process need to be discriminated and recognized as actual flow features with physical support in the geometry and flow conditions of the problem being solved, or as numerical or non-physical errors arising from mis-application of solver code and its parameters, gridding strategies, or discretization. interpolations. The fundamental concept is to develop an intelligent computational system that can accept the engineer's definition of the problem and construct an optimal CFD solution. To do this requires capturing both the knowledge of how to apply the various CFD tools and how to adapt the application of those tools to flow structures as they evolve during the flow simulation. Embedded within this adaptive system approach is the additional desire to automatically identify and quantify the quality of resolution of the pertinent flow structures, be they genuine or error-induced, and then to adjust the solution strategy accordingly. This paper discusses the status of that prototyping effort.

Description: We present views and analysis of the execution of several PVM codes for Computational Fluid Dynamics on a network of Sparcstations, including (a) NAS Parallel benchmarks CG and MG (White, Alund and Sunderam 1993); (b) a multi-partitioning algorithm for NAS Parallel Benchmark SP (Wijngaart 1993); and (c) an overset grid flowsolver (Smith 1993). These views and analysis were obtained using our Automated Instrumentation and Monitoring System (AIMS) version 3.0, a toolkit for debugging the performance of PVM programs. We will describe the architecture, operation and application of AIMS. The AIMS toolkit contains (a) Xinstrument, which can automatically instrument various computational and communication constructs in message-passing parallel programs; (b) Monitor, a library of run-time trace-collection routines; (c) VK (Visual Kernel), an execution-animation tool with source-code clickback; and (d) Tally, a tool for statistical analysis of execution profiles. Currently, Xinstrument can handle C and Fortran77 programs using PVM 3.2.x; Monitor has been implemented and tested on Sun 4 systems running SunOS 4.1.2; and VK uses X11R5 and Motif 1.2. Data and views obtained using AIMS clearly illustrate several characteristic features of executing parallel programs on networked workstations: (a) the impact of long message latencies; (b) the impact of multiprogramming overheads and associated load imbalance; (c) cache and virtual-memory effects; and (4significant skews between workstation clocks. Interestingly, AIMS can compensate for constant skew (zero drift) by calibrating the skew between a parent and its spawned children. In addition, AIMS' skew-compensation algorithm can adjust timestamps in a way that eliminates physically impossible communications (e.g., messages going backwards in time). Our current efforts are directed toward creating new views to explain the observed performance of PVM programs. Some of the features planned for the near future include: (a) ConfigView, showing the physical topology of the virtual machine, inferred using specially formatted IP (Internet Protocol) packets; and (b) LoadView, synchronous animation of PVM-program execution and resource-utilization patterns.

Description: Compressibility plays a significant role in the development of separation on airfoils experiencing unsteady motion, even at moderately compressible free-stream flow velocities. This effect can result in completely changed stall characteristics compared to those observed at incompressible speed, and can dramatically affect techniques used to control separation. There has been a significant effort in recent years directed toward better understanding; of this process, and its impact on possible techniques for control of separation in this complex environment. A review of existing research in this area will be presented, with emphasis on the physical mechanisms that play such an important role in the development of separation on airfoils. The increasing impact of compressibility on the stall process will be discussed as a function of free-stream Mach number, and an analysis of the changing flow physics will be presented. Examples of the effect of compressibility on dynamic stall will be selected from both recent and historical efforts by members of the aerospace community, as well as from the ongoing research program of the present authors. This will include a presentation of a sample of high speed filming of compressible dynamic stall which has recently been created using real-time interferometry.

Description: This paper will describe the Airbreathing Hypersonic Research Program at NASA Ames Research Center. A main theme will be the "From Computation Through Flight" research effort. General research areas covered will include systems analysis, aerodynamics and aerothermodynamics, propulsion, materials, and flight research. Illustrative results from each discipline will be presented. The synergism between computational and experimental research will be demonstrated by examples. All examples given will have been published in the open literature.

Description: It is well known that slender bodies of revolution will develop an asymmetric, unsteady flow pattern in experimental tests, if the angle of incidence to the oncoming stream is above a critical value. It has been suggested that the origin of these asymmetric flows may stem from geometric imperfections of the model being tested, or from disturbances in the oncoming stream. In numerical simulations, it is possible to generate bodies of revolution which are perfectly symmetric about their longitudinal axis, and to impose uniform flow conditions which are free from disturbances. The current work presents numerical simulations of the flow about an ogive-cylinder configuration at 40 and 60 degree angle of incidence. These simulations. were performed using numerical algorithms which are also symmetric about the lateral plane of the cylinder body. The flowfields at 40 degree angle of attack were seen to remain symmetric to the round-off accuracy of the computer. At 60 degree angle of attack, a lateral force coefficient developed of O(1) which progressed to an alternate vortex shedding in time. The nature of this lateral force generation and vortex shedding was dependent on the choice of numerical algorithm. The origin of the asymmetries observed in the 60 degree angle of attack computations were traced to round-off errors in the implicit block-matrix inverter. A means of inverting the implicit operator matrices, which maintains the symmetry of the overall numerical algorithm was implemented.

Description: The goal of this work is to add insight about the flow within expansion tubes by using computational fluid dynamics. This is accomplished by comparing the results of axisymmetric numerical simulations with finite-rate chemistry to data from the HYPULSE expansion tube facility which was previously the NASA Langley expansion tube. The numerical simulations begin at the opening of the primary diaphragm and compute the flow throughout the whole facility and, thus, are able to follow and assess the effect of many of the flow features created during operation of the facility. One particular issue that will be investigated is the effect of boundary layer formation in the acceleration tube on the test gas volume and test gas conditions. Both laminar and turbulent boundary layers will be implemented. The effect of momentary shock reflection off the secondary diaphragm will also be investigated. There is concern that such a reflection will stagnate the test gas and create high levels of dissociated molecules. This is particularly important in propulsion experiments where a freestream composition different from flight conditions may influence ignition and burning data. Several different models of diaphragm rupture will be implemented in order to help understand the importance of this issue.

Description: A developed method has been applied to calculate accurately the viscous flow about airfoils normal to the free-stream flow. This method has special application to the analysis of tilt rotor aircraft in the evaluation of download. In particular, the flow about an XV-15 airfoil with and without deflected leading and trailing edge flaps at -90 degrees incidence is evaluated. The multi-element aspect of the method provides for the evaluation of slotted flap configurations which may lead to decreased drag. The method solves for turbulent flow at flight Reynolds numbers. The flow about the XV-15 airfoil with and without flap deflections has been calculated and compared with experimental data at a Reynolds number of one million. The comparison between the calculated and measured pressure distributions are very good, thereby, verifying the method. The aerodynamic evaluation of multielement airfoils will be conducted to determine airfoil/flap configurations for reduced airfoil drag. Comparisons between the calculated lift, drag and pitching moment on the airfoil and the airfoil surface pressure will also be presented.

Description: Steady and unsteady viscous, three-dimensional flowfields are calculated using a thin layer approximation of Navier-Stokes equations in conjunction with Chimera overset grids. The finite-difference numerical scheme uses structured grids and a pentadiagonal flow solver called "OVERFLOW". The configuration of Boeing 747-200 has been chosen as one of configurations to be used as a platform for the SOFIA (Stratospheric Observatory For Infrared Astronomy). Initially, the steady flowfield of the full aircraft is calculated for the clean configuration (without a cavity to house telescope). This solution is then used to start the unsteady flowfield of a configuration containing cavity housing the observation telescope and its peripheral units. Analysis of unsteady flowfield in the cavity and its influence on the tail empennage, as well as the noise due to turbulence and optical quality of the flow are the main focus of this study. For the configuration considered here, the telescope housing cavity is located slightly downstream of the portwing. The entire flow-field is carefully constructed using 45 overset grids and consists of nearly 4 million grid points. All the computations axe done at one freestream flow condition of M(sub infinity) = 0.85, alpha = 2.5deg, and a Reynolds of Re = 1.85x10deg

Description: Detailed experimental data have been obtained on several advanced TUFI systems during their exposure to high temperature high shear supersonic turbulent flow. Dimensional stability of these systems were determined at surface temperatures above 3000 F. Effect of step gap configuration on the thermal performance of the TUFI systems were also evaluated.

Description: An alternative theoretical model of joint filtration flow of immiscible incompressible fluids is presented. The model takes into account relaxation processes due to the interchange of the fluids between pores of difference sizes which is driven by capillary forces. The fluids occupy connected regions in a four-dimensional space formed by three coordinates and the pore length scale. When the fluid exchange between pores of given sizes is effected by way of successive flow through pores of all the intermediate sizes, the pressure within each region is governed by a hyperbolic equation, the role of time being played by the pore linear scale. Pressure jumps across hypersurfaces separating the regions equal corresponding values of the capillary pressure. A supplementary condition at any such hypersurface requires the speed of its displacement in the four-dimensional space to coincide with the normal velocity components of both the adjoining fluids. As a result, a principally new statement of multiphase filtration flow problems is gained with allowance for capillary relaxation in the porous space.

Description: The existing pseudocompressibility method for the system of incompressible Navier-Stokes equations is extended to heat transfer problems by including the energy equation. The solution method is based on the pseudo compressibility approach and uses an implicit-upwind differencing scheme together with the Gauss-Seidel line relaxation method. Current computations use one-equation Baldwin-Barth turbulence model which is derived from a simplified form of the standard k-epsilon model equations. Both forced and natural convection problems are examined. Numerical results from turbulent reattaching flow behind a backward-facing step will be compared against experimental measurements for the forced convection case. The validity of Boussinesq approximation to simplify the buoyancy force term will be investigated. The natural convective flow structure generated by heat transfer in a vertical rectangular cavity will be studied. The numerical results will be compared by experimental measurements by Morrison and Tran.

Description: The proposed paper presents flow visualization performed during experiments conducted on a full-scale F/A-18 aircraft in the 80- by 120-Foot Wind-Tunnel at NASA Ames Research Center. This investigation used both surface and off-surface flow visualization techniques to examine the flow field on the forebody, canopy, leading edge extensions (LEXs), and wings. The various techniques used to visualize the flow field were fluorescent tufts, flow cones treated with reflective material, smoke in combination with a laser light sheet, and a video imaging system. The flow visualization experiments were conducted over an angle of attack range from 20deg to 45deg and over a sideslip range from -10deg to 10deg. The results show regions of attached and separated flow on the forebody, canopy, and wings. Additionally, the vortical flow is clearly visible over the leading-edge extensions, canopy, and wings.

Description: It is stated that the aerodynamic forces on the vehicle being aerocaptured are controlled by "altering the angle of attack" and thereby controlling the lift coefficient. Furthermore, the resulting variation of drag coefficient with angle of attack was ignored. The purpose of this Comment is to point out that an aerodynamic control method that is much more effective than the pitch modulation has been studied and utilized during entries for many years. During aerocapture, it is desirable to have a large range of lift coefficients available, while keeping the vehicle's ballistic coefficients constant. This is accomplished by modulating the vehicle's bank angle, i.e., by rolling the vehicle about its velocity vector. By this method, the angle of attack can be held constant (at the trim angle, if desired), and the C(sub D) and the ballistic coefficient remain constant. Furthermore, the vertical component of the normal force vector (essentially the lift) can be varied over its entire range, from maximum positive to maximum negative values. Reaction controls, rather than aerodynamic ones, are usually utilized to change the bank angle of the vehicle, thus requiring the use of fuel. However, the fuel expenditure that is required to change the bank angle is far less than the amount that would have to be used to continuously hold the vehicle at pitch angles that differ significantly from its trim angle of attack. Also, it has been shown that bank angle modulation to vary the lift can enlarge the entry corridor by increasing the entry angle for the undershoot boundary, where both the heating rate and deceleration reach a maximum. Finally, the crew's deceleration tolerance can be increased somewhat when the bank angle is varied, as opposed to the pitch angle. For bank modulation, the deceleration force vector can be kept at a constant angle with respect to the occupants whose tolerance to g loads is highest when the force is applied in a direction normal to the upper torso. The advantages of bank angle variation to modulate the lift vector were recognized long ago, and this method of control was used successfully on the Apollo command module during lunar return' and, more recently, for the Space Shuttle Orbiter.

Description: We present views and analysis of the execution of several PVM (Parallel Virtual Machine) codes for Computational Fluid Dynamics on a networks of Sparcstations, including: (1) NAS Parallel Benchmarks CG and MG; (2) a multi-partitioning algorithm for NAS Parallel Benchmark SP; and (3) an overset grid flowsolver. These views and analysis were obtained using our Automated Instrumentation and Monitoring System (AIMS) version 3.0, a toolkit for debugging the performance of PVM programs. We will describe the architecture, operation and application of AIMS. The AIMS toolkit contains: (1) Xinstrument, which can automatically instrument various computational and communication constructs in message-passing parallel programs; (2) Monitor, a library of runtime trace-collection routines; (3) VK (Visual Kernel), an execution-animation tool with source-code clickback; and (4) Tally, a tool for statistical analysis of execution profiles. Currently, Xinstrument can handle C and Fortran 77 programs using PVM 3.2.x; Monitor has been implemented and tested on Sun 4 systems running SunOS 4.1.2; and VK uses XIIR5 and Motif 1.2. Data and views obtained using AIMS clearly illustrate several characteristic features of executing parallel programs on networked workstations: (1) the impact of long message latencies; (2) the impact of multiprogramming overheads and associated load imbalance; (3) cache and virtual-memory effects; and (4) significant skews between workstation clocks. Interestingly, AIMS can compensate for constant skew (zero drift) by calibrating the skew between a parent and its spawned children. In addition, AIMS' skew-compensation algorithm can adjust timestamps in a way that eliminates physically impossible communications (e.g., messages going backwards in time). Our current efforts are directed toward creating new views to explain the observed performance of PVM programs. Some of the features planned for the near future include: (1) ConfigView, showing the physical topology of the virtual machine, inferred using specially formatted IP (Internet Protocol) packets: and (2) LoadView, synchronous animation of PVM-program execution and resource-utilization patterns.

Description: FlowViz is a flow visualization application that uses Line Integral Convolution and the texture mapping capabilities of a graphics workstation to create an animation of flow over a curvilinear grid surface.

Description: This paper discusses the development of a general method for the determination of very low leak rates from limiting enclosures. There are many methods that can be used to detect and repair leaks from enclosures. Many methods have also been proposed that allow the estimation of actual leak rates, usually expressed as enclosure volume turnover. The proposed method combines measurements of the state variables (pressure, temperature, and volume) as well as the change in the concentration of a tracer gas to estimate the leak rate. The method was applied to the containment enclosure of the Engineering Development Unit of the CELSS Test Facility, currently undergoing testing at the NASA Ames Research Center.

Description: The case of isotropic compressible turbulence subjected to rapid isotropic compression is studied using inviscid rapid distortion theory and direct numerical simulation. An exact solution to the rapid distortion problem is given, and results are compared to those of direct numerical simulation. Implications for modelling turbulent flows are discussed.

Description: Combining multiple engineering workstations into a network-based heterogeneous parallel computer allows application of aerodynamic optimization with advance computational fluid dynamics codes, which is computationally expensive in mainframe supercomputer. This paper introduces a nonlinear quasi-Newton optimizer designed for this network-based heterogeneous parallel computer on a software called Parallel Virtual Machine. This paper will introduce the methodology behind coupling a Parabolized Navier-Stokes flow solver to the nonlinear optimizer. This parallel optimization package has been applied to reduce the wave drag of a body of revolution and a wing/body configuration with results of 5% to 6% drag reduction.

Description: In order to gain insights into the strong dependence of numerical solutions on initial data for finite time steps, a set of nonlinear test problems rich enough to capture the behavior of difference schemes were recently identified and the numerical basins of attraction for these problems were computed using commonly used time discretizations in CFD. Our study revealed a wealth of surprisingly nonlinear behavior of numerical schemes that were not observed before, in particular for the implicit time discretizations that are commonly used in CFD. The goal of this work is to apply these tools to study a practical model from non-equilibrium flowfield relaxation. This type of problem arises in chemically nonequilibrium hypersonic flows such as in a shock tube experiment or an expanding nozzle. Here we consider a reacting mixture of (N2, N) for an inviscid one-dimensional steady model. Preliminary numerical results indicate that, aside from the possibility of spurious numerical solutions being introduced by the time discretizations, limitations on the model for physical or accurate solutions may also play a part in the dynamics observed.

Description: Study of sonic and supersonic jet plumes are relevant to understanding such phenomenon as jet-noise, plume signatures, and rocket base-heating and radiation. Jet plumes are simple to simulate and yet, have complex flow structures such as Mach disks, triple points, shear-layers, barrel shocks, shock-shear-layer interaction, etc. Experimental and computational simulation of sonic and supersonic jet plumes have been performed for under- and over-expanded, axisymmetric plume conditions. The computational simulation compare very well with the experimental observations of schlieren pictures. Experimental data such as temperature measurements with hot-wire probes are yet to be measured and will be compared with computed values. Extensive analysis of the computational simulations presents a clear picture of how the complex flow structure develops and the conditions under which self-similar flow structures evolve. From the computations, the plume structure can be further classified into many sub-groups. In the proposed paper, detail results from the experimental and computational simulations for single, axisymmetric, under- and over-expanded, sonic and supersonic plumes will be compared and the fluid dynamic aspects of flow structures will be discussed.

Description: The present paper addresses some topical issues in modeling compressible turbulent shear flows. The work is based on direct numerical simulation of two supersonic fully developed channel flows between very cold isothermal walls. Detailed decomposition and analysis of terms appearing in the momentum and energy equations are presented. The simulation results are used to provide insights into differences between conventional time-and Favre-averaging of the mean-flow and turbulent quantities. Study of the turbulence energy budget for the two cases shows that the compressibility effects due to turbulent density and pressure fluctuations are insignificant. In particular, the dilatational dissipation and the mean product of the pressure and dilatation fluctuations are very small, contrary to the results of simulations for sheared homogeneous compressible turbulence and to recent proposals for models for general compressible turbulent flows. This provides a possible explanation of why the Van Driest density-weighted transformation is so successful in correlating compressible boundary layer data. Finally, it is found that the DNS data do not support the strong Reynolds analogy. A more general representation of the analogy is analysed and shown to match the DNS data very well.

Description: A shear-stress--sensitive liquid crystal coating (LCC) was used to visualize the surface shear stress distribution on the flat sidewall of the MSU quiet supersonic tunnel as a function of flow stagnation pressure. Under conditions of quiet operation, the LCC color-change response indicated the existence of a nonuniform surface shear stress distribution. This shear pattern was characterized by an elongated, down st ream-point ing triangular region of relatively low shear, with its apex on and its axis coincident with, the sidewall centerline. This low-shear zone was bounded symmetrically by two regions of relatively higher shear; these high-shear zones originated within the concave-curvature portion of the nozzle, in the corners between the flat sidewall and the contoured upper and lower nozzle surfaces. A 3-D Navier-Stokes code was used to compute the pressure and surface shear distributions on the sidewall. Flow-expansion-induced transverse pressure gradients on the nozzle sidewall generated symmetric inflows from the corners towards the sidewall centerline; these inflows caused a thickening of the sidewall boundary layer along the centerline, resulting in lower shear stresses consistent with the liquid crystal results. Peripherally nonuniform laminar boundary layer development, and the associated stability of such complex 3-D flows, must be considered in quiet-tunnel applications using rectangular nozzles. A color video will be shown.

Description: Three direct numerical simulations of time-evolving turbulent plane wakes with velocity deficit Reynolds numbers of about 2,000 have been simulated using a spectral numerical method with up to 600 x 260 x 160 modes. The initial conditions for the simulations are generated from direct numerical simulations of a turbulent boundary layer (momentum thickness Reynolds number of 670), and varying amounts of additional two- dimensional, forcing. In order to preserve the self-similar flow evolution, the forcing is implemented by multiplying all the two-dimensional modes in the initial condition by a constant factor. In the "natural" case no additional forcing is used; in the "forced" and "heavily forced" cases this factor is 5 and 20, respectively. The wake spreading rate Is increased by factors of 1.7 and 7.1 for the two forced cases. The Reynolds stresses are also increased by a similar or even larger factor. These results indicate that the plane wake is much more sensitive to initial forcing than the plane mixing layer. As in the plane mixing layer, two-dimensional forcing promotes more organized large-scale vortical flow structures and these structures axe sometimes separated by "braid regions" containing streamwise "rib" vortices, unlike in the unforced wake.

Description: Large-eddy simulation of the incompressible Navier-Stokes equations has been used to examine the long-time development of initially isotropic turbulence subjected to solid-body rotation. The simulations were carried out using a pseudo-spectral method with 128 x 128 x 512 collocation points in a computational domain that is four times larger along the rotation axis than in the other directions; subgrid-scale motions were parameterized using a spectral eddy viscosity model modified for system rotation. Simulation results show that the correlation length along the rotation am's of velocities orthogonal to the rotation vector exhibits rapid growth while the integral length-scale of velocities aligned with the rotation axis is relatively unaffected by rotation. Examination of the energy spectrum of two-dimensional, two-component motions indicates the presence of an inverse cascade of energy. System rotation also causes an alignment of vorticity along the rotation axis with relatively stronger cyclonic vorticity than anticyclonic. The onset of anisotropic effects are well characterized by Rossby numbers defined in terms of both macroscopic and microscopic quantities.

Description: This paper will review the advances made recently in the Navier-Stokes CFD methods to simulate aerodynamics and aeroacoustics of helicopter rotors and rotor-body flows. Although a complete flowfield simulation of full helicopter is currently not feasible with these methods, impressive gains have been made in analyzing individual components of this complex problem in a very detailed manner. The use of the state-of-the-art numerical algorithms in solution methods, in conjunction with powerful supercomputers, like the Cray-2, have enabled noticeable progress to be made in modeling viscous-inviscid interactions, blade-vortex interactions, tip-vortex: simulation and wake effects, as well as high speed impulsive noise in hover and forward flight for isolated rotor blades. This paper will critically evaluate the presently available Euler and Navier-Stokes methods, both finite-difference and finite volume methods using structured and unstructured grids for helicopter applications for accuracy, suitability, and computational efficiency. The review will also include the recent progress made using overset grids to model rotor-body flows. All the material for this review will be drawn from the published material shown below.

Description: In recent years significant advances have been made for parallel computers in both hardware and software. Now parallel computers have become viable tools in computational mechanics. Many application codes developed on conventional computers have been modified to benefit from parallel computers. Significant speedups in some areas have been achieved by parallel computations. For single-discipline use of both fluid dynamics and structural dynamics, computations have been made on wing-body configurations using parallel computers. However, only a limited amount of work has been completed in combining these two disciplines for multidisciplinary applications. The prime reason is the increased level of complication associated with a multidisciplinary approach. In this work, procedures to compute aeroelasticity on parallel computers using direct coupling of fluid and structural equations will be investigated for wing-body configurations. The parallel computer selected for computations is an Intel iPSC/860 computer which is a distributed-memory, multiple-instruction, multiple data (MIMD) computer with 128 processors. In this study, the computational efficiency issues of parallel integration of both fluid and structural equations will be investigated in detail. The fluid and structural domains will be modeled using finite-difference and finite-element approaches, respectively. Results from the parallel computer will be compared with those from the conventional computers using a single processor. This study will provide an efficient computational tool for the aeroelastic analysis of wing-body structures on MIMD type parallel computers.

Description: 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.

Description: Experiments were performed to study the evolution of the heat transfer structure in a separated free shear layer region of an incompressible separated turbulent boundary layer flow behind a backward-facing step. While there is an abundance of velocity field measurements of separated flows, heat transfer measurements are rather scarce, thus limiting assessment of the heat transfer physics and its accurate modeling. The purpose of the paper is twofold: to improve an understanding of effects of flow separation on heat transfer characteristics, and to provide data for turbulence modeling and computation. The boundary layer upstream of the step was turbulent and fully developed. A constant temperature surface boundary condition was imposed upstream and downstream of the step for the heat transfer study. An internal mixing-layer like flow forms and grows from the step lip within the original boundary layer. The turbulent structure of the flow evolving downstream, however, does not switch immediately to that of a mixing layer over the entire shear layer thickness. Measurements of mean and fluctuating velocity and temperature fields indicate that the internal layer spreads gradually in the transverse direction while the outer part of the original boundary layer is effectively unperturbed. The results in this paper have not been previously reported.

Description: Direct simulation Monte Carlo (DSMC) calculations of rarefied flows about entry bodies typically employ a fixed surface temperature or a radiative-equilibrium energy balance to compute that temperature. Such boundary conditions neglect any effects of heat capacitance and heat conduction in the spacecraft heat shield and, therefore, provide an upper bound for the surface temperature. Such calculations also neglect pyrolysis from the heat shield which can be significant for a high-energy incident flow at very low densities. Accurate prediction of both heating and aerodynamic forces requires including pyrolysis and surface heat transfer in the models for gas-surface interaction employed in DSMC methods. Although these physical models have long appeared in various continuum flow calculation codes, they have only recently appeared in DSMC codes which are required to simulate rarefied flows during entry at high altitudes. In the current implementation, routines from the widely distributed Charring Material Thermal Response and Ablation (CMA) program are coupled into a DSMC code to calculate the one-dimensional heat transfer into the carbon phenolic heat shield at each point on a vehicle surface. Temperature-dependent material properties, surface re-radiation, and in-depth pyrolysis were included in the calculation, but surface ablation was neglected. Sample calculations for entry of the Galileo probe into the atmosphere of Jupiter demonstrate that including pyrolysis in the model leads to significant differences in predicted aerodynamics. Granted, the drag coefficient does not depend strongly on the surface temperature which can itself be significantly below the radiative equilibrium value during entry. However, the surface mass flux due to pyrolysis of the material is significant once the probe drops to altitudes characterized by transition flow. This leads to a noticeable increase in drag and a decrease in heating compared to a body without pyrolysis.

Description: A fluid channeling system includes a fluid ejector, a heat exchanger, and a fluid pump disposed in series flow communication The ejector includes a primary inlet for receiving a primary fluid, and a secondary inlet for receiving a secondary fluid which is mixed with the primary fluid and discharged therefrom as ejector discharge. Heat is removed from the ejector discharge in the heat exchanger, and the heat exchanger discharge is compressed in the fluid pump and channeled to the ejector secondary inlet as the secondary fluid In an exemplary embodiment, the temperature of the primary fluid is greater than the maximum operating temperature of a fluid motor powering the fluid pump using a portion of the ejector discharge, with the secondary fluid being mixed with the primary fluid so that the ejector discharge temperature is equal to about the maximum operating temperature of the fluid motor.

Description: This paper describes a parallel implementation of the direct simulation Monte Carlo (DSMC) method. Runtime library support is used for scheduling and execution of communication between nodes, and domain decomposition is performed dynamically to maintain a good load balance. Performance tests are conducted using the code to evaluate various remapping and remapping-interval policies, and it is shown that a one-dimensional chain-partitioning method works best for the problems considered. The parallel code is then used to simulate the Mach 20 nitrogen flow over a finite-thickness flat plate. It is shown that the parallel algorithm produces results which compare well with experimental data. Moreover, it yields significantly faster execution times than the scalar code, as well as very good load-balance characteristics.

Description: The effect of far-field boundary conditions on the evolution of a finite-amplitude two-dimensional wave in the Blasius boundary layer is assessed. With the use of the parabolized stability equations (PSE) theory for the numerical computations, either asymptotic, Dirichlet, Neumann or mixed boundary conditions are imposed at various distances from the wall. The results indicate that asymptotic and mixed boundary conditions yield the most accurate mean-flow distortion and unsteady instability modes in comparison with the results obtained with either Dirichlet or Neumann conditions.

Description: The Interface Configuration Experiment (ICE) was carried out on USML-1 to investigate liquid-gas interfaces in certain rotationally-symmetric containers having prescribed, mathematically derived shapes. These containers have the property that they admit an entire continuum of distinct equilibrium rotationally-symmetric interfaces for a given liquid volume and contact angle. Furthermore, it can be shown that none of these interfaces can be stable. It was found, after the containers were filled in orbit, that an initial equilibrium interface from the symmetric continuum re-oriented, when perturbed, to a stable interface that was not rotationally symmetric, in accordance with the mathematical theory.

Description: A numerical study was performed to investigate the shock-wave/boundary-layer interactions on a flat plate with bleed through one or more circular holes that vent into a plenum. The bleed-hole patterns considered for the study include in-line multiple holes and staggered multiple-row holes that are configured to simulate the patterns used in inlet bleed systems of high performance aircraft. The focus of the study was to examine how the bleed through multiple holes affect bleed rate and the pressure and Mach number distributions. Since the bleed performance was found sensitive to the change in bleed conditions, a computational procedure was developed to give a good turnaround computational time for parametric studies involving changes in bleed hole geometry and the structure of shock-wave/boundary-layer flowfield. The procedure includes the grid-generation methodology and the flow simulation with solutions from the Navier-Stokes equations. The computational techniques permit analysis of complex bleed systems and make possible the investigation of a broader range of design variables associated with inlet bleed operation.

Description: A gas turbine engine hot section combustor liner is provided a non-film cooled portion of a heat transfer wall having a hot surface and a plurality of longitudinally extending micro-grooves disposed in the portion of the wall along the hot surface in a direction parallel to the direction of the hot gas flow. The depth of the micro-grooves is very small and on the order of magnitude of a predetermined laminar sublayer of a turbulent boundary layer. The micro-grooves are sized so as to inhibit heat transfer from the hot gas flow to the hot surface of the wall while reducing NOx emissions of the combustor relative to an otherwise similar combustor having a liner wall portion including film cooling apertures. In one embodiment the micro-grooves are about 0.001 inches deep and have a preferred depth range of from about 0.001 inches to 0.005 inches and which are square, rectangular, or triangular in cross-section and the micro-grooves are spaced about one width apart.

Description: The weakly nonlinear evolution of an inviscid marginally unstable wave growing on a boundary layer supporting a streamwise vortex structure is investigated. The nonlinear growth of the wave is found to be controlled by the diffusion layer located at the edge of the critical layer associated with the wave. The evolution equation is found to depend on the upstream history of the wave and the solution of the equation suggests that the wave either restructures the mean state so as to make it stable or develops a singularity at a finite distance downstream of the point of neutral stability.

Description: One or the key objectives of the Applied Research Branch in the Numerical Aerodynamic Simulation (NAS) Systems Division at NASA Allies Research Center is the accelerated introduction of highly parallel machines into a full operational environment. In this report we discuss the performance results obtained from the implementation of some computational fluid dynamics (CFD) applications on the Connection Machine CM-2 and the Intel iPSC/860. We summarize some of the experiences made so far with the parallel testbed machines at the NAS Applied Research Branch. Then we discuss the long term computational requirements for accomplishing some of the grand challenge problems in computational aerosciences. We argue that only massively parallel machines will be able to meet these grand challenge requirements, and we outline the computer science and algorithm research challenges ahead.

Description: Combining multiple engineering workstations into a network-based heterogeneous parallel computer allows application of aerodynamic optimization with advance computational fluid dynamics codes, which is computationally expensive in mainframe supercomputer. This paper introduces a nonlinear quasi-Newton optimizer designed for this network-based heterogeneous parallel computer on a software called Parallel Virtual Machine. This paper will introduce the methodology behind coupling a Parabolized Navier-Stokes flow solver to the nonlinear optimizer. This parallel optimization package has been applied to reduce the wave drag of a body of revolution and a wing/body configuration with results of 5% to 6% drag reduction.

Description: The performance of a workstation cluster used for the solution of the Reynolds-averaged Navier-Stokes equations is compared with a conventional vector supercomputer architecture. The application simulation of the steady flowfield about a transonic transport was computed using an implicit diagonal scheme in an overset mesh framework. Static load balancing was used, while coarse grain decomposition was achieved by solution of a grid zone per processor. Price/performance ratios are estimated for several scenarios in which such clusters may be utilized.

Description: A subsonic wind tunnel investigation of pneumatic vortex flow control on a chined forebody using slots was accomplished at a dynamic pressure of 50 psf resulting in a R(n)/ft of 1.3 x 10(exp 6). Data were acquired from angles of attack ranging from -4deg to +34deg at side slips of +0.4deg and +10.4deg. The test article used in this study was the 10% scale Fighter Lift and Control (FLAC) advanced diamond winged, vee-tailed fighter configuration. Three different slot blowing concepts were evaluated; outward, downward, and tangential with ail blowing accomplished asymmetrically. The results of three different mass flows (0.067, 0.13, and 0.26 lbm/s; C(sub mu)'s of less than or equal to 0.006, 0.011. and 0.022 respectively) were analyzed and reported. Test data are presented on the effects of mass flows, slot lengths and positions and blowing concepts on yawing moment and side force generation. Results from this study indicate that the outward and downward blowing slots developed yawing moment and side force increments in the direction opposite of the blowing side while the tangential blowing slots generated yawing moment and side force increments in the direction towards the blowing side. The outward and downward blowing slots typically produced positive pitching moment increments while the tangential blowing slots typically generated negative pitching moment increments. The slot blowing nearest the forebody apex was most effective at generating the largest increments and as the slot was moved aft or increased in length, its effectiveness at generating forces and moments diminished.

Description: The aim of the present investigation is to characterize the motion of dendrite fragments falling under the influence of gravity in a uniform liquid medium at low Reynolds number. In an earlier study, Zakhem, Weidman and de Groh (1992) reported on the settling speed of model equiaxed dendrite grains released along their axis of symmetry. In this follow-up study uniaxial model dendrite grains were released off-axis to observe and document their motion at different orientations. It was hypothesized that the dendrite models might rotate when released off-axis in which case an attempt would be made to document the ensuing unsteady motion. This latter event turned out to be in fact true: at the small but finite Reynolds numbers that existed, each uniaxial dendrite slowly rotated towards its equilibrium orientation while failing under the influence of gravity. In addition to completing the original goal, we have made use of a beads-on-a shell Stokes flow code to numerically determine the drag coefficient for capsules, i.e.. uniaxial dendrites without arms. The drag on horizontally and vertically falling capsules are reported and compared with measurements.

Description: Supersonic jet plumes were studied using a two-equation turbulence model employing corrections for compressible dissipation and pressure-dilatation. A space-marching procedure based on an upwind numerical scheme was used to solve the governing equations and turbulence transport equations. The computed results indicate that two-equation models employing corrections for compressible dissipation and pressure-dilatation yield improved agreement with the experimental data. In addition, the numerical study demonstrates that the computed results are sensitive to the effect of grid refinement and insensitive to the type of velocity profiles used at the inflow boundary for the cases considered in the present study.

Description: Results are obtained for cylindrical leading edges of proposed transatmospheric vehicles by employing a two-dimensional viscous shock-layer code for nonequilibrium gas flows. The accuracy and efficiency of the planar code is verified through detailed comparisons with other predictions. This study includes results for 6-deg half-angle bodies with nose radii ranging from 0.01 to 2.0 ft for both cylindrically blunted wedges and spherically blunted cones (included for comparison). Some results are presented as a ratio of the noncatalytic to the corresponding fully catalytic heating value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flows. Generally, this ratio and the individual heating rates are smaller for cylindrically blunted wedges with small nose radii as compared to the spherically blunted cones (for the same nose radius). Therefore, a larger potential exists for heating reduction in cylindrically blunted as compared with the spherically blunted surfaces. However, the results presented at higher altitudes (where the slip effects become important) show that the spherically, blunted nose gives lower stagnation-point heating due to stronger merged shock-layer effects as compared with a cylindrically blunted nose.

Description: The three-dimensional Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate nonsteady vortical flow about a 65 degree sweep delta wing at 30 degrees angle of attack. Two large-amplitude, high-rate, forced-roll motions and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are compared with experimental forces, moments, and roll-angle time histories. The overall agreement is good. Vortex breakdown is present in each case, which causes significant time lags in the vortex breakdown motions relative to the body motions. This behavior strongly influences the dynamic forces and moments.

Description: A study is described that evaluates the accuracy of vortex-lattice methods when they are used to compute the loads induced on aircraft as they encounter lift-generated wakes. The evaluation is accomplished by use of measurements made in the 80- by 120-foot wind tunnel of the lift, rolling-moment, and downwash in the wake of three configurations of a model of a subsonic transport aircraft. The downwash measurements are used as input for a vortex-lattice code in order to compute the lift and rolling moment induced on wings that have a span of 0.186, 0.510, or 1.022 times the span of the wake-generating model. Comparison of the computed results with the measured lift and rolling moment distributions are used to determine the accuracy of the vortex-lattice code. It was found that the vortex-lattice method is very reliable as long as the span of the encountering of following wing is less than about 0.2 of the generator span. As the span of the following wing increases above 0.2, the vortex-lattice method continues to correctly predict the trends and nature of the induced loads, but it overpredicts the magnitude of the loads by increasing amounts. The increase in deviation of the computed from the measured loads with size of the following wing is attributed to the increase in distortion of the structure of the vortex wake as it approaches and passes the larger following wings.

Description: The viscous flow field near the surface of a hovering rotor blade was studied for blade twist distributions typical of a till rotor blade and a conventional helicopter rotor blade. Three blade geometries were studied, including a tilt rotor blade twist distribution (baseline), conventional helicopter rotor blade twist distribution, and the baseline twist distribution with 2 deg of precone. The results give insight into the delayed stall phenomenon often observed for highly twisted rotors. Calculations were performed for a high thrust condition near stall using the thin-layer Navier-Stokes CFD code TURNS. Effects of built-in twist on section force coefficients, skin friction, velocities, surface pressures, and boundary layer shape factor are discussed. Although the rotor thrust coefficient was nominally the same for the cases using the two twist distributions, large differences were found in the section in-plane and normal force coefficients. These preliminary results imply that the blade outboard region, rather than the inboard region, provides the majority of the performance advantage of the baseline case over the low twist case. Skin friction, velocities near the blade, and surface pressures for the two twist distributions reveal significant differences in the blade outboard region.

Description: A hybrid method for computing compressible viscous flows is presented. This method divides the computational domain into two zones. In the outer zone, the unsteady full-potential equation (FPE) is solved. In the inner zone, the Navier-Stokes equations are solved. The two zones are tightly coupled so that steady and unsteady flows may be efficiently solved. The resulting CPU times are less than 50 percent of the required for a full-blown Navier-Stokes analysis. Sample applications of the method to an unswept iced wing at 4 deg and 8 deg angle of attack are presented. Surface pressures are in good agreement with the measurements obtained by Bragg et al. at the University of Illinois.

Description: Strong interactions between flow about an aircraft wing and the wing structure can result in aeroelastic phenomena which significantly impact aircraft performance. Time-accurate methods for solving the unsteady Navier-Stokes equations have matured to the point where reliable results can be obtained with reasonable computational costs for complex non-linear flows with shock waves, vortices and separations. The ability to combine such a flow solver with a general finite element structural model is key to an aeroelastic analysis in these flows. Earlier work involved time-accurate integration of modal structural models based on plate elements. A finite element model was developed to handle three-dimensional wing boxes, and incorporated into the flow solver without the need for modal analysis. Static condensation is performed on the structural model to reduce the structural degrees of freedom for the aeroelastic analysis. Direct incorporation of the finite element wing-box structural model with the flow solver requires finding adequate methods for transferring aerodynamic pressures to the structural grid and returning deflections to the aerodynamic grid. Several schemes were explored for handling the grid-to-grid transfer of information. The complex, built-up nature of the wing-box complicated this transfer. Aeroelastic calculations for a sample wing in transonic flow comparing various simple transfer schemes are presented and discussed.

Description: A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.

Description: 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 under-blowing 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 sub-scale 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 flow field about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.

Description: The phenomena of rotational relaxation of nitrogen has been examined by numerous investigators over many years. One of the experiments which has been performed examines nonequilibrium flow in low-density free jet expansions. Data have been taken in such flows using a variety of techniques, including time-of-flight methods and electron beam fluorescence spectroscopy. The direct flow properties measured in these different investigations, such as density and translational, rotational and vibrational temperatures generally show reasonable agreement. However, this kind of correlation from experiment to experiment tends to be lost when these data are analyzed to obtain rotational relaxation time or collision number. The goal of such data analyses is to generate a succinct model for rotational relaxation in nitrogen which is essential for the computation of nonequilibrium rarefied flows. The objective of the present work is to process a large body of experimental data in a consistent manner to yield relaxation model parameters of the greatest utility for flow computations.

Description: This study focuses on the suppression of instability growth using an automated active-control technique. The evolution of 2D disturbances that are spatially growing in a flat-plate boundary layer are computed with a spatial DNS code. A controller receives wall sensor information (pressure or shear) as input and provides a signal that controls an actuator response as output. The control law assumes that wave cancellation is valid. 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: An apparatus comprising a rotatable mass of structured packing for mass or heat transfer between two contacting fluids of different densities wherein the packing mass is made up of corrugated sheets of involute shape relative to the axis of the packing mass and form a logarithmic spiral curved counter to the direction of rotation.

Description: An experimental survey of supersonic wing tip vortices has been conducted at Mach 2.5 using small performed 2.25 chords down-stream of a semi-span rectangular wing at angle of attack of 5 and 10 degrees. The main objective of the experiments was to determine the Mach number, flow angularity and total pressure distribution in the core region of supersonic wing tip vortices. A secondary aim was to demonstrate the feasibility of using cone probes calibrated with a numerical flow solver to measure flow characteristics at supersonic speeds. Results showed that the numerically generated calibration curves can be used for 4-hole cone probes, but were not sufficiently accurate for conventional 5-hole probes due to nose bluntness effects. Combination of 4-hole cone probe measurements with independent pitot pressure measurements indicated a significant Mach number and total pressure deficit in the core regions of supersonic wing tip vortices, combined with an asymmetric 'Burger like' swirl distribution.

Description: Study of turbulent flows in rotating reference frames has proven to be one of the more challenging areas of turbulence research. The large number of theoretical, experimental, and computational studies performed over the years have demonstrated that the effect of solid-body rotation on turbulent flows is subtle and remains exceedingly difficult to predict. Because of the complexities associated with non-homogeneous turbulence, it is worthwhile to examine the effect of steady system rotation on the evolution of an initially isotropic turbulent flow. The assumption of statistical homogeneity considerably simplifies analysis and computation; calculation of homogeneous turbulence is further motivated since it possesses the essential physics found in more complex rotating flows. The principal objectives of the present study have therefore been to increase our fundamental understanding of turbulent flows in rotating reference frames through an examination of the asymptotic state of homogeneous rotating turbulence; particularly as to the existence of an asymptotic state which is self similar. Knowledge of an asymptotic similarity state permits prediction of the ultimate statistical evolution of the flow without requiring detailed knowledge of the complex, and not well understood, non-linear transfer processes. Aside from examination of possible similarity states in rotating turbulence, of further interest in this study has been an examination of the degree to which solid-body rotation induces a two-dimensional state in an initially isotropic flow.

Description: Digital flight records from reported clear-air turbulence incidents are used to determine winds and turbulence, to determine maneuver g loads, and to analyze control problems. Many cases of severe turbulence are found downwind of mountains and thunderstorms where sharp, sudden jolts are associated with vortices in atmospheric waves. Other cases of severe turbulence are round in strong updrafts above thunderstorm buildups that may be undetected by onboard weather radar. An important finding is that there are large maneuvering loads in over half of the reported clear-air turbulence incidents. Maneuvering loads are determined through an analysis of the short-term variations in elevator deflection and aircraft pitch angle. For altitude control in mountain waves the results indicate that small pitch angle changes with proper timing are sufficient to counter variations in vertical wind. For airspeed control in strong mountain waves, however, there is neither the available thrust nor the quickness in engine response necessary to counter the large variations in winds.

Description: The effect of forebody tangential slot blowing on the flowfield about an F/A-18 aircraft is investigated numerically using solutions of the Navier-Stokes equations. Computed solutions are obtained for an aircraft geometry which includes the fuselage, a wing with deflected leading-edge flap, empennage, and a faired-over engine inlet. The computational slot geometry corresponds to that used in full-scale wind-tunnel tests. Solutions are computed using flight test conditions and jet mass flow ratios equivalent to wind-tunnel test conditions. The effect of slot location is analyzed by computing two nontime-accurate solutions with a 16-in. slot located 3 in. and 11 in. aft of the nose of the aircraft. These computations resolve the trends observed in the full-scale wind-tunnel test data. The flow aft of the leading-edge extension vortex burst is unsteady. A time-accurate solution is obtained to investigate the flow characteristics aft of the vortex burst, including the effect of blowing on tail buffet.