ALBERT

Description: This chapter is concerned with three-dimensional imaging of fluid flows. Although relatively young, this field of research has already yielded an enormous range of techniques. These vary widely in cost and complexity, with the cheapest light sheet systems being within the budgets of most laboratories, and the most expensive Magnetic Resonance Imaging systems available to a select few. Taking the view that the most likely systems to be developed are those using light sheets, the authors will relate their knowledge and experience of such systems. Other systems will be described briefly and references provided. Flows are inherently three-dimensional in structure; even those generated around nominally 2-D surface geometry. It is becoming increasingly apparent to scientists and engineers that the three-dimensionalities, both large and small scale, are important in terms of overall flow structure and species, momentum, and energy transport. Furthermore, we are accustomed to seeing the world in three dimensions, so it is natural that we should wish to view, measure and interpret flows in three-dimensions. Unfortunately, 3-D images do not lend themselves to convenient presentation on the printed page, and this task is one of the challenges facing us.

Description: The objectives of this effort were to determine (if possible) the best method: 1) for forcing the boundary layer to transition, 2) for assessing trip effectiveness, 3) for quantifying trip drag, 4) for testing at Reynolds numbers per foot from 5 million to maximum available rather than I to 5 million, and 5) for boundary layer state determination.

Description: NASA Langley Research Center (LaRC) has field tested a 2.0 gm, 100 Hertz, pulsed coherent lidar to detect and characterize wake vortices and to measure atmospheric winds and turbulence. The quantification of aircraft wake-vortex hazards is being addressed by the Wake Vortex Lidar (WVL) Project as part of Aircraft Vortex Spacing System (AVOSS), which is under the Reduced Spacing Operations Element of the Terminal Area Productivity (TAP) Program. These hazards currently set the minimum, fixed separation distance between two aircraft and affect the number of takeoff and landing operations on a single runway under Instrument Meteorological Conditions (IMC). The AVOSS concept seeks to safely reduce aircraft separation distances, when weather conditions permit, to increase the operational capacity of major airports. The current NASA wake-vortex research efforts focus on developing and validating wake vortex encounter models, wake decay and advection models, and wake sensing technologies. These technologies will be incorporated into an automated AVOSS that can properly select safe separation distances for different weather conditions, based on the aircraft pair and predicted/measured vortex behavior. The sensor subsystem efforts focus on developing and validating wake sensing technologies. The lidar system has been field-tested to provide real-time wake vortex trajectory and strength data to AVOSS for wake prediction verification. Wake vortices, atmospheric winds, and turbulence products have been generated from processing the lidar data collected during deployments to Norfolk (ORF), John F. Kennedy (JFK), and Dallas/Fort Worth (DFW) International Airports.

Description: Subgrid analysis of a transitional temporal mixing layer with evaporating droplets has been performed using three sets of results from a Direct Numerical Simulation (DNS) database, with Reynolds numbers (based on initial vorticity thickness) as large as 600 and with droplet mass loadings as large as 0.5. In the DNS, the gas phase is computed using a Eulerian formulation, with Lagrangian droplet tracking. The Large Eddy Simulation (LES) equations corresponding to the DNS are first derived, and key assumptions in deriving them are first confirmed by computing the terms using the DNS database. Since LES of this flow requires the computation of unfiltered gas-phase variables at droplet locations from filtered gas-phase variables at the grid points, it is proposed to model these by assuming the gas-phase variables to be the sum of the filtered variables and a correction based on the filtered standard deviation; this correction is then computed from the Subgrid Scale (SGS) standard deviation. This model predicts the unfiltered variables at droplet locations considerably better than simply interpolating the filtered variables. Three methods are investigated for modeling the SGS standard deviation: the Smagorinsky approach, the Gradient model and the Scale-Similarity formulation. When the proportionality constant inherent in the SGS models is properly calculated, the Gradient and Scale-Similarity methods give results in excellent agreement with the DNS.

Description: A three-dimensional Navier Stokes code has been used to compute the heat transfer coefficient on a film-cooled, rotating turbine blade. The blade chosen is the ACE rotor with five rows containing 93 film cooling holes covering the entire span. This is the only film-cooled rotating blade over which experimental data is available for comparison. Over 2.278 million grid points are used to compute the flow over the blade including the tip clearance region. using Wilcox's k-omega model, Coakley's q-omega model, and the zero-equation Baldwin-Lomax (B-L) model. A reasonably good comparison with the experimental data is obtained on the suction surface for all the turbulence models. At the leading edge, the B-L model yields a better comparison than tile two-equation models. On the pressure surface however the comparison between the experimental data and the prediction from the k-omega model is much better than from the other two models. Overall, the k-omega model provides the best comparison with the experimental data. However, the two-equation models require at least 40% more computational resources than the B-L model.

Description: We report measurements of the depression of the superfluid transition temperature by a heat current (1 less than or = Q less than or = 100 microW/sq cm) along the lambda-line (SVP less than or = P less than or = 21.6 bar). At P = 21.6 bar, measurements were also performed in a reduced gravity (0.2g). Experimental results show that the pressure dependence of the depression and the gravity effect on the measurements are small, in qualitative agreement with theoretical predictions. Keywords: superfluid helium; Lambda transition; heat current

Description: The differences between subcritical liquid drop and supercritical fluid drop behavior are discussed. Under subcritical, evaporative high emission rate conditions, a film layer is present in the inner part of the drop surface which contributes to the unique determination of the boundary conditions; it is this film layer which contributes to the solution's convective-diffusive character. In contrast, under supercritical condition as the boundary conditions contain a degree of arbitrariness due to the absence of a surface, and the solution has then a purely diffusive character. Results from simulations of a free fluid drop under no-gravity conditions are compared to microgravity experimental data from suspended, large drop experiments at high, low and intermediary temperatures and in a range of pressures encompassing the sub-and supercritical regime. Despite the difference between the conditions of the simulations and experiments (suspension vs. free floating), the time rate of variation of the drop diameter square is remarkably well predicted in the linear curve regime. The drop diameter is determined in the simulations from the location of the maximum density gradient, and agrees well with the data. It is also shown that the classical calculation of the Lewis number gives qualitatively erroneous results at supercritical conditions, but that an effective Lewis number previously defined gives qualitatively correct estimates of the length scales for heat and mass transfer at all pressures.

Description: We report here, preliminary data from an experiment studying flow of superfluid helium through a slit orifice (of sub-micron width) very close to T(sub lambda). Critical supercurrent (I(sub c)) data is obtained from a step function drive to the diaphragm in a Helmholtz resonator cell. The superfluid density (rho(sub s)) data can be obtained from the resonant frequency of the Helmholtz oscillator, as determined by transfer function of the resonator or from the free ringing after the step function excitation. Preliminary data shows that I(sub c) is proportional to (rho(sub s))(exp 1.27) and rho(sub s)) is proportional to tau(exp 0.73), where tau is the reduced temperature. However, the magnitude of I(sub c) is much larger than expected, indicating a possible parallel flow path. Further investigations are in progress. Keywords: superfluid; hydrodynamics; critical exponent

Description: This paper describes and discusses the textbook, Fundamentals of Computational Fluid Dynamics by Lomax, Pulliam, and Zingg, which is intended for a graduate level first course in computational fluid dynamics. This textbook emphasizes fundamental concepts in developing, analyzing, and understanding numerical methods for the partial differential equations governing the physics of fluid flow. Its underlying philosophy is that the theory of linear algebra and the attendant eigenanalysis of linear systems provides a mathematical framework to describe and unify most numerical methods in common use in the field of fluid dynamics. Two linear model equations, the linear convection and diffusion equations, are used to illustrate concepts throughout. Emphasis is on the semi-discrete approach, in which the governing partial differential equations (PDE's) are reduced to systems of ordinary differential equations (ODE's) through a discretization of the spatial derivatives. The ordinary differential equations are then reduced to ordinary difference equations (O(Delta)E's) using a time-marching method. This methodology, using the progression from PDE through ODE's to O(Delta)E's, together with the use of the eigensystems of tridiagonal matrices and the theory of O(Delta)E's, gives the book its distinctiveness and provides a sound basis for a deep understanding of fundamental concepts in computational fluid dynamics.

Description: Buoyant flows play an important role in various technological and environmental issues. For example, dispersal of pollutants, smoke, or volcano exhaust in the atmosphere, vertical motion of air, formation of clouds and other weather systems, and flows in cooling towers and fires are all determined primarily by buoyancy effects. The buoyancy force in such flows can originate from either a heat source or due to different densities between a fluid and its surroundings. Whatever the cause, the flow can be understood by studying the effects of the tight coupling between the thermal and the velocity fields since density differences can be characterized as temperature differences.

Description: Our objective is to predict droplet size distributions created by fuel injector nozzles in Jet turbines. These results will be used to determine the initial conditions for numerical simulations of the combustion process in gas turbine combustors. To predict the droplet size distribution, we are currently constructing a numerical model to understand the instability and breakup of thin conical liquid sheets. This geometry serves as a simplified model of the liquid jet emerging from a real nozzle. The physics of this process is difficult to study experimentally as the time and length scales are very short. From existing photographic data, it does seem clear that three-dimensional effects such as the formation of streamwise ligaments and the pulling back of the sheet at its edges under the action of surface tension are important.

Description: A detailed investigation of the flow physics occurring on the suction side of a simulated Low Pressure Turbine (LPT) blade was performed. A contoured upper wall was designed to simulate the y pressure distribution of an actual LPT blade onto a flat plate. The experiments were carried out at Reynolds numbers of 100,000 and 250,000 with three levels of freestream turbulence. The main emphasis in this paper is placed on flow field surveys performed at a y Reynolds number of 100,000 with levels of freestream turbulence ranging from 0.8% to 3%. Smoke-wire flow visualization data was used to confirm that the boundary layer was separated and formed a bubble. The transition process over the separated flow region is observed to be similar to a laminar free shear layer flow with the formation of a large coherent eddy structure. For each condition, the locations defining the separation bubble were determined by careful examination of pressure and mean velocity profile data. Transition onset location and length determined from intermittency profiles decrease as freestream turbulence levels increase. Additionally, the length and height of the laminar separation bubbles were observed to be inversely proportional to the levels of freestream turbulence.

Description: A distinct boundary between turbulent and non-turbulent regions in a fluid of otherwise constant properties is found in many laboratory and engineering turbulent flows, including jets, mixing layers, boundary layers and wakes. Generally, the flow has mean shear in at least one direction within t he turbulent zone, but the non-turbulent zones have no shear (adjacent laminar shear is a different case, e.g. transition in a boundary layer). There may be purely passive differences between the turbulent and non-turbulent zones, e.g. small variations in temperature or scalar concentration, for which turbulent mixing is an important issue. The boundary has several major characteristics of interest for the present study. Firstly, the boundary advances into the non-turbulent fluid, or in other words, nonturbulent fluid is entrained. Secondly, the change in turbulence properties across the boundary is remarkably abrupt; strong turbulent motions come close to the nonturbulent fluid, promoting entrainment. Thirdly, the boundary is irregular with a continually changing convoluted shape, which produces statistical intermittency. Its shape is contorted at all scales of the turbulent motion.

Description: Turbulent flow simulation methods based on finite differences are attractive for their simplicity, flexibility and efficiency, but not always for accuracy or stability. This report demonstrates that a good compromise is possible with the Advected Grid Explicit (AGE) method. AGE has proven to be both efficient and accurate for simulating turbulent free-shear flows, including planar mixing layers and planar jets. Its efficiency results from its localized fully explicit finite difference formulation (Bisset 1998a,b) that is very straightforward to compute, outweighing the need for a fairly small timestep. Also, most of the successful simulations were slightly under-resolved, and therefore they were, in effect, large-eddy simulations (LES) without a sub-grid-scale (SGS) model, rather than direct numerical simulations (DNS). The principle is that the role of the smallest scales of turbulent motion (when the Reynolds number is not too low) is to dissipate turbulent energy, and therefore they do not have to be simulated when the numerical method is inherently dissipative at its resolution limits. Such simulations are termed 'auto-LES' (LES with automatic SGS modeling) in this report.

Description: Detailed flow field measurements have been carried out for a turbulent circular jet perturbed by tabs and artificial excitation. Two "delta tabs" were placed at the nozzle exit at diametricall opposite y locations. The excitation condition involved subharmonic resonance that manifested in a periodic vortex pairing in the near flow field. While the excitation and the tabs independently increased jet spreading, a combination of the two diminished the effect. The jet spreading was most pronounced with the tabs but was reduced when excitation was applied to the tabbed jet. The tabs generated streamwise vortex pairs that caused a lateral spreading of the jet in a direction perpendicular to the plane containing the tabs. ne excitation, on the other hand, organized the azimuthal vorticity into coherent ring structures whose evolution and pairing also increased entrainment by the jet. In the tabbed case, the excitation produced coherent azimuthal structures that were distorted and asymmetric in shape. The self-induction of these structures produced an effect that opposed the tendency for the lateral spreading of the streamwise vortex pairs. The passage of the distorted vortices, and their pairing, also had a cancellation effect on the time-averaged streamwise vorticity field. These led to the reduction in jet spreading.

Description: The proposed paper will describe the state of the art associated with numerical solution of the full or exact velocity potential equation for solving transonic, external-aerodynamic flows. The presentation will begin with a review of the literature emphasizing research activities of the past decade. Next, the various forms of the full or exact velocity potential equation, the equation's corresponding mathematical characteristics, and the derivation assumptions will be presented and described in detail. Impact of the derivation assumptions on simulation accuracy, especially with respect to shock wave capture, will be presented and discussed relative to the more complete Euler or Navier-Stokes formulations. The technical presentation will continue with a description of recently developed full potential numerical approach characteristics. This description will include governing equation nondimensionalization, physical-to-computational-domain mapping procedures, a limited description of grid generation requirements, the spatial discretization scheme, numerical implementation of boundary conditions, and the iteration scheme. The next portion of the presentation will present and discuss numerical results for several two- and three-dimensional aerodynamic applications. Included in the results section will be a discussion and demonstration of a typical grid refinement analysis for determining spatial convergence of the numerical solution and level of solution accuracy. Computer timings for a variety of full potential applications will be compared and contrasted with similar results for the Euler equation formulation. Finally. the presentation will end with concluding remarks and recommendations for future work.

Description: North American Marine Jet (NAMJ), Inc. received assistance from Marshall Space Flight Center engineers in the Computational Fluid Dynamics (CFD) branch of the Structure and Dynamics Laboratory in improving the proposed design of a new impeller for their jet-propulsion systems. Marshall used advanced CFD techniques, which included creating a three-dimensional computer model of the impeller for analysis. With Marshall input, the company modified the design, then Marshall used a computer model to make a solid polycarbonate model. The rapid prototyping allowed the company to avoid many time- consuming and costly steps in creating the impeller model. NAMJ is now able to compete with Pacific-area and European manufacturers who have traditionally dominated the market.

Description: Small Business Innovation Research contracts from Goddard Space Flight Center to Thermacore Inc. have fostered the company work on devices tagged "heat pipes" for space application. To control the extreme temperature ranges in space, heat pipes are important to spacecraft. The problem was to maintain an 8-watt central processing unit (CPU) at less than 90 C in a notebook computer using no power, with very little space available and without using forced convection. Thermacore's answer was in the design of a powder metal wick that transfers CPU heat from a tightly confined spot to an area near available air flow. The heat pipe technology permits a notebook computer to be operated in any position without loss of performance. Miniature heat pipe technology has successfully been applied, such as in Pentium Processor notebook computers. The company expects its heat pipes to accommodate desktop computers as well. Cellular phones, camcorders, and other hand-held electronics are forsible applications for heat pipes.

Description: Under SBIR (Small Business Innovative Research) contracts with Lewis Research Center, Nektonics, Inc., developed coating process simulation tools, known as Nekton. This powerful simulation software is used specifically for the modeling and analysis of a wide range of coating flows including thin film coating analysis, polymer processing, and glass melt flows. Polaroid, Xerox, 3M, Dow Corning, Mead Paper, BASF, Mitsubishi, Chugai, and Dupont Imaging Systems are only a few of the companies that presently use Nekton.

Description: A common problem occurs when refrigerant fluids wick inside the insulation of thermocouple wires through a compressor's casing feedthrough and then leak into the adjacent disconnect box outside the casing. Leaking fluids create an unfavorable situation inside the disconnect box and may contaminate the fluids. To address this problem, NASA Lewis Research Center s Manufacturing Engineering Division developed a customized hermetic feedthrough for a bank of Worthington compressors. In these compressors, bearing temperatures are measured by internal thermocouples embedded in bearings located inside the compressor casings. The thermocouple wires need to be routed outside the casing and read at another location. These wires are short and are terminated to a disconnect strip inside the casing. The bearings operate at about 170 F, but because the casing is filled with R12 refrigerant oil, the casing has a maximum temperature of about 100 F. The operating conditions of these compressors permit the use of an epoxy that is compatible with the R12 fluid. The desired finished product is a stainless steel tube that has been filled solid with epoxy after thermocouple wires bonded and sealed by epoxy have been inserted through its length. Shrink tubing extends from both ends of the tube. The process that was developed to isolate the thermocouple wires from the R12 fluid follows. For this application, use an 8-in.-long piece of 0.500-in. 304 stainless steel tube with six pairs of 24-gauge stranded, PTFE-insulated (polytetrafluoroethylene) type "T" thermocouple wires for each feedthrough. Use shrink tubing to strain relief the insulated wires at their exit from the stainless steel tube. Cut the wire to length and identify the location of the stainless steel tube sleeve with masking tape. Then, remove the outer insulation from a 2-in. section of wire that will be inside the tube, and carefully strip to bare wire a 1-in. section in the middle of the section with the outer insulation removed. For an effective seal, the epoxy must penetrate between the strands when stranded conductors are used. Make the seal with epoxy bond on the bare wire. The bare wire must be encapsulated with a thin layer of the epoxy that leaves only a very low profile. These encapsulated wires must cure before the assembly can be continued. Then, inspect the cured wires for complete encapsulation before going to the next step. Insert the wires in the stainless steel tube and orient them so that the epoxied stripped sections are staggered within the tube; then, apply shrink tubing to one end of the cleaned wires, positioning it inside the edge of the tube. The small gaps between the wires on the other end will be used to inject the epoxy into the tube. Let the epoxy cure inside the tube, free of any voids. Then, continue to fill the tube until the entire 8-in. length is nearly filled, allowing room for the other strain-relieving shrink tubing. Since this first design, the process has been adjusted to fit many needs and situations. Customized feedthroughs have been assembled from various wire types, wire gauges, and/or stainless steel tube passages. The fittings selected to mount these feedthroughs allow their use in other areas, such as pressure or vacuum systems.

Description: High-Speed Civil Transport (HSCT) engine combustor liners are subjected to complex thermal environments and have to endure these for thousands of hours with assured reliability. In the past, several deterministic analyses have been performed, including detailed heat transfer analyses to obtain thermal profiles and deterministic stress analyses to identify critical locations of high stresses. Actual rig tests also have been performed for segments by simulating these loading situations as closely as possible. However, it is well known that many uncertainties exist in loading (primarily thermal loads due to heat transfer), boundary conditions (end fixity unknowns), and material properties (moduli, thermal-expansion coefficients, and conductivities). The present in-house effort at the NASA Lewis Research Center is directed toward accounting for these in a formal way to assess the performance of liner components under complex and uncertain loading conditions as well as subject to other geometry- and material-related uncertainties.

Description: A major portion of the energy produced in the world today comes from the burning of liquid hydrocarbon fuels in the form of droplets. Understanding the fundamental physical processes involved in droplet combustion is not only important in energy production but also in propulsion, in the mitigation of combustion-generated pollution, and in the control of the fire hazards associated with handling liquid combustibles. Microgravity makes spherically symmetric combustion possible, allowing investigators to easily validate their droplet models without the complicating effects of gravity. The Fiber-Supported Droplet Combustion (FSDC-2) investigation was conducted in the Microgravity Glovebox facility of the shuttles' Spacelab during the reflight of the Microgravity Science Laboratory (MSL- 1R) on STS-94 in July 1997. FSDC-2 studied fundamental phenomena related to liquid fuel droplet combustion in air. Pure fuels and mixtures of fuels were burned as isolated single and duo droplets with and without forced air convection. FSDC-2 is sponsored by the NASA Lewis Research Center, whose researchers are working in cooperation with several investigators from industry and academia. The rate at which a droplet burns is important in many commercial applications. The classical theory of droplet burning assumes that, for an isolated, spherically symmetric, single-fuel droplet, the gas-phase combustion processes are much faster than the droplet surface regression rate and that the liquid phase is at a uniform temperature equal to the boiling point. Recent, more advanced models predict that both the liquid and gas phases are unsteady during a substantial portion of the droplet's burning history, thus affecting the instantaneous and average burning rates, and that flame radiation is a dominant mechanism that can extinguish flames in a microgravity environment. FSDC-2 has provided well-defined, symmetric droplet burning data including radiative emissions to validate these theoretical models for heptane, decane, ethanol, and methanol fuels. Since most commercial combustion systems burn droplets in a convective environment, data were obtained without and with convective flow over the burning droplet (see the following photos).

Description: This in-house experiment on the gust response of an annular turbine cascade was particularly designed to provide data to compare with the results of a typical, linearized gust-response analysis. Reduced frequency, Mach number, and incidence were varied independently. Except for the lowest reduced frequency, the gust velocity distribution was nearly sinusoidal. For the high inlet-velocity series of tests, the cascade was near choking. The mean flow was documented by measuring blade surface pressures and the cascade exit flow, and high-response pressure transducers were used to measure the unsteady pressure distribution. Inlet-velocity components and turbulence parameters were measured using hot wire anemometry. In addition to the synchronous time-averaged pressure spectra, typical power spectra are included for several representative conditions.

Description: Through a combination of experimental and computational studies, researchers at the NASA Lewis Research Center investigated the effect of upstream blade-row wake passing on the showerhead (leading edge) film cooling of a downstream turbine. The experiments were performed in a steady-flow annular turbine cascade facility equipped with an upstream rotating row of cylindrical rods to produce a periodic wake field similar to that found in an actual turbine. Spanwise, chordwise, and temporal resolution of the blade surface temperature were achieved through the use of an array of nickel thin-film surface gauges covering one unit cell of a showerhead film hole pattern. Film effectiveness and Nusselt numbers were determined for a test matrix of various injectants, injectant blowing ratios, and wake Strouhal numbers (St).

Description: The NASA Lewis Research Center is developing analytical methods and software tools to create a bridge between the controls and computational fluid dynamics (CFD) disciplines. Traditionally, control design engineers have used coarse nonlinear simulations to generate information for the design of new propulsion system controls. However, such traditional methods are not adequate for modeling the propulsion systems of complex, high-speed vehicles like the High Speed Civil Transport. To properly model the relevant flow physics of high-speed propulsion systems, one must use simulations based on CFD methods. Such CFD simulations have become useful tools for engineers that are designing propulsion system components. The analysis techniques and software being developed as part of this effort are an attempt to evolve CFD into a useful tool for control design as well. One major aspect of this research is the generation of linear models from steady-state CFD results. CFD simulations, often used during the design of high-speed inlets, yield high resolution operating point data. Under a NASA grant, the University of Akron has developed analytical techniques and software tools that use these data to generate linear models for control design. The resulting linear models have the same number of states as the original CFD simulation, so they are still very large and computationally cumbersome. Model reduction techniques have been successfully applied to reduce these large linear models by several orders of magnitude without significantly changing the dynamic response. The result is an accurate, easy to use, low-order linear model that takes less time to generate than those generated by traditional means. The development of methods for generating low-order linear models from steady-state CFD is most complete at the one-dimensional level, where software is available to generate models with different kinds of input and output variables. One-dimensional methods have been extended somewhat so that linear models can also be generated from two- and three-dimensional steady-state results. Standard techniques are adequate for reducing the order of one-dimensional CFD-based linear models. However, reduction of linear models based on two- and three-dimensional CFD results is complicated by very sparse, ill-conditioned matrices. Some novel approaches are being investigated to solve this problem.

Description: Coupled Aerodynamic-Thermal-Structural (CATS) Analysis is a focused effort within the Numerical Propulsion System Simulation (NPSS) program to streamline multidisciplinary analysis of aeropropulsion components and assemblies. Multidisciplinary analysis of axial-flow compressor performance has been selected for the initial focus of this project. CATS will permit more accurate compressor system analysis by enabling users to include thermal and mechanical effects as an integral part of the aerodynamic analysis of the compressor primary flowpath. Thus, critical details, such as the variation of blade tip clearances and the deformation of the flowpath geometry, can be more accurately modeled and included in the aerodynamic analyses. The benefits of this coupled analysis capability are (1) performance and stall line predictions are improved by the inclusion of tip clearances and hot geometries, (2) design alternatives can be readily analyzed, and (3) higher fidelity analysis by researchers in various disciplines is possible. The goals for this project are a 10-percent improvement in stall margin predictions and a 2:1 speed-up in multidisciplinary analysis times. Working cooperatively with Pratt & Whitney, the Lewis CATS team defined the engineering processes and identified the software products necessary for streamlining these processes. The basic approach is to integrate the aerodynamic, thermal, and structural computational analyses by using data management and Non-Uniform Rational B-Splines (NURBS) based data mapping. Five software products have been defined for this task: (1) a primary flowpath data mapper, (2) a two-dimensional data mapper, (3) a database interface, (4) a blade structural pre- and post-processor, and (5) a computational fluid dynamics code for aerothermal analysis of the drum rotor. Thus far (1) a cooperative agreement has been established with Pratt & Whitney, (2) a Primary Flowpath Data Mapper has been prototyped and delivered to General Electric Aircraft Engines and Pratt & Whitney for evaluation, (3) a collaborative effort has been initiated with the National Institute of Standards and Testing to develop a Standard Data Access Interface, and (4) a blade tip clearance capability has been implemented into the Structural Airfoil Blade Engineering Routine (SABER) program. We plan to continue to develop the data mappers and data management tools. As progress is made, additional efforts will be made to apply these tools to propulsion system applications.

Description: Gaseous oxygen and hydrogen propellants used in a special engine energy cycle called Full-Flow Staged Combustion are believed to significantly increase the lifetime of a rocket engine's pumps. The cycle can also reduce the operating temperatures of the engine. Improving the lifetime of the hardware reduces its overall maintenance and operations costs, and is critical to reducing costs for the joint NASA/industry Reusable Launch Vehicle (RLV). The work in this project will demonstrate the performance and lifetime of one-element and many-element combustors with gaseous O2/H2 injectors. This work supporting the RLV program is a cooperative venture of the NASA Lewis Research Center, the NASA Marshall Space Flight Center, Rocketdyne, and the Pennsylvania State University. Information about gas-gas rocket injector performance with O2/H2 is very limited. Because of this paucity of data, new testing is needed to improve the knowledge base for testing and designing new injectors for the RLV and to improve computer models that predict the combusting gas flows of new injector designs. Therefore, detailed observations and measurements of the combusting flow from many-element injectors in a rocket engine are being sought. These observations and measurements will be done with three different tools: schlieren photography, ultraviolet imaging, and Raman spectroscopy. The schlieren system will take photos of the density differences in combusting flow, the ultraviolet movies will determine the location of the hydroxyl (OH) radical in the combustion flow, and the Raman spectroscopic measurements will provide the combustion temperature and amount of water (H2O), hydrogen (H2), and oxygen (O2) in the combustor. Marshall is providing overall program management, design and computational fluid dynamics (CFD) analyses, as well as funding for the work at Penn State. An existing, windowed combustor and several injectors will be provided by Rocketdyne--two injectors for the initial screening tests and one with an optimized design based on the best design found in the screening tests. Lewis will provide a nozzle and several injectors for the screening test program. The configuration of the injectors will be based on a design chosen by all the participants, and their elements will be based on the coaxial and impinging flow. Lewis also will provide the instrumentation for the flow-field measurements: schlieren, ultraviolet imaging, and Raman spectroscopy. In addition, thermocouples will measure heat flow on the injector face. Other traditional measurements of rocket performance will be made as well: chamber pressure, mass flow of each propellant, purge flow, and the barrier cooling gas flow. Penn State will conduct single-element testing with the injector elements from both the Rocketdyne and the jointly designed injectors. A wide variety of traditional and nontraditional injector designs will be tested in this program. The results will be valuable in computational fluid dynamics code validation and overall rocket combustion efficiency measurements. Correlations between combustion efficiency, laser measurements of species, and ultraviolet and visible light photography will also be made. Thus far, several different single-element injectors have been tested at Penn State and Lewis. The experimental setup of a rocket engine with a viewing window is shown. The combusting flow is shown. The results are helping engineers design the many element injectors.

Description: On a swept wing, contamination along the leading edge, Tollmien-Schlichting waves, stationary or traveling crossflow vortices, and/or Taylor-Gortler vortices can cause the catastrophic breakdown of laminar to turbulent flow, which leads to increased skin-friction drag for the aircraft. The discussion in this Note will be limited to disturbances which evolve along the attachment line (leading edge of swept wing). If the Reynolds number of the attachment-line boundary layer is greater than some critical value, then the complete wing is inevitably engulfed in turbulent flow. Essentially, there are two critical Reynolds number points that must be considered. The first is for small-amplitude disturbances, and the second is for bypass transition. The present study will use direct numerical simulations to validate a linear 2D-eigenvalue prediction method based on parabolized stability equations by Lin and Malik. This method is considered because it suggests that a number of symmetric and asymmetric modes exist and are stable or unstable on the attachment line depending on the Reynolds number. If validated, the approach would predict a number of modes which are linearly damped in the Reynolds number regime 100 to 245; however, these modes may grow nonlinearly and provide an explanation to this region.

Description: The entropy associated with absolute equilibrium ensemble theories of ideal, homogeneous, fluid and magneto-fluid turbulence is discussed and the three-dimensional fluid case is examined in detail. A sigma-function is defined, whose minimum value with respect to global parameters is the entropy. A comparison is made between the use of global functions sigma and phase functions H (associated with the development of various H-theorems of ideal turbulence). It is shown that the two approaches are complimentary though conceptually different: H-theorems show that an isolated system tends to equilibrium while sigma-functions allow the demonstration that entropy never decreases when two previously isolated systems are combined. This provides a more complete picture of entropy in the statistical mechanics of ideal fluids.

Description: The spatial evolution of three-dimensional disturbances in an attachment-line boundary layer is computed by direct numerical simulation of the unsteady, incompressible Navier-Stokes equations. Disturbances are introduced into the boundary layer by harmonic sources that involve unsteady suction and blowing through the wall. Various harmonic- source generators are implemented on or near the attachment line, and the disturbance evolutions are compared. Previous two-dimensional simulation results and nonparallel theory are compared with the present results. The three-dimensional simulation results for disturbances with quasi-two-dimensional features indicate growth rates of only a few percent larger than pure two-dimensional results; however, the results are close enough to enable the use of the more computationally efficient, two-dimensional approach. However, true three-dimensional disturbances are more likely in practice and are more stable than two-dimensional disturbances. Disturbances generated off (but near) the attachment line spread both away from and toward the attachment line as they evolve. The evolution pattern is comparable to wave packets in at-plate boundary-layer flows. Suction stabilizes the quasi-two-dimensional attachment-line instabilities, and blowing destabilizes these instabilities; these results qualitatively agree with the theory. Furthermore, suction stabilizes the disturbances that develop off the attachment line. Clearly, disturbances that are generated near the attachment line can supply energy to attachment-line instabilities, but suction can be used to stabilize these instabilities.

Description: Results are presented of an experimental program that investigated the use of a secondary air stream to control the amount of flow through a convergent-divergent nozzle. These static tests utilized high pressure, ambient temperature air that was injected at the throat of the nozzle through an annular slot. Multiple injection slot sizes and injection angles were tested. The introduction of secondary flow was made in an opposing direction to the primary flow and the resulting flow field caused the primary stream to react as though the physical throat size had been reduced. The percentage reduction in primary flow rate was generally about twice the injected flow rate. The most effective throttling was achieved by injecting through the smallest slot in an orientation most nearly opposed to the approaching primary flow. Thrust edliciency, as measured by changes in nozzle thrust coefficient, was highest at high nozzle pressure ratios, NPR. The static test results agreed with predictions obtained prior from PABSD, a fully viscous computational fluid dynamics program. Since use of such an injection system on gas turbine engine exhaust nozzles would be primarily at high NPRs, it was concluded that fluidic control holds promise for reducing nozzle weight and complexity on future systems.

Description: The component integration of a class of hypersonic high-lift configurations known as waveriders into hypersonic cruise vehicles was evaluated. A wind-tunnel model was developed which integrates realistic vehicle components with two waverider shapes, referred to as the straight-wing and cranked-wing shapes. Both shapes were conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) predictions were obtained over a Mach number range of 1.6 to 4.63 at a Reynolds number of 2.0 x 10(exp 6) per foot. The CFD predictions and flow visualization data confirmed the shock attachment characteristics of the baseline waverider shapes and illustrated the waverider flow-field properties. Experimental data showed that no significant performance degradations, in terms of maximum lift-to-drag ratios, occur at off-design Mach numbers for the waverider shapes and the integrated configurations. A comparison of the fully-integrated waverider vehicles to the baseline shapes showed that the performance was significantly degraded when all of the components were added to the waveriders, with the most significant degradation resulting from aftbody closure and the addition of control surfaces. Both fully-integrated configurations were longitudinally unstable over the Mach number range studied with the selected center of gravity location and for unpowered conditions. The cranked-wing configuration provided better lateral-directional stability characteristics than the straight-wing configuration.

Description: A new procedure seeks to combine the thin-layer Navier-Stokes solver LAURA with the parabolized Navier-Stokes solver UPS for the aerothermodynamic solution of chemically-reacting air flow fields. The interface protocol is presented and the method is applied to two slender, blunted shapes. Both axisymmetric and three-dimensional solutions are included with surface pressure and heat transfer comparisons between the present method and previously published results. The case of Mach 25 flow over an axisymmetric six degree sphere-cone with a non-catalytic wall is considered to 100 nose radii. A stability bound on the marching step size was observed with this case and is attributed to chemistry effects resulting from the non-catalytic wall boundary condition. A second case with Mach 28 flow over a sphere-cone-cylinder-flare configuration is computed at both two and five degree angles of attack with a fully-catalytic wall. Surface pressures are seen to be within five percent with the present method compared to the baseline LAURA solution and heat transfers are within 10 percent. The effect of grid resolution is investigated in both the radial and streamwise directions. The procedure demonstrates significant, order of magnitude reductions in solution time and required memory for the three-dimensional case in comparison to an all thin-layer Navier-Stokes solution.

Description: The linear and the nonlinear stability of disturbances that propagate along the attachment line of a three-dimensional boundary layer is considered. The spatially evolving disturbances in the boundary layer are computed by direct numerical simulation (DNS) of the unsteady, incompressible Navier-Stokes equations. Disturbances are introduced either by forcing at the in ow or by applying suction and blowing at the wall. Quasi-parallel linear stability theory and a nonparallel theory yield notably different stability characteristics for disturbances near the critical Reynolds number; the DNS results con rm the latter theory. Previously, a weakly nonlinear theory and computations revealed a high wave-number region of subcritical disturbance growth. More recent computations have failed to achieve this subcritical growth. The present computational results indicate the presence of subcritically growing disturbances; the results support the weakly nonlinear theory. Furthermore, an explanation is provided for the previous theoretical and computational discrepancy. In addition, the present results demonstrate that steady suction can be used to stabilize disturbances that otherwise grow subcritically along the attachment line.

Description: The year 1971 saw the publication of one of the landmark papers in computational aerodynamics, that of Murman and Cole. As with many seminal works, its significance lies not so much in the specific problem that it addressed| small disturbance, plane transonic flow - but in the identification of a general approach to the solution of a technically important and theoretically difficult problem. The key features of Murman and Cole's work were the use of type- dependent differencing to correctly account for the proper domain of dependence of a mixed elliptic/hyperbolic equation, and the introduction of line relaxation to solve the steady flow equation. All subsequent work in transonic potential flows was based on these concepts. Jameson extended Murman and Cole's ideas to the full potential equation with two important contributions. First, he introduced the rotated difference stencil, which generalized the Murman and Cole type-dependent difference operator to general coordinates. Second, he used the interpretation, introduced by Garabedian, of relaxation as an iteration in artificial time to construct stable relaxation schemes, generalizing the original line relaxation method of Reference. The decade of the 1970s saw an explosion of activity in the solution of transonic potential flows, which has been summarized in the review article of Caughey.

Description: The paper summarizes Active Flow Control projects currently underway at the NASA Langley Research Center. Technology development is being pursued within a multidisciplinary, cooperative approach, involving the classical disciplines of fluid mechanics, structural mechanics, material science, acoustics, and stability and control theory. Complementing the companion papers in this session, the present paper will focus on projects that have the goal of extending the state-of-the-art in the measurement, prediction, and control of unsteady, nonlinear aerodynamics. Toward this goal, innovative actuators, micro and macro sensors, and control strategies are considered for high payoff flow control applications. The target payoffs are outlined within each section below. Validation of the approaches range from bench-top experiments to wind-tunnel experiments to flight tests. Obtaining correlations for future actuator and sensor designs are implicit in the discussion. The products of the demonstration projects and design tool development from the fundamental NASA R&D level technology will then be transferred to the Applied Research components within NASA, DOD, and US Industry. Keywords: active flow control, separation control, MEMS, review

Description: This paper presents the results of a numerical study of shock/shock interactions that include both the Edney type IV and type III interactions, with emphasis on the type IV interactions. Computations are made using the direct simulation Monte Carlo (DSMC) method of Bird for Mach 10 air flow, as produced in the ONERA R5Ch low-density wind tunnel. The simulations include the flow about a shock generator which creates a relatively weak oblique shock that impinges on a much stronger cylinder bow shock. The sensitivity and characteristics of the interactions are examined by varying the horizontal distance separating the shock generator leading edge and cylinder. Results of the simulation for one separation distance are compared with wind tunnel measurements. Comparisons are made for surface heating and pressure and for flow-field values of density and rotational temperatures, as obtained with the Dual-line Coherent Anti-Stokes Scattering (DL-CARS) technique. The comparisons between experiment and calculation yield a consistent description of the shock interaction features and a consistent description of the surface heating and pressure distributions, with the exception of the peak values-the computed values being greater than the measured values.

Description: Problems concerning laminar-turbulent transition are addressed by solving a series of initial value problems. Solutions to the temporal, initial-value problem .with an inhomogeneous forcing term imposed upon the flow are sought. It is shown that: (1) A transient disturbance lying located outside of the boundary layer can lead to the growth of an unstable Tollmein-Schlicting wave; (2) A resonance with the continuous spectrum may provide a mechanism for bypass transition; and (3) The continuum modes of a disturbance feed directly into the Tollmein-Schlicting wave downstream through non-parallel effects.

Description: At the NASA Lewis Research Center, a comprehensive assessment was made of the predictive capability of the average passage flow model as applied to multistage axial-flow compressors. This model, which describes the time-averaged flow field within a typical passage of a blade row embedded in a multistage configuration, is being widely used throughout U.S. aircraft industry as an integral part of their design systems. Rotor flow-angle deviation. In this work, detailed data taken within a four and one-half stage large low-speed compressor were used to assess the weaknesses and strengths of the predictive capabilities of the average passage flow model. The low-speed compressor blading is of modern design and employs stator end-bends. Measurements were made with slow- and high response instrumentation. The high-response measurements revealed the velocity components of both the rotor and stator wakes. From the measured wake profiles, we found that the flow exiting the rotors deviated from the rotor exit metal angle to a lesser degree than was predicted by the average passage flow model. This was found to be due to blade boundary layer transition, which recently has been shown to exist on multistage axial compressor rotor and stator blades, but was not accounted for in the average passage model. Consequently, a model that mimics the effects of blade boundary layer transition, Shih k-epsilon model, was incorporated into the average passage model. Simulations that incorporated this transition model showed a dramatic improvement in agreement with data. The altered model thus improved predictive capability for multistage axial-flow compressors, and this was verified by detailed experimental measurement.

Description: A thorough understanding of dynamic interactions between inlets and compressors is extremely important to the design and development of propulsion control systems, particularly for supersonic aircraft such as the High-Speed Civil Transport (HSCT). Computational fluid dynamics (CFD) codes are routinely used to analyze individual propulsion components. By coupling the appropriate CFD component codes, it is possible to investigate inlet-compressor interactions. The objectives of this work were to gain a better understanding of inlet-compressor interaction physics, formulate a more realistic compressor-face boundary condition for time-accurate CFD simulations of inlets, and to take a first step toward the CFD simulation of an entire engine by coupling multidimensional component codes. This work was conducted at the NASA Lewis Research Center by a team of civil servants and support service contractors as part of the High Performance Computing and Communications Program (HPCCP).

Description: The design and implementation of Doppler Global Velocimetry (DGV) for testing in the Langley Unitary Plan Wind Tunnel is presented. The discussion begins by outlining the characteristics of the tunnel and the test environment, with potential problem areas highlighted. Modifications to the optical system design to implement solutions for these problems are described. Since this tunnel entry was the first ever use of DGV in a supersonic wind tunnel, the test series was divided into three phases, each with its own goal. Phase I determined if condensation provided sufficient scattered light for DGV applications. Phase II studied particle lag by measuring the flow about an oblique shock above an inclined flat plate. Phase III investigated the supersonic vortical flow field above a 75-degree delta wing at 24-degrees angle of attack. Example results from these tests are presented.

Description: As opposed to previous explanations based on the effects of anharmonicity of simple diatomic molecules, traces of water vapor are suggested to be the most likely cause of the anomalously fast vibrational relaxation of such gases observed in supersonic and hypersonic nozzles. The mechanism is the strong V-VR coupling with H2O molecules that dramatically facilitates the collisional transfer of vibrational energy. Slight moisture content is thus a real world aspect of gas dynamics that must be considered in characterizations of shock tubes, reflected shock tunnels, and expansion tubes.

Description: A computational method for accurately predicting the static aeroelastic deformations of typical transonic transport wind tunnel models is described. The method utilizes a finite element method (FEM) for predicting the deformations. Extensive calibration/validation of this method was carried out using a novel wind-off wind tunnel model static loading experiment and wind-on optical wing twist measurements obtained during a recent wind tunnel test in the National Transonic Facility (NTF) at NASA LaRC. Further validations were carried out using a Navier-Stokes computational fluid dynamics (CFD) flow solver to calculate wing pressure distributions about several aeroelastically deformed wings and comparing these predictions with NTF experimental data. Results from this aeroelastic deformation method are in good overall agreement with experimentally measured values. Including the predicted deformations significantly improves the correlation between CFD predicted and experimentally measured wing & pressures.

Description: Boundary-layer bleed is an invaluable tool for controlling the airflow in supersonic aircraft engine inlets. Incoming air is decelerated to subsonic speeds prior to entering the compressor via a series of oblique shocks. The low momentum flow in the boundary layer interacts with these shocks, growing in thickness and, under some conditions, leading to flow separation. To remedy this, bleed holes are strategically located to remove mass from the boundary layer, reducing its thickness and helping to maintain uniform flow to the compressor. The bleed requirements for any inlet design are unique and must be validated by extensive wind tunnel testing to optimize performance and efficiency. To accelerate this process and reduce cost, researchers at the NASA Lewis Research Center initiated an experimental program to study the flow phenomena associated with bleed holes. Knowledge of these flow properties will be incorporated into computational fluid dynamics (CFD) models that will aid engine inlet designers in optimizing bleed configurations before any hardware is fabricated. This ongoing investigation is currently examining two hole geometries, 90 and 20 (both with 5-mm diameters), and various flow features.

Description: As part of a contract with the NASA Lewis Research Center, a simple, accurate method of predicting the performance characteristics of a nozzle design has been developed for use in conceptual design studies. The Nozzle Performance Analysis Code (NPAC) can predict the on- and off-design performance of axisymmetric or two-dimensional convergent and convergent-divergent nozzle geometries. NPAC accounts for the effects of overexpansion or underexpansion, flow divergence, wall friction, heat transfer, and small mass addition or loss across surfaces when the nozzle gross thrust and gross thrust coefficient are being computed. NPAC can be used to predict the performance of a given nozzle design or to develop a preliminary nozzle system design for subsequent analysis. The input required by NPAC consists of a simple geometry definition of the nozzle surfaces, the location of key nozzle stations (entrance, throat, exit), and the nozzle entrance flow properties. NPAC performs three analysis "passes" on the nozzle geometry. First, an isentropic control volume analysis is performed to determine the gross thrust and gross thrust coefficient of the nozzle. During the second analysis pass, the skin friction and heat transfer losses are computed. The third analysis pass couples the effects of wall shear and heat transfer with the initial internal nozzle flow solutions to produce a system of equations that is solved at steps along the nozzle geometry. Small mass additions or losses, such as those resulting from leakage or bleed flow, can be included in the model at specified geometric sections. A final correction is made to account for divergence losses that are incurred if the nozzle exit flow is not purely axial.

Description: Several schemes for introducing an artificial dissipation into a central difference approximation to the Euler and Navier Stokes equations are considered. The focus of the paper is on the convective upwind and split pressure (CUSP) scheme, which is designed to support single interior point discrete shock waves. This scheme is analyzed and compared in detail with scalar dissipation and matrix dissipation (MATD) schemes. Resolution capability is determined by solving subsonic, transonic, and hypersonic flow problems. A finite-volume discretization and a multistage time-stepping scheme with multigrid are used to compute solutions to the flow equations. Numerical solutions are also compared with either theoretical solutions or experimental data. For transonic airfoil flows the best accuracy on coarse meshes for aerodynamic coefficients is obtained with a simple MATD scheme. The coarse-grid accuracy for the original CUSP scheme is improved by modifying the limiter function used with the scheme, giving comparable accuracy to that obtained with the MATD scheme. The modifications reduce the background dissipation and provide control over the regions where the scheme can become first order.

Description: Aircraft laminar flow control (LFC) from the 1930's through the 1990's is reviewed and the current status of the technology is assessed. Examples are provided to demonstrate the benefits of LFC for subsonic and supersonic aircraft. Early studies related to the laminar boundary-layer flow physics, manufacturing tolerances for laminar flow, and insect-contamination avoidance are discussed. LFC concept studies in wind-tunnel and flight experiments are the major focus of the paper. LFC design tools are briefly outlined for completeness.

Description: NASA Lewis Research Center's Advanced Controls and Dynamics Technology Branch is investigating active control strategies to mitigate or eliminate the combustion instabilities prevalent in lean-burning, low-emission combustors. These instabilities result from coupling between the heat-release mechanisms of the burning process and the acoustic flow field of the combustor. Control design and implementation require a simulation capability that is both fast and accurate. It must capture the essential physics of the system, yet be as simple as possible. A quasi-one-dimensional, computational fluid dynamics (CFD) based simulation has been developed which may meet these requirements. The Euler equations of mass, momentum, and energy have been used, along with a single reactive species transport equation to simulate coupled thermoacoustic oscillations. A very simple numerical integration scheme was chosen to reduce computing time. Robust boundary condition procedures were incorporated to simulate various flow conditions (e.g., valves, open ends, and choked inflow) as well as to accommodate flow reversals that may arise during large flow-field oscillations. The accompanying figure shows a sample simulation result. A combustor with an open inlet, a choked outlet, and a large constriction approximately two thirds of the way down the length is shown. The middle plot shows normalized, time-averaged distributions of the relevant flow quantities, and the bottom plot illustrates the acoustic mode shape of the resulting thermoacoustic oscillation. For this simulation, the limit cycle peak-to-peak pressure fluctuations were 13 percent of the mean. The simulation used 100 numerical cells. The total normalized simulation time was 50 units (approximately 15 oscillations), which took 26 sec on a Sun Ultra2.

Description: Spacecraft solar dynamic power systems typically use high-temperature phase-change materials to efficiently store thermal energy for heat engine operation in orbital eclipse periods. Lithium fluoride salts are particularly well suited for this application because of their high heat of fusion, long-term stability, and appropriate melting point. Considerable attention has been focused on the development of thermal energy storage (TES) canisters that employ either pure lithium fluoride (LiF), with a melting point of 1121 K, or eutectic composition lithium-fluoride/calcium-difluoride (LiF-20CaF2), with a 1040 K melting point, as the phase-change material. Primary goals of TES canister development include maximizing the phase-change material melt fraction, minimizing the canister mass per unit of energy storage, and maximizing the phase-change material thermal charge/discharge rates within the limits posed by the container structure.

Description: Pulse Phase Thermography (PPT) has been reported as a novel powerful technique of the thermal NDE. It employs application of the Discrete Fourier Transform (DFT) to thermal images obtained following flash heating of the front surface of a specimen. The computed phasegrams are excellent for defect visualization in a wide range of materials. This is in part due to their low sensitivity to uneven heating. This work is an attempt to analyze advantages and limitations of PPT. Results of application of the DFT to simulated temperature decays are presented. The temperature evolution on a surface has been simulated based on an analytical solution of the one-dimensional heat diffusion problem. A more sophisticated study has been done for different sizes of defects using numerical solution of the three-dimensional mathematical model. Capabilities of PPT for in-depth scanning and for monitoring of the material loss are discussed. The recommendations for the practical application of the PPT are presented. Experimental results obtained following these recommendations are reported.

Description: Inertial and gravitational forces determine a particle's motion in a turbulent flow field. Gravity plays the dominant role in this motion by pulling the particles through adjacent regions of fluid turbulence. To better understand and model how a particle's inertia effects its displacement, one must examine the dispersion in a turbulent flow in the absence of gravity. In this paper, we present the particle experiments planned for NASA's KC-135 Reduced-Gravity Aircraft, which generates microgravity conditions for about 20 seconds. We also predict the particle behavior using simulation and ground-based experiments. We will release particles with Stokes numbers of 0.1, 1, and 10 into an enclosed tank of near-isotropic, stationary, and homogenous turbulence. These particle Stoke numbers cover a broad range of flow regimes of interest. Two opposed grids oscillating back and forth generate the turbulent field in the tank with a range of turbulence scales that covers about three orders of magnitude and with turbulence intensities of about ten times the mean velocity. The motion of the particles will be tracked using a stereo image velocimetry technique.

Description: Experimental results are presented for drops and bubbles levitated in a liquid host, with particular attention given to the effect of shape oscillations and capillary waves on the local flow fields. Some preliminary results are also presented on the use of streaming flows for the control of evaporation rate and rotation of electrostatically levitated droplets in 1 g. The results demonstrate the potential for the technological application of acoustic methods to active control of forced convection in microgravity.

Description: We have developed a new instrument that can measure fast transient birefringence and polymer chain orientation angle in complex fluids. The instrument uses a dual-crystal transverse electro-optic modulator with the second crystal's modulation voltage applied 180 deg out of phase from that of the first crystal. In this manner, the second crystal compensates for the intrinsic static birefringence of the first crystal, and it doubles the modulation depth. By incorporating a transverse electro-optic modulator with two lithium-niobate (LiNbO3) crystals oriented orthogonal to each other with a custom-designed optical system, we have produced a very small robust instrument capable of fast transient retardation measurements. By measuring the sample thickness or optical path length through the sample, we can calculate the transient birefringence. This system can also measure dichroism. We have compared the calibration results and retardation and orientation angle measurements of this instrument with those of a photoelastic modulator (PEM) based system using a quarter wave plate and a high-precision 1/16-wave plate to simulate a birefringent sample. Transient birefringence measurements on the order of 10(exp -9) can be measured using either modulator.

Description: The bulk behavior of dispersed, fluidized, or undispersed stationary granular systems cannot be fully understood in terms of adhesive/cohesive properties without understanding the role of electrostatic forces acting at the level of the grains themselves. When grains adhere to a surface, or come in contact with one another in a stationary bulk mass, it is difficult to measure the forces acting on the grains, and the forces themselves that induced the cohesion and adhesion are changed. Even if a single grain were to be scrutinized in the laboratory, it might be difficult, perhaps impossible, to define the distribution and character of surface charging and the three-dimensional relationship that charges (electrons, holes) have to one another. The hypothesis that we propose to test in microgravity (for dielectric materials) is that adhesion and cohesion of granular matter are mediated primarily by dipole forces that do not require the presence of a net charge; in fact, nominally electrically neutral materials should express adhesive and cohesive behavior when the neutrality results from a balance of positive and negative charge carriers. Moreover, the use of net charge alone as a measure of the electrical nature of grain-to-grain relationships within a granular mass may be misleading. We believe that the dipole forces arise from the presence of randomly-distributed positive and negative fixed charge carriers on grains that give rise to a resultant dipole moment. These dipole forces have long-range attraction. Random charges are created whenever there is triboelectrical activity of a granular mass, that is, whenever the grains experience contact/separation sequences or friction.

Description: The present work is aimed at the experimental studies and numerical modeling of the bubble growth mechanisms of a single bubble attached to a heating surface and of a bubble sliding along an inclined heated plate. Single artificial cavity of 10 microns in diameter was made on the polished Silicon wafer which was electrically heated at the back side in order to control the surface nucleation superheat. Experiments with a sliding bubble were conducted at different inclination angles of the downward facing heated surface for the purpose of studying the effect of magnitude of components of gravity acting parallel to and normal to the heat transfer surface. Information on the bubble shape and size, the bubble induced liquid velocities as well as the surface temperature were obtained using the high speed imaging and hydrogen bubble techniques. Analytical/numerical models were developed to describe the heat transfer through the micro-macro layer underneath and around a bubble formed at a nucleation site. In the micro layer model the capillary and disjoining pressures were included. Evolution of the bubble-liquid interface along with induced liquid motion was modeled. As a follow-up to the studies at normal gravity, experiments are being conducted in the KC-135 aircraft to understand the bubble growth/detachment under low gravity conditions. Experiments have been defined to be performed under long duration of microgravity conditions in the space shuttle. The experiment in the space shuttle will provide bubble growth and detachment data at microgravity and will lead to validation of the nucleate boiling heat transfer model developed from the preceding studies conducted at normal and low gravity (KC-135) conditions.

Description: In this work, we study steady and oscillatory thermocapillary and natural convective flows generated by a bubble on a heated solid surface. The interaction between gas and vapor bubbles with the surrounding fluid is of interest for both space and ground-based processing. A combined numerical-experimental approach is adopted here. The temperature field is visualized using Mach-Zehnder and/or Wollaston Prism Interferometry and the flow field is observed by a laser sheet flow visualization technique. A finite element numerical model is developed which solves the transient two-dimensional continuity, momentum, and energy equations and includes the effects of temperature-dependent surface tension and bubble surface deformation. Below the critical Marangoni number, the steady state low-g and 1-g temperature and velocity fields predicted by the finite element model are in excellent agreement with both the visualization experiments in our laboratory and recently published experimental results in the literature. Above the critical Marangoni number, the model predicts an oscillatory flow which is also closely confirmed by experiments. It is shown that the dynamics of the oscillatory flow are directly controlled by the thermal and hydrodynamic interactions brought about by combined natural and thermocapillary convection. Therefore, as numerical simulations show, there are considerable differences between the 1-g and low-g temperature and flow fields at both low and high Marangoni numbers. This has serious implications for both materials processing and fluid management in space.

Description: The Advanced Subsonic Combustion Rig (ASCR), a unique, state-of-the-art facility for conducting combustion research, is located at the NASA Lewis Research Center in Cleveland, Ohio. The ASCR, which was nearing completion at the close of 1995, will be capable of simulating the very high pressure and high temperature conditions that are expected to exist in future, advanced subsonic gas turbine (jet) engines. Future environmental regulations will require much cleaner burning (more environmentally friendly) aircraft engines. The ASCR is critical to the development of these cleaner engines. It will allow NASA and U.S. aircraft engine industry researchers to identify and test promising clean-burning gas turbine engine combustion concepts under the pressure and temperature conditions that are expected for those future engines. Combustion processes will be investigated for a variety of next-generation aircraft engine sizes, including engines for large, long-range aircraft (with typical trip lengths of about 3000 mi) and for regional aircraft (with typical trip lengths of about 400 mi). The ASCR design was conceived and initiated in 1993, and fabrication and construction of the rig, including the buildup of an advanced control room, took place throughout 1994 and 1995. In early 1996, the ASCR will be operational for obtaining research data. The ASCR is an intricate part of the NASA Advanced Subsonic Technology Propulsion Program, which is aimed at developing technologies critical to the next generation of gas turbine engines. This effort is in collaboration with the U.S. aircraft gas turbine engine industry. A goal of the Advanced Subsonic Technology Propulsion Program is to develop combustion concepts and technologies that will result in gas turbine engines that produce 50 percent less nitrous oxide (NO_x) pollutants than current engines do. This facility is unique in its capability to simulate advanced subsonic engine pressure, temperature, and air flow rate conditions. Specifically, it will provide operating temperatures up to 3000 F and pressures up to 60 atm. Under these conditions, researchers will obtain detailed combustion temperatures, pressures, and flow velocities as well as the chemical compositions of the combustion exhaust. Researchers also will be able to obtain data by using nonintrusive laser diagnostic techniques. The ASCR facility will be used to test fundamental combustion configurations (flametubes) for detailed study of combustion processes, to test sectors of gas turbine combustors to study the process in configurations more like those of aircraft engines, and in some cases to test full annular combustors.

Description: The success of any solution methodology for studying gas-turbine combustor flows depends a great deal on how well it can model various complex, rate-controlling processes associated with turbulent transport, mixing, chemical kinetics, evaporation and spreading rates of the spray, convective and radiative heat transfer, and other phenomena. These phenomena often strongly interact with each other at disparate time and length scales. In particular, turbulence plays an important role in determining the rates of mass and heat transfer, chemical reactions, and evaporation in many practical combustion devices. Turbulence manifests its influence in a diffusion flame in several forms depending on how turbulence interacts with various flame scales. These forms range from the so-called wrinkled, or stretched, flamelets regime, to the distributed combustion regime. Conventional turbulence closure models have difficulty in treating highly nonlinear reaction rates. A solution procedure based on the joint composition probability density function (PDF) approach holds the promise of modeling various important combustion phenomena relevant to practical combustion devices such as extinction, blowoff limits, and emissions predictions because it can handle the nonlinear chemical reaction rates without any approximation. In this approach, mean and turbulence gas-phase velocity fields are determined from a standard turbulence model; the joint composition field of species and enthalpy are determined from the solution of a modeled PDF transport equation; and a Lagrangian-based dilute spray model is used for the liquid-phase representation with appropriate consideration of the exchanges of mass, momentum, and energy between the two phases. The PDF transport equation is solved by a Monte Carlo method, and existing state-of-the-art numerical representations are used to solve the mean gasphase velocity and turbulence fields together with the liquid-phase equations. The joint composition PDF approach was extended in our previous work to the study of compressible reacting flows. The application of this method to several supersonic diffusion flames associated with scramjet combustor flow fields provided favorable comparisons with the available experimental data. A further extension of this approach to spray flames, three-dimensional computations, and parallel computing was reported in a recent paper. The recently developed PDF/SPRAY/computational fluid dynamics (CFD) module combines the novelty of the joint composition PDF approach with the ability to run on parallel architectures. This algorithm was implemented on the NASA Lewis Research Center's Cray T3D, a massively parallel computer with an aggregate of 64 processor elements. The calculation procedure was applied to predict the flow properties of both open and confined swirl-stabilized spray flames.

Description: The traditional approach to the design of engine inlets for commercial transport aircraft is a tedious process that ends with a less-than-optimum design. With the advent of high-speed computers and the availability of more accurate and reliable computational fluid dynamics (CFD) solvers, numerical optimization processes can effectively be used to design an aerodynamic inlet lip that enhances engine performance. The designers' experience at Boeing Corporation showed that for a peak Mach number on the inlet surface beyond some upper limit, the performance of the engine degrades excessively. Thus, our objective was to optimize efficiency (minimize the peak Mach number) at maximum cruise without compromising performance at other operating conditions. Using a CFD code NPARC, the NASA Lewis Research Center, in collaboration with Boeing, developed an integrated procedure at Lewis to find the optimum shape of a subsonic inlet lip and a numerical optimization code, ADS. We used a GRAPE-based three-dimensional grid generator to help automate the optimization procedure. The inlet lip shape at the crown and the keel was described as a superellipse, and the superellipse exponents and radii ratios were considered as design variables. Three operating conditions: cruise, takeoff, and rolling takeoff, were considered in this study. Three-dimensional Euler computations were carried out to obtain the flow field. At the initial design, the peak Mach numbers for maximum cruise, takeoff, and rolling takeoff conditions were 0.88, 1.772, and 1.61, respectively. The acceptable upper limits on the takeoff and rolling takeoff Mach numbers were 1.55 and 1.45. Since the initial design provided by Boeing was found to be optimum with respect to the maximum cruise condition, the sum of the peak Mach numbers at takeoff and rolling takeoff were minimized in the current study while the maximum cruise Mach number was constrained to be close to that at the existing design. With this objective, the optimum design satisfied the upper limits at takeoff and rolling takeoff while retaining the desirable cruise performance. Further studies are being conducted to include static and cross-wind operating conditions in the design optimization procedure. This work was carried out in collaboration with Dr. E.S. Reddy of NYMA, Inc.

Description: The WIND code is a general-purpose, structured, multizone, compressible flow solver that can be used to analyze steady or unsteady flow for a wide range of geometric configurations and over a wide range of flow conditions. WIND is the latest product of the NPARC Alliance, a formal partnership between the NASA Lewis Research Center and the Air Force Arnold Engineering Development Center (AEDC). WIND Version 1.0 was released in February 1998, and Version 2.0 will be released in February 1999. The WIND code represents a merger of the capabilities of three existing computational fluid dynamics codes--NPARC (the original NPARC Alliance flow solver), NXAIR (an Air Force code used primarily for unsteady store separation problems), and NASTD (the primary flow solver at McDonnell Douglas, now part of Boeing).

Description: Results from a recently completed enhanced mixing program are summarized in the two technical papers. These studies were parts of a High Speed Research (HSR)-supported joint Government/industry/university program that involved, in addition to the NASA Lewis Research Center, researchers at United Technologies Research Center, Allison Engine Company, CFD Research Corporation, and the University of California, Irvine. The studies investigated the mixing of jets injected normal to a confined subsonic mainsteam in both rectangular and cylindrical ducts. Experimental and computational studies were performed in both nonreacting and reacting flows. The orifice geometries and flow conditions were selected as typical of the complex three-dimensional flows in the combustion chambers in low-emission gas turbine engines. The principal conclusion from both the experiments and modeling was that the momentum-flux ratio J and orifice spacing S/H were the most significant flow and geometry variables, respectively. Conserved scalar distributions were similar-independent of reaction, orifice diameter H/d, and shape-when the orifice spacing and the square root of the momentum-flux ratio were inversely proportional. Jet penetration was critical, and penetration decreased as either momentum-flux ratio or orifice spacing decreased. We found that planar averages must be considered in context with the distributions. The mass-flow ratios and the orifices investigated were often very large. The jet-to-mainstream mass-flow ratio was varied from significantly less than 1 to greater than 1. The orifice-area to mainstream-cross-sectional-area was varied from approx. 0 to 0.5, and the axial planes of interest were often just downstream of the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations. As an example of the results, the accompanying figure shows the effects of different rates of mass addition on the opposite walls of a rectangular duct.

Description: Experimental results of tests made at the Langley Memorial Aeronautical Laboratory are presented to show how heat-transfer coefficients can he increased by a method utilizing the high rate of heat transfer known to exist on any heat-transfer surface in the region adjacent to the edge on which the cooling or heating fluid impinges. The results show that, for the same pressure drop, the average surface heat-transfer.coefficient can be increased 50 to 100 percent when a cooling surface having a length of four inches in the direction of fluid flow is cut to form twenty fins with a length of 0.2 inch in the direction of fluid flow and the fins are sharpened and staggered in the air stream. The percentage of increase in the surface heat-transfer coefficient obtained as a result of shortening the length of the cooling surface varies with the pressure drop of the cooling fluid in passing the surface, the increase being largest when small pressure drop is used and smallest when high pressure drop is used.

Description: An analysis is made to simplify pressure-drop calculations for nonadiabatic and adiabatic friction flow of air in smooth cylindrical tubes when the density changes due to heat transfer and pressure drop are appreciable. Solutions of the equation of motion are obtained by the use of Reynolds' analogy between heat transfer and skin friction. Charts of the solutions are presented for making pressure-drop calculations. A technique of using the charts to determine the position of a normal shock in a tube is described.

Description: About half of all scientific and commercial spacecraft spin during some part of their mission. Although spinning has many benefits (increasing stability, controlling the location of liquid propellants, and distributing solar heat loads), it also creates problems because a precession (or wobble) motion is unavoidable. For modern spacecraft, by far the largest source of wobble is energy dissipation in the liquid of partially filled tanks. The liquid s energy dissipation cannot, however, be quantified adequately by any ground test. Current analytical models are also inadequate because fundamental data on fluid motion in low gravity are needed to validate them. Consequently, spacecraft attitude-control systems are designed and operated very conservatively. Nonetheless, spacecraft often still perform poorly in orbit, and some have been lost because of a rapid increase of the wobble rate. The Liquid Motion Experiment (LME) was designed to provide spacecraft designers accurate data on the wobble dynamics of spacecraft that contain large quantities of mobile liquids. LME, which was flown on the space shuttle mission STS-84, was built under contract to the NASA Lewis Research Center by the Southwest Research Institute of San Antonio, Texas. Major accomplishments for 1998 include reduction of the flight data and publication of the experimental results. LME was essentially a spin table that created a realistic nutation motion of scale-model tanks containing liquid. Two spherical and two cylindrical transparent tanks were tested simultaneously, and three sets of such tanks were employed to vary liquid viscosity, fill level, and propellant management device (PMD) design. All the tanks were approximately 4.5 in. in diameter. The primary test measurements were the radial and tangential torques exerted on the tanks by the liquid. These torques could not be measured on the ground because of the masking effects of gravity.

Description: For the last several years, LeRC-HT, a three-dimensional computational fluid dynamics (CFD) computer code for analyzing gas turbine flow and convective heat transfer, has been evolving at the NASA Lewis Research Center. The code is unique in its ability to give a highly detailed representation of the flow field very close to solid surfaces. This is necessary for an accurate representation of fluid heat transfer and viscous shear stresses. The code has been used extensively for both internal cooling passage flows and hot gas path flows--including detailed film cooling calculations, complex tip-clearance gap flows, and heat transfer. In its current form, this code has a multiblock grid capability and has been validated for a number of turbine configurations. The code has been developed and used primarily as a research tool (at least 35 technical papers have been published relative to the code and its application), but it should be useful for detailed design analysis. We now plan to make this code available to selected users for further evaluation.

Description: The overall performance of a centrifugal compressor depends on the performance of the impeller and diffuser as well as on the interactions occurring between these components. Accurate measurements of the flow fields in each component are needed to develop computational models that can be used in compressor design codes. These measurements must be made simultaneously over an area that covers both components so that researchers can understand the interactions occurring between the two components. Optical measurement techniques are being used at the NASA Lewis Research Center to measure the velocity fields present in both the impeller and diffuser of a 4:1 pressure ratio centrifugal compressor operating at several conditions ranging from design flow to surge. Laser Doppler Velocimetry (LDV) was used to measure the intrablade flows present in the impeller, and the results were compared with analyses obtained from two three-dimensional viscous codes. The development of a region of low throughflow velocity fluid within this high-speed impeller was examined and compared with a similar region first observed in a large low-speed centrifugal impeller at Lewis. Particle Image Velocimetry (PIV) is a relatively new technique that has been applied to measuring the diffuser flow fields. PIV can collect data rapidly in the diffuser while avoiding the light-reflection problems that are often encountered when LDV is used. The Particle Image Velocimeter employs a sheet of pulsed laser light that is introduced into the diffuser in a quasi-radial direction through an optical probe inserted near the diffuser discharge. The light sheet is positioned such that its centerline is parallel to the hub and shroud surfaces and such that it is parallel to the diffuser vane, thereby avoiding reflections from the solid surfaces. Seed particles small enough to follow the diffuser flow are introduced into the compressor at an upstream location. A high-speed charge-coupled discharge (CCD) camera is synchronized to the laser pulse rate; this allows it to capture images of seed particle position that are separated by a small increment in time. A crosscorrelation of a particle's position in two consecutive images provides an estimate of flow velocity and direction. Multiple image pairs obtained in rapid succession at a particular flow condition provide enough measurements for statistical significance. PIV provides simultaneous velocity measurements over the entire plane that is illuminated by the light sheet instead of at a single point, as is the case when LDV is used. PIV has a further advantage in that the laser light pulse can be triggered by an external source such as a high-response pressure transducer. This feature will allow PIV to synchronize flow imaging to physical phenomena such as rotating stall or stall precursor waves. We hope that this technique can be used to obtain images of the flow field during and just prior to stall.

Description: Recently, refractive secondary solar concentrator systems were developed for solar thermal power and propulsion (ref. 1). Single-crystal oxides-such as yttria-stabilized zirconia (Y2O3-ZrO2), yttrium aluminum garnet (Y3Al5O12, or YAG), magnesium oxide (MgO), and sapphire (Al2O3)-are candidate refractive secondary concentrator materials. However, the refractive concentrator system will experience high-temperature thermal cycling in the solar thermal engine during the sun/shade transition of a space mission. The thermal mechanical reliability of these components in severe thermal environments is of great concern. Simulated mission tests are important for evaluating these candidate oxide materials under a variety of transient and steady-state heat flux conditions. In this research at the NASA Lewis Research Center, a controlled heat flux test approach was developed for investigating the thermal mechanical stability of the candidate oxide. This approach used a 3.0-kW continuous-wave (wavelength, 10.6 mm) carbon dioxide (CO2) laser (ref. 2). The CO2 laser is especially well-suited for single-crystal thermal shock tests because it can directly deliver well-characterized heat energy to the oxide surfaces. Since the oxides are opaque at the 10.6-mm wavelength of the laser beam, the light energy is absorbed at the surfaces rather than transmitting into the crystals, and thus generates the required temperature gradients within the specimens. The following figure is a schematic diagram of the test rig.

Description: A new turbomachinery analysis code called SWIFT has been developed at the NASA Lewis Research Center. SWIFT solves thin-layer Navier-Stokes equations with the Baldwin-Lomax turbulence model and an explicit finite-difference scheme. Preconditioning allows the code to be used for all speed ranges, from incompressible to supersonic flows. Multiblock capability allows three types of grids to be patched together to simulate many types of turbomachinery geometries, including hub and tip clearances. In addition, code users can analyze multistage turbomachinery by using a steady averaging-plane approach. This approach uses linearized characteristic boundary conditions to pass information accurately between the stages.

Description: Capillary-pumped loops (CPL's) are devices that are used to transport heat from one location to another--specifically to transfer heat away from something. In low-gravity applications, such as satellites (and possibly the International Space Station), CPL's are used to transfer heat from electrical devices to space radiators. This is accomplished by evaporating one liquid surface on the hot side of the CPL and condensing the vapor produced onto another liquid surface on the cold side. Capillary action, the phenomenon that causes paper towels to absorb spilled liquids, is used to "pump" the liquid back to the evaporating liquid surface (hot side) to complete the "loop." CPL's require no power to operate and can transfer heat over distances as large as 30 ft or more. Their reliance upon evaporation and condensation to transfer heat makes them much more economical in terms of weight than conventional heat transfer systems. Unfortunately, they have proven to be unreliable in space operations, and the explanation for this unreliability has been elusive. The Capillary-Driven Heat Transfer (CHT) experiment is investigating the fundamental fluid physics phenomena thought to be responsible for the failure of CPL's in low-gravity operations. If the failure mechanism can be identified, then appropriate design modifications can be developed to make capillary phase-change heat-transport devices a more viable option in space applications. CHT was conducted onboard the Space Shuttle Columbia during the first Microgravity Science Laboratory (MSL-1) mission, STS-94, which flew from July 1 to 17, 1997. The CHT glovebox investigation, which was conceived by Dr. Kevin Hallinan and Jeffrey Allen of the University of Dayton, focused on studying the dynamics associated with the heating and cooling at the evaporating meniscus within a capillary phase-change device in a low-gravity environment. The CHT experimental hardware was designed by a small team of engineers from Aerospace Design & Fabrication (ADF), the NASA Lewis Research Center, and the University of Dayton. The hardware consisted of two experiment modules that each contained an instrumented test loop (idealized capillary-pumped loop), a base unit for power conversion and backlighting, a display unit with 15 LED's (light-emitting diodes) to display temperatures, pressure, heater power, and time, a control unit to select heaters and heater settings, a cooling fan, and associated cables.

Description: Applying smart materials in aeropropulsion systems may improve the performance of aircraft engines through a variety of vibration, noise, and shape-control applications. To facilitate the experimental characterization of these smart structures, researchers have been focusing on developing analytical models to account for the coupled mechanical, electrical, and thermal response of these materials. One focus of current research efforts has been directed toward incorporating a comprehensive thermal analysis modeling capability. Typically, temperature affects the behavior of smart materials by three distinct mechanisms: Induction of thermal strains because of coefficient of thermal expansion mismatch 1. Pyroelectric effects on the piezoelectric elements; 2. Temperature-dependent changes in material properties; and 3. Previous analytical models only investigated the first two thermal effects mechanisms. However, since the material properties of piezoelectric materials generally vary greatly with temperature (see the graph), incorporating temperature-dependent material properties will significantly affect the structural deflections, sensory voltages, and stresses. Thus, the current analytical model captures thermal effects arising from all three mechanisms through thermopiezoelectric constitutive equations. These constitutive equations were incorporated into a layerwise laminate theory with the inherent capability to model both the active and sensory response of smart structures in thermal environments. Corresponding finite element equations were formulated and implemented for both the beam and plate elements to provide a comprehensive thermal effects modeling capability.

Description: The Surface Tension Driven Convection Experiment (STDCE) was designed to study basic fluid mechanics and heat transfer on thermocapillary flows generated by temperature variations along the free surfaces of liquids in microgravity. STDCE first flew on the USML-1 mission in July 1992 and was rebuilt for the USML-2 mission that was launched in October 1995. This was a collaborative project with principal investigators from Case Western Reserve University (CWRU), Professors Simon Ostrach and Yasuhiro Kamotani, along with a team from the NASA Lewis Research Center composed of civil servants and contractors from Aerospace Design & Fabrication, Inc. (ADF), Analex, and NYMA, Inc.

Description: In our world-class, optically accessible combustion facility at the NASA Lewis Research Center, we have developed the unique capability of making three-dimensional fuel distribution measurements of aviation gas turbine fuel injectors at actual operating conditions. These measurements are made in situ at the actual operating temperatures and pressures using the JP-grade fuels of candidate next-generation advanced aircraft engines for the High Speed Research (HSR) and Advanced Subsonics Technology (AST) programs. The inlet temperature and pressure ranges used thus far are 300 to 1100 F and 80 to 250 psia. With these data, we can obtain the injector spray angles, the fuel mass distributions of liquid and vapor, the degree of fuel vaporization, and the degree to which fuel has been consumed. The data have been used to diagnose the performance of injectors designed both in-house and by major U.S. engine manufacturers and to design new fuel injectors with overall engine performance goals of increased efficiency and reduced environmental impact. Mie scattering is used to visualize the liquid fuel, and laser-induced fluorescence is used to visualize both liquid and fuel vapor.

Description: Sprays occur in a wide variety of industrial and power applications and in the processing of materials. A liquid spray is a phase flow with a gas as the continuous phase and a liquid as the dispersed phase (in the form of droplets or ligaments). Interactions between the two phases, which are coupled through exchanges of mass, momentum, and energy, can occur in different ways at different times and locations involving various thermal, mass, and fluid dynamic factors. An understanding of the flow, combustion, and thermal properties of a rapidly vaporizing spray requires careful modeling of the rate-controlling processes associated with the spray's turbulent transport, mixing, chemical kinetics, evaporation, and spreading rates, as well as other phenomena. In an attempt to advance the state-of-the-art in multidimensional numerical methods, we at the NASA Lewis Research Center extended our previous work on sprays to unstructured grids and parallel computing. LSPRAY, which was developed by M.S. Raju of Nyma, Inc., is designed to be massively parallel and could easily be coupled with any existing gas-phase flow and/or Monte Carlo probability density function (PDF) solver. The LSPRAY solver accommodates the use of an unstructured mesh with mixed triangular, quadrilateral, and/or tetrahedral elements in the gas-phase solvers. It is used specifically for fuel sprays within gas turbine combustors, but it has many other uses. The spray model used in LSPRAY provided favorable results when applied to stratified-charge rotary combustion (Wankel) engines and several other confined and unconfined spray flames. The source code will be available with the National Combustion Code (NCC) as a complete package.

Description: The Interface Configuration Experiment (ICE) is part of a multifaceted study that is exploring the often striking behavior of liquid-vapor interfaces in low-gravity environments. Although the experiment was posed largely as a test of current mathematical theory, applications of the results should be manifold. In space almost every fluid system is affected, if not dominated, by capillarity (the effects of surface tension). As a result, knowledge of fluid interface behavior, in particular an equilibrium interface shape from which any analysis must begin, is fundamental--from the control of liquid fuels and oxygen in storage tanks to the design and development of inspace thermal systems, such as heat pipes and capillary pumped loops. ICE has increased, and should continue to increase, such knowledge as it probes the specific peculiarities of current theory upon which our present understanding rests. Several versions of ICE have been conducted in the drop towers at the NASA Lewis Research Center, on the space shuttles during the first and second United States Microgravity Laboratory missions (USML-1 and USML-2), and most recently aboard the Russian Mir space station. These studies focused on interfacial problems concerning the existence, uniqueness, configuration, stability, and flow characteristics of liquid-vapor interfaces. Results to date have clearly demonstrated the value of the present theory and the extent to which it can predict the behavior of capillary systems.

Description: Particle Imaging Velocimetry (PIV) is an optical technique whereby a pulsed laser sheet is used to illuminate particles entrained in a fluid across an extended planar cross section of a flow field. Electronic recording of the particle positions at two closely timed laser pulses permits the computation of the flow velocity. PIV captures the instantaneous flow field, permitting the study of unsteady flow phenomena. Mean flow statistics can be computed by acquiring several hundred images and averaging the results. The first-ever successful application of PIV to acquire measurements in a high-speed rotating turbomachinery blade row was completed in NASA Lewis Research Center's W-8 Single Stage Axial Compressor Facility. Measurements were acquired in a 20-in.-diameter transonic compressor rotor operating at 17,188 rpm. A custom-designed light-sheet generating probe was used to insert the high-energy, pulsed light-sheet illumination required for recording the unblurred images of particles entrained in the fluid. Measurements of the shock wave formed within the rotor blade passage and of unsteady structures within the blade wakes were acquired. These measurements provide insight into unsteady spatial structures in the flow field which cannot be measured with the more conventional laser anemometry technique. The PIV technique provides both instantaneous and average velocity data in a transonic compressor in an order of magnitude less time than required for other conventional optical diagnostic techniques.

Description: Ceramic thermal barrier coatings have received increasing attention for advanced gas turbine and diesel engine applications because of their ability to provide thermal insulation to engine components. However, the durability of these coatings under the severe thermal cycling conditions encountered in a diesel engine (ref. 1) still remains a major issue. In this research at the NASA Lewis Research Center, a high-power laser was used to investigate the thermal fatigue behavior of a yttria-stabilized zirconia coating system under simulated diesel engine conditions. The mechanisms of fatigue crack initiation and propagation, and of coating failure under complex thermal low-cycle fatigue (LCF, representing stop/start cycles) and thermal high-cycle fatigue (HCF, representing operation at 1300 rpm) are described. Continuous wave and pulse laser modes were used to simulate pure LCF and combined LCF/HCF, respectively (ref. 2). The LCF mechanism was found to be closely related to the coating sintering and creep at high temperatures. These creep strains in the ceramic coating led to a tensile stress state during cooling, thus providing the major driving force for crack growth under LCF conditions. The combined LCF/HCF tests induced more severe coating surface cracking, microspallation, and accelerated crack growth than did the pure LCF test. HCF thermal loads also facilitated lateral crack branching and ceramic/bond coat interface delaminations. HCF is associated with the cyclic stresses originating from the high-frequency temperature fluctuation at the ceramic coating surface. The HCF thermal loads act on the crack by a wedging mechanism (ref. 1), resulting in continuous crack growth at temperature. The HCF stress intensity factor amplitude increases with the interaction depth and temperature swing, and decreases with the crack depth. HCF damage also increases with the thermal expansion coefficient and the Young's modulus of the ceramic coating (refs. 1 and 3).

Description: The number of optical techniques that may potentially be used during a given wind tunnel test is continually growing. These include parameter sensitive paints that are sensitive to temperature or pressure, several different types of off-body and on-body flow visualization techniques, optical angle-of-attack (AoA), optical measurement of model deformation, optical techniques for determining density or velocity, and spectroscopic techniques for determining various flow field parameters. Often in the past the various optical techniques were developed independently of each other, with little or no consideration for other techniques that might also be used during a given test. Recently two optical techniques have been increasingly requested for production measurements in NASA wind tunnels. These are the video photogrammetric (or videogrammetric) technique for measuring model deformation known as the video model deformation (VMD) technique, and the parameter sensitive paints for making global pressure and temperature measurements. Considerations for, and initial attempts at, simultaneous measurements with the pressure sensitive paint (PSP) and the videogrammetric techniques have been implemented. Temperature sensitive paint (TSP) has been found to be useful for boundary-layer transition detection since turbulent boundary layers convect heat at higher rates than laminar boundary layers of comparable thickness. Transition is marked by a characteristic surface temperature change wherever there is a difference between model and flow temperatures. Recently, additional capabilities have been implemented in the target-tracking videogrammetric measurement system. These capabilities have permitted practical simultaneous measurements using parameter sensitive paint and video model deformation measurements that led to the first successful unified test with TSP for transition detection in a large production wind tunnel.

Description: A study of a flow through a straight converging-diverging nozzle of simple design has been made preliminary to studies of other supersonic flows. The diverging part of the nozzle was designed by the Prandtl-Busemann method to give a uniform pressure at its exit of 0.298 times the initial total head, that is, to give a Mach number of 1.436. Schlieren photographs of the flow and pressure-distribution measurements along the diverging part of the nozzle were made. A comparison of the theory with these measurements is presented.

Description: Microscopic spheres suspended in liquid become highly ordered under the proper conditions. Such collections of particles, called colloidal suspensions or colloids, are the subject of a series of ongoing microgravity experiments at the NASA Lewis Research Center. By studying the way these colloidal suspensions order themselves, scientists can better understand how atoms of a liquid become ordered to form a solid. In addition, highly ordered colloids have special properties that may make them useful in future hightech applications. Work is underway at Lewis to develop an optical microscope to view these colloidal suspensions sphere by sphere in microgravity.

Description: Under the Advanced Subsonic Technology (AST) Program, an aeroelastic analysis code (TURBO-AE) based on Navier-Stokes equations is currently under development at NASA Lewis Research Center s Machine Dynamics Branch. For a blade row, aeroelastic instability can occur in any of the possible interblade phase angles (IBPA s). Analyzing small IBPA s is very computationally expensive because a large number of blade passages must be simulated. To reduce the computational cost of these analyses, we used time shifted, or phase-lagged, boundary conditions in the TURBO-AE code. These conditions can be used to reduce the computational domain to a single blade passage by requiring the boundary conditions across the passage to be lagged depending on the IBPA being analyzed. The time-shifted boundary conditions currently implemented are based on the direct-store method. This method requires large amounts of data to be stored over a period of the oscillation cycle. On CRAY computers this is not a major problem because solid-state devices can be used for fast input and output to read and write the data onto a disk instead of storing it in core memory.

Description: The NASA Lewis Research Center is managing a task to numerically simulate overnight, on a parallel computing testbed, the aerodynamic flow in the complete low-pressure subsystem (LPS) of a gas turbine engine. The model solves the three-dimensional Navier- Stokes flow equations through all the components within the LPS, as well as the external flow around the engine nacelle. The LPS modeling task is being performed by Allison Engine Company under the Small Engine Technology contract. The large computer simulation was evaluated on networked computer systems using 8, 16, and 32 processors, with the parallel computing efficiency reaching 75 percent when 16 processors were used.

Description: One of the most challenging areas of research in aerodynamics is the reduction of skin friction, especially for turbulent flow. Reduced skin friction means less drag. For aircraft, less drag can lead to less fuel burned or to a greater flight range for a fixed amount of fuel. Many techniques and methods have been tried; however, none of them has significantly reduced skin friction in the flight environment. An innovative skin-friction reduction technique, the Microblowing Technique (MBT), was invented in 1993. This is a unique concept in which an extremely small amount of air is blown vertically at a surface through very small holes. It can be used for aircraft or marine vehicles, such as submarines (where water is blown through the holes instead of air). As shown in the figure, the outer layer, which controls vertical flow, is a plate with high-vertical holes. The inner layer, which produces evenly distributed flow, is a low-permeability porous plate. Microblowing reduces the surface roughness and changes the flow velocity profile on the surface, thereby reducing skin friction.

Description: Renewed interest in hypersonic propulsion systems has led to research programs investigating combined cycle engines that are designed to operate efficiently across the flight regime. The Rocket-Based Combined Cycle Engine is a propulsion system under development at the NASA Lewis Research Center. This engine integrates a high specific impulse, low thrust-to-weight, airbreathing engine with a low-impulse, high thrust-to-weight rocket. From takeoff to Mach 2.5, the engine operates as an air-augmented rocket. At Mach 2.5, the engine becomes a dual-mode ramjet; and beyond Mach 8, the rocket is turned back on. One Rocket-Based Combined Cycle Engine variation known as the "Strut-Jet" concept is being investigated jointly by NASA Lewis, the U.S. Air Force, Gencorp Aerojet, General Applied Science Labs (GASL), and Lockheed Martin Corporation. Work thus far has included wind tunnel experiments and computational fluid dynamics (CFD) investigations with the NPARC code. The CFD method was initiated by modeling the geometry of the Strut-Jet with the GRIDGEN structured grid generator. Grids representing a subscale inlet model and the full-scale demonstrator geometry were constructed. These grids modeled one-half of the symmetric inlet flow path, including the precompression plate, diverter, center duct, side duct, and combustor. After the grid generation, full Navier-Stokes flow simulations were conducted with the NPARC Navier-Stokes code. The Chien low-Reynolds-number k-e turbulence model was employed to simulate the high-speed turbulent flow. Finally, the CFD solutions were postprocessed with a Fortran code. This code provided wall static pressure distributions, pitot pressure distributions, mass flow rates, and internal drag. These results were compared with experimental data from a subscale inlet test for code validation; then they were used to help evaluate the demonstrator engine net thrust.

Description: The National PARC (NPARC) Alliance was established by the NASA Lewis Research Center and the Air Force Arnold Engineering Development Center to provide the U.S. aeropropulsion community with a reliable Navier-Stokes code for simulating the nonrotating components of propulsion systems. Recent improvements to the turbulence model capabilities of the NPARC code have significantly improved its capability to simulate turbulent flows. Specifically, the Chien k-epsilon and Wilcox k-omega turbulence models were implemented at Lewis. Lewis researchers installed the Chien k-epsilon model into NPARC to improve the code's ability to calculate turbulent flows with attached wall boundary layers and free shear layers. Calculations with NPARC have demonstrated that the Chien k-epsilon model provides more accurate calculations than those obtained with algebraic models previously available in the code. Grid sensitivity investigations have shown that computational grids must be packed against the solid walls such that the first point off of the wall is placed in the laminar sublayer. In addition, matching the boundary layer and momentum thicknesses entering mixing regions is necessary for an accurate prediction of the free shear-layer growth.

Description: In a continuing in-house program at the NASA Lewis Research Center, analytical and numerical methods are being developed to apply radiative analysis to predict transient temperature distributions and heat flows in partially transmitting materials. Results have been obtained for a single plane layer, and a transient analysis is being developed for a two-layer composite where each layer has a different refractive index. Because the ceramic refractive indices are larger than one, internal reflections are produced at the surfaces and at the internal interface. Reflections tend to distribute energy within a layer, and this affects the transient temperature distributions.

Description: A gas-jet diffusion flame is similar to the flame on a Bunsen burner, where a gaseous fuel (e.g., propane) flows from a nozzle into an oxygen-containing atmosphere (e.g., air). The difference is that a Bunsen burner allows for (partial) premixing of the fuel and the air, whereas a diffusion flame is not premixed and gets its oxygen (principally) by diffusion from the atmosphere around the flame. Simple gas-jet diffusion flames are often used for combustion studies because they embody the mechanisms operating in accidental fires and in practical combustion systems. However, most practical combustion is turbulent (i.e., with random flow vortices), which enhances the fuel/air mixing. These turbulent flames are not well understood because their random and transient nature complicates analysis. Normal gravity studies of turbulence in gas-jet diffusion flames can be impeded by buoyancy-induced instabilities. These gravitycaused instabilities, which are evident in the flickering of a candle flame in normal gravity, interfere with the study of turbulent gas-jet diffusion flames. By conducting experiments in microgravity, where buoyant instabilities are avoided, we at the NASA Lewis Research Center hope to improve our understanding of turbulent combustion. Ultimately, this could lead to improvements in combustor design, yielding higher efficiency and lower pollutant emissions. Gas-jet diffusion flames are often researched as model flames, because they embody mechanisms operating in both accidental fires and practical combustion systems (see the first figure). In normal gravity laboratory research, buoyant air flows, which are often negligible in practical situations, dominate the heat and mass transfer processes. Microgravity research studies, however, are not constrained by buoyant air flows, and new, unique information on the behavior of gas-jet diffusion flames has been obtained.

Description: The Interface Configuration Experiment (ICE) is actually a series of experiments that explore the striking behavior of liquid-vapor interfaces (i.e., fluid surfaces) in a low gravity environment under which major shifts in liquid position can arise from small changes in container shape or contact angle. Although these experiments are designed to test current mathematical theory, there are numerous practical applications that could result from these studies. When designing fluid management systems for space-based operations, it is important to be able to predict the locations and configurations that fluids will assume in containers under low-gravity conditions. The increased ability to predict, and hence control, fluid interfaces is vital to systems and/or processes where capillary forces play a significant role both in space and on the Earth. Some of these applications are in general coating processes (paints, pesticides, printing, etc.), fluid transport in porous media (ground water flows, oil recovery, etc.), liquid propellant systems in space (liquid fuel and oxygen), capillary-pumped loops and heat pipes, and space-based life-support systems. In space, almost every fluid system is affected, if not dominated, by capillarity. Knowledge of the liquid-vapor interface behavior, and in particular the interface shape from which any analysis must begin, is required as a foundation to predict how these fluids will react in microgravity and on Earth. With such knowledge, system designs can be optimized, thereby decreasing costs and complexity, while increasing performance and reliability. ICE has increased, and will continue to increase this knowledge, as it probes the specific peculiarities of current theory upon which our current understanding of these effects is based. Several versions of ICE were conducted in NASA Lewis Research Center's drop towers and on the space shuttle during the first and second United States Microgravity Laboratory missions (USML-1 and USML-2). Additional tests are planned for the space shuttle and for the Russian Mir space station. These studies will focus on interfacial problems concerning surface existence, uniqueness, configuration, stability, and flow characteristics.

Description: The development of thermodynamics and statistical mechanics is very important in the history of physics, and it underlines the difficulty in dealing with systems involving many bodies, even if those bodies are identical. Macroscopic systems of atoms typically contain so many particles that it would be virtually impossible to follow the behavior of all of the particles involved. Therefore, the behavior of a complete system can only be described or predicted in statistical ways. Under a grant to the NASA Lewis Research Center, scientists at the Case Western Reserve University have been examining the use of modern computing techniques that may be able to investigate and find the behavior of complete systems that have a large number of particles by tracking each particle individually. This is the study of molecular dynamics. In contrast to Monte Carlo techniques, which incorporate uncertainty from the outset, molecular dynamics calculations are fully deterministic. Although it is still impossible to track, even on high-speed computers, each particle in a system of a trillion trillion particles, it has been found that such systems can be well simulated by calculating the trajectories of a few thousand particles. Modern computers and efficient computing strategies have been used to calculate the behavior of a few physical systems and are now being employed to study important problems such as supersonic flows in the laboratory and in space. In particular, an animated video (available in mpeg format--4.4 MB) was produced by Dr. M.J. Woo, now a National Research Council fellow at Lewis, and the G-VIS laboratory at Lewis. This video shows the behavior of supersonic shocks produced by pistons in enclosed cylinders by following exactly the behavior of thousands of particles. The major assumptions made were that the particles involved were hard spheres and that all collisions with the walls and with other particles were fully elastic. The animated video was voted one of two winning videos in a competition held at the meeting of the American Physical Society's Division of Fluid Dynamics, held in Atlanta, Georgia, in November 1994. Of great interest was the result that in every shock there were a few high-speed precursor particles racing ahead of the shock, carrying information about its impending arrival. Most recently, Dr. Woo has been applying molecular dynamics techniques to the problem of determining the drag produced by the space station truss structure as it flies through the thin residual atmosphere of low-Earth orbit. This problem is made difficult by the complex structure of the truss and by the extreme supersonic nature of the flow. A fully filled section of the truss has already been examined, and drag predictions have been made. Molecular dynamics techniques promise to make realistic drag calculations possible even for very complex partially filled truss segments flying at arbitrary angles.

Description: No abstract available

Description: A complete "geometry to drag-polar" analysis capability for three-dimensional high-lift configurations is described. The approach is based on the use of unstructured meshes in order to enable rapid turnaround for complicated geometries which arise in high-lift con gurations. Special attention is devoted to creating a capability for enabling analyses on highly resolved grids. Unstructured meshes of several million vertices are initially generated on a work-station, and subsequently refined on a supercomputer. The flow is solved on these refined meshes on large parallel computers using an unstructured agglomeration multigrid algorithm. Good prediction of lift and drag throughout the range of incidences is demonstrated on a transport take-off configuration using up to 24.7 million grid points. The feasibility of using this approach in a production environment on existing parallel machines is demonstrated, as well as the scalability of the solver on machines using up to 1450 processors.

Description: The "CFD General Notation System" (CGNS) consists of a collection of conventions, and conforming software, for the storage and retrieval of Computational Fluid Dynamics (CFD) data. It facilitates the exchange of data between sites and applications, and helps stabilize the archiving of aerodynamic data. This effort was initiated in order to streamline the procedures in exchanging data and software between NASA and its customers, but the goal is to develop CGNS into a National Standard for the exchange of aerodynamic data. The CGNS development team is comprised of members from Boeing Commercial Airplane Group, NASA-Ames, NASA-Langley, NASA-Lewis, McDonnell-Douglas Corporation (now Boeing-St. Louis), Air Force-Wright Lab., and ICEM-CFD Engineering. The elements of CGNS address all activities associated with the storage of data on external media and its movement to and from application programs. These elements include: - The Advanced Data Format (ADF) Database manager, consisting of both a file format specification and its I/O software, which handles the actual reading and writing of data from and to external storage media; - The Standard Interface Data Structures (SIDS), which specify the intellectual content of CFD data and the conventions governing naming and terminology; - The SIDS-to-ADF File Mapping conventions, which specify the exact location where the CFD data defined by the SIDS is to be stored within the ADF file(s); and - The CGNS Mid-level Library, which provides CFD-knowledgeable routines suitable for direct installation into application codes. The CGNS Mid-level Library was designed to ease the implementation of CGNS by providing developers with a collection of handy I/O functions. Since knowledge of the ADF core is not required to use this library, it will greatly facilitate the task of interfacing with CGNS. There are currently 48 user callable functions that comprise the Mid-level library and are described in the Users Guide. The library is written in C, but each function has a FORTRAN counterpart.

Description: The flow fields of vortices, whether bouyancy-driven or lift-generated, are fascinating fluid-dynamic phenomena which often possess intense swirl velocities and complex time-dependent behavior. As part of the on-going study of vortex behavior, this paper presents a historical overview of the research conducted on the structure and modification of the vortices generated by the lifting surfaces of subsonic transport aircraft. It is pointed out that the characteristics of lift-generated vortices are related to the aerodynamic shapes that produce them and that various arrangements of surfaces can be used to produce different vortex structures. The primary purpose of the research to be described is to find a way to reduce the hazard potential of lift-generated vortices shed by subsonic transport aircraft in the vicinity of airports during landing and takeoff operations. It is stressed that lift-generated vortex wakes are so complex that progress towards a solution requires application of a combined theoretical and experimental research program because either alone often leads to incorrect conclusions. It is concluded that a satisfactory aerodynamic solution to the wake-vortex problem at airports has not yet been found but a reduction in the impact of the wake-vortex hazard on airport capacity may become available in the foreseeable future through wake-vortex avoidance concepts currently under study. The material to be presented in this overview is drawn from aerospace journals that are available publicly.

Description: The "CFD General Notation System" (CGNS) consists of a collection of conventions, and conforming software, for the storage and retrieval of Computational Fluid Dynamics (CFD) data. It facilitates the exchange of data between sites and applications, and helps stabilize the archiving of aerodynamic data. This effort was initiated in order to streamline the procedures in exchanging data and software between NASA and its customers, but the goal is to develop CGNS into a National Standard for the exchange of aerodynamic data. The CGNS development team is comprised of members from Boeing Commercial. Airplane Group, NASA-Ames, NASA-Langley, NASA-Lewis, McDonnell-Douglas Corporation (now Boeing-St. Louis), Air Force-Wright Lab., and ICEM-CFD Engineering. The elements of CGNS address all activities associated with the storage of data on external media and its movement to and from application programs. These elements include: 1) The Advanced Data Format (ADF) Database manager, consisting of both a file format specification and its 1/0 software, which handles the actual reading and writing of data from and to external storage media; 2) The Standard Interface Data Structures (SIDS), which specify the intellectual content of CFD data and the conventions governing naming and terminology; 3) The SIDS-to-ADF File Mapping conventions, which specify the exact location where the CFD data defined by the SIDS is to be stored within the ADF file(s); and 4) The CGNS Mid-level Library, which provides CFD-knowledgeable routines suitable for direct installation into application codes. The ADF is a generic database manager with minimal intrinsic capability. It was written for the purpose of storing large numerical datasets in an efficient, platform independent manner. To be effective, it must be used in conjunction with external agreements on how the data will be organized within the ADF database such defined by the SIDS. There are currently 34 user callable functions that comprise the ADF Core library and are described in the Users Guide. The library is written in C, but each function has a FORTRAN counterpart.

Description: Development of HPF versions of NPB and ARC3D showed that HPF has potential to be a high level language for parallelization of CFD applications. The use of HPF requires an intimate knowledge of the applications and a detailed analysis of data affinity, data movement and data granularity. Since HPF hides data movement from the user even with this knowledge it is easy to overlook pieces of the code causing low performance of the application. In order to simplify and accelerate the task of developing HPF versions of existing CFD applications we have designed and partially implemented ADAPT (Automatic Data Distribution and Placement Tool). The ADAPT analyzes a CFD application working on a single structured grid and generates HPF TEMPLATE, (RE)DISTRIBUTION, ALIGNMENT and INDEPENDENT directives. The directives can be generated on the nest level, subroutine level, application level or inter application level. ADAPT is designed to annotate existing CFD FORTRAN application performing computations on single or multiple grids. On each grid the application can considered as a sequence of operators each applied to a set of variables defined in a particular grid domain. The operators can be classified as implicit, having data dependences, and explicit, without data dependences. In order to parallelize an explicit operator it is sufficient to create a template for the domain of the operator, align arrays used in the operator with the template, distribute the template, and declare the loops over the distributed dimensions as INDEPENDENT. In order to parallelize an implicit operator, the distribution of the operator's domain should be consistent with the operator's dependences. Any dependence between sections distributed on different processors would preclude parallelization if compiler does not have an ability to pipeline computations. If a data distribution is "orthogonal" to the dependences of an implicit operator then the loop which implements the operator can be declared as INDEPENDENT.

Description: Experiments and computations were carried out on the adiabatic laminar boundary layer developing along the surfaces of a two-dimensional supersonic nozzle, consisting of upper and lower contoured nozzle blocks and flat sidewalls. Two- and three-dimensional Navier-Stokes codes, as well as two-dimensional boundary-layer codes were employed. These codes were adapted to the characteristics of a specific wind tunnel nozzle, so that their numerical results could be directly compared with experimental data obtained in the same nozzle. Such comparisons were made for the boundary-layer growth on the nozzle contoured surfaces, and for the boundary-layer growth, surface streamlines and surface shear on the sidewalls. The three-dimensional Navier-Stokes code was found to be the only one to correctly predict the mean laminar boundary-layer flow on both the sidewalls and the contoured surfaces. Theory and experiment both indicated that the sidewall flow is highly three-dimensional, with non-uniform shear, corner vortices and a boundary layer strongly distorted by cross flows induced by lateral pressure gradients.

Description: This summary presents the motivation for the Special Section on the credibility of computational fluid dynamics (CFD) simulations, its objective, its background and context, its content, and its major conclusions. Verification and validation (V&V) are the processes for establishing the credibility of CFD simulations. Validation assesses whether correct things are performed and verification assesses whether they are performed correctly. Various aspects of V&V are discussed. Progress is made in verification of simulation models. Considerable effort is still needed for developing a systematic validation method that can assess the credibility of simulated reality.

Description: An experimental and numerical investigation of the thermochemical state of arcjet flows is currently being conducted at NASA Ames Research Center. The experimental approach relies on the use of laser- and emission-spectroscopic diagnostic techniques in three regions of the flow. A fiber optic sensor is used to record spectrally resolved emission signals from the electrode package region, where the flow is most likely to be in thermochemical equilibrium. A second emission diagnostic measurement is made in the shock layer formed over a blunt-body test article placed in the stream, and a CCD camera is used to simultaneously record spectral emission from several measurement locations along the stagnation streamline. Downstream of the nozzle exit, but upstream of the test article, Laser-Induced Fluorescence (LIF) of atomic nitrogen is used to assess the nonequilibrium distribution of flow enthalpy in the free stream. Results from the measurements are compared with predictions from a two-temperature, axisymmetric flow model that solves the nozzle and shock-layer flows.

Description: Implicit methods have been the workhorse for the Euler and Navier-Stokes equations for the last 25 years. The ground breaking work of Dr. Joe Steger in implementing such techniques in practical Euler and Navier-Stokes codes provided the basis for all the success in this area. This presentation will highlight his contribution and technical excellence in the area of implicit methods for CFD.

Description: Workload characterization is used for modeling and evaluating of computing systems at different levels of detail. We present workload characterization for a class of Computational Fluid Dynamics (CFD) applications that solve Partial Differential Equations (PDEs). This workload characterization focuses on three high performance computing platforms: SGI Origin2000, EBM SP-2, a cluster of Intel Pentium Pro bases PCs. We execute extensive measurement-based experiments on these platforms to gather statistics of system resource usage, which results in workload characterization. Our workload characterization approach yields a coarse-grain resource utilization behavior that is being applied for performance modeling and evaluation of distributed high performance metacomputing systems. In addition, this study enhances our understanding of interactions between PDE solver workloads and high performance computing platforms and is useful for tuning these applications.

Description: The use of Direct numerical simulations (DNS) data in developing and testing turbulence models is reviewed. The data is used to test turbulence models at all levels: algebraic, one-equation, two-equation and full Reynolds stress models were tested. Particular examples on the development of models for the dissipation rate equation are presented. Homogeneous flows are used to test new scaling arguments for the various terms in the dissipation rate equation. The channel flow data is used to develop modifications to the equation model that take into account near-wall effects. DNS of compressible flows under mean compression are used in testing new compressible modifications to the two-equation models.