ALBERT

Description: Large-eddy simulation (LES) is conducted of a three-dimensional temporal mixing layer whose lower stream is initially laden with liquid drops which may evaporate during the simulation. The gas-phase equations are written in an Eulerian frame for two perfect gas species (carrier gas and vapour emanating from the drops), while the liquid-phase equations are written in a Lagrangian frame. The effect of drop evaporation on the gas phase is considered through mass, species, momentum and energy source terms. The drop evolution is modelled using physical drops, or using computational drops to represent the physical drops. Simulations are performed using various LES models previously assessed on a database obtained from direct numerical simulations (DNS). These LES models are for: (i) the subgrid-scale (SGS) fluxes and (ii) the filtered source terms (FSTs) based on computational drops. The LES, which are compared to filtered-and-coarsened (FC) DNS results at the coarser LES grid, are conducted with 64 times fewer grid points than the DNS, and up to 64 times fewer computational than physical drops. It is found that both constant-coefficient and dynamic Smagorinsky SGS-flux models, though numerically stable, are overly dissipative and damp generated small-resolved-scale (SRS) turbulent structures. Although the global growth and mixing predictions of LES using Smagorinsky models are in good agreement with the FC-DNS, the spatial distributions of the drops differ significantly. In contrast, the constant-coefficient scale-similarity model and the dynamic gradient model perform well in predicting most flow features, with the latter model having the advantage of not requiring a priori calibration of the model coefficient. The ability of the dynamic models to determine the model coefficient during LES is found to be essential since the constant-coefficient gradient model, although more accurate than the Smagorinsky model, is not consistently numerically stable despite using DNS-calibrated coefficients. With accurate SGS-flux models, namely scale-similarity and dynamic gradient, the FST model allows up to a 32-fold reduction in computational drops compared to the number of physical drops, without degradation of accuracy; a 64-fold reduction leads to a slight decrease in accuracy.

Description: New programs are forcing American propulsion system designers into unfamiliar territory. For instance, industry s answer to the cost and reliability goals set out by the Next Generation Launch Technology Program are engine concepts based on the Oxygen- Rich Staged Combustion Cycle. Historical injector design tools are not well suited for this new task. The empirical correlations do not apply directly to the injector concepts associated with the ORSC cycle. These legacy tools focus primarily on performance with environment evaluation a secondary objective. Additionally, the environmental capability of these tools is usually one-dimensional while the actual environments are at least two- and often three-dimensional. CFD has the potential to calculate performance and multi-dimensional environments but its use in the injector design process has been retarded by long solution turnaround times and insufficient demonstrated accuracy. This paper has documented the parallel paths of program support and technology development currently employed at Marshall Space Flight Center in an effort to move CFD to the forefront of injector design. MSFC has established a long-term goal for use of CFD for combustion devices design. The work on injector design is the heart of that vision and the Combustion Devices CFD Simulation Capability Roadmap that focuses the vision. The SRL concept, combining solution fidelity, robustness and accuracy, has been established as a quantitative gauge of current and desired capability. Three examples of current injector analysis for program support have been presented and discussed. These examples are used to establish the current capability at MSFC for these problems. Shortcomings identified from this experience are being used as inputs to the Roadmap process. The SRL evaluation identified lack of demonstrated solution accuracy as a major issue. Accordingly, the MSFC view of code validation and current MSFC-funded validation efforts were discussed in some detail. The objectives of each effort were noted. Issues relative to code validation for injector design were discussed in some detail. The requirement for CFD support during the design of the experiment was noted and discussed in terms of instrumentation placement and experimental rig uncertainty. In conclusion, MSFC has made significant progress in the last two years in advancing CFD toward the goal of application to injector design. A parallel effort focused on program support and technology development via the SCIT Task have enabled the progress.

Description: Advances in micro-fabrication processes have generated tremendous interests in miniaturizing chemical and biomedical analyses into integrated microsystems (Lab-on-Chip devices). To successfully design and operate the micro fluidics system, it is essential to understand the fundamental fluid flow phenomena when channel sizes are shrink to micron or even nano dimensions. One important phenomenon is the electro kinetic effect in micro/nano channels due to the existence of the electrical double layer (EDL) near a solid-liquid interface. Not only EDL is responsible for electro-osmosis pumping when an electric field parallel to the surface is imposed, EDL also causes extra flow resistance (the electro-viscous effect) and flow anomaly (such as early transition from laminar to turbulent flow) observed in pressure-driven microchannel flows. Modeling and simulation of electro-kinetic effects on micro flows poses significant numerical challenge due to the fact that the sizes of the double layer (10 nm up to microns) are very thin compared to channel width (can be up to 100 s of m). Since the typical thickness of the double layer is extremely small compared to the channel width, it would be computationally very costly to capture the velocity profile inside the double layer by placing sufficient number of grid cells in the layer to resolve the velocity changes, especially in complex, 3-d geometries. Existing approaches using "slip" wall velocity and augmented double layer are difficult to use when the flow geometry is complicated, e.g. flow in a T-junction, X-junction, etc. In order to overcome the difficulties arising from those two approaches, we have developed a sub-grid integration method to properly account for the physics of the double layer. The integration approach can be used on simple or complicated flow geometries. Resolution of the double layer is not needed in this approach, and the effects of the double layer can be accounted for at the same time. With this approach, the numeric grid size can be much larger than the thickness of double layer. Presented in this report are a description of the approach, methodology for implementation and several validation simulations for micro flows.

Description: The recent bold initiatives to expand the human presence in space require innovative approaches to the design of propulsion systems whose underlying technology is not yet mature. The space propulsion community has identified a number of candidate concepts. A short list includes solar sails, high-energy-density chemical propellants, electric and electromagnetic accelerators, solar-thermal and nuclear-thermal expanders. For each of these, the underlying physics are relatively well understood. One could easily cite authoritative texts, addressing both the governing equations, and practical solution methods for, e.g. electromagnetic fields, heat transfer, radiation, thermophysics, structural dynamics, particulate kinematics, nuclear energy, power conversion, and fluid dynamics. One could also easily cite scholarly works in which complete equation sets for any one of these physical processes have been accurately solved relative to complex engineered systems. The Advanced Concepts and Analysis Office (ACAO), Space Transportation Directorate, NASA Marshall Space Flight Center, has recently released the first alpha version of a set of computer utilities for performing the applicable physical analyses relative to candidate deep-space propulsion systems such as those listed above. PARSEC, Preliminary Analysis of Revolutionary in-Space Engineering Concepts, enables rapid iterative calculations using several physics tools developed in-house. A complete cycle of the entire tool set takes about twenty minutes. PARSEC is a level-zero/level-one design tool. For PARSEC s proof-of-concept, and preliminary design decision-making, assumptions that significantly simplify the governing equation sets are necessary. To proceed to level-two, one wishes to retain modeling of the underlying physics as close as practical to known applicable first principles. This report describes results of collaboration between ACAO, and Embry-Riddle Aeronautical University (ERAU), to begin building a set of level-two design tools for PARSEC. The "CFD Multiphysics Tool" will be the propulsive element of the tool set. The name acknowledges that space propulsion performance assessment is primarily a fluid mechanics problem. At the core of the CFD Multiphysics Tool is an open-source CFD code, HYP, under development at ERAU. ERAU is renowned for its undergraduate degree program in Aerospace Engineering the largest in the nation. The strength of the program is its applications-oriented curriculum, which culminates in one of three two-course Engineering Design sequences: Aerospace Propulsion, Spacecraft, or Aircraft. This same philosophy applies to the HYP Project, albeit with fluid physics modeling commensurate with graduate research. HYP s purpose, like the Multiphysics Tool s, is to enable calculations of real (three-dimensional; geometrically complex; intended for hardware development) applications of high speed and propulsive fluid flows.

Description: Volatile Removal Assembly (VRA) is a subsystem of the Closed Environment Life Support System (CELSS) installed in the International Space Station. It is used for removing contaminants (volatile organics) in the wastewater produced by the space station crews. The major contaminants are formic acid, ethanol, and propylene glycol. The VRA contains a slim packbed reactor (3.5 cm diameter and four 28 cm long tubes in series) to perform catalyst oxidation of wastewater at elevated pressure and temperature under microgravity conditions. In the reactor, the contaminants are burned with oxygen gas (O2) to form water and carbon dioxide (CO2) that dissolves in the water stream. Optimal design of the reactor requires a thorough understanding about how the reactor performs under microgravity conditions. The objective of this study was to develop a mathematical model to interpret experimental data obtained from normal and microgravity conditions, and to predict the performance of VRA reactor under microgravity conditions. Catalyst oxidation kinetics and the total oxygen-water contact area control the efficiency of catalyst oxidation for mass transfer, which depends on oxygen gas holdup and distribution in the reactor. The process involves bubbly flow in porous media with chemical reactions in microgravity environment. This presents a unique problem in fluid dynamics that has not been studied. Guo et al. (2004) developed a mathematical model that predicts oxygen holdup in the VRA reactor. No mathematical model has been found in the literature that can be used to predict the efficiency of catalyst oxidation under microgravity conditions.

Description: The development and analysis of turbomachinery components for industrial and aerospace applications has been greatly enhanced in recent years through the advent of computational fluid dynamics (CFD) codes and techniques. Although the use of this technology has greatly reduced the time required to perform analysis and design, there still remains much room for improvement in the process. In particular, there is a steep learning curve associated with most turbomachinery CFD codes, and the computation times need to be reduced in order to facilitate their integration into standard work processes. Two turbomachinery codes have recently been developed by Dr. Daniel Dorney (MSFC) and Dr. Douglas Sondak (Boston University). These codes are entitled Aardvark (for 2-D and quasi 3-D simulations) and Phantom (for 3-D simulations). The codes utilize the General Equation Set (GES), structured grid methodology, and overset O- and H-grids. The codes have been used with success by Drs. Dorney and Sondak, as well as others within the turbomachinery community, to analyze engine components and other geometries. One of the primary objectives of this study was to establish a set of parametric input values which will enhance convergence rates for steady state simulations, as well as reduce the runtime required for unsteady cases. The goal is to reduce the turnaround time for CFD simulations, thus permitting more design parametrics to be run within a given time period. In addition, other code enhancements to reduce runtimes were investigated and implemented. The other primary goal of the study was to develop enhanced users manuals for Aardvark and Phantom. These manuals are intended to answer most questions for new users, as well as provide valuable detailed information for the experienced user. The existence of detailed user s manuals will enable new users to become proficient with the codes, as well as reducing the dependency of new users on the code authors. In order to achieve the objectives listed, the following tasks were accomplished: 1) Parametric Study Of Preconditioning Parameters And Other Code Inputs; 2) Code Modifications To Reduce Runtimes; 3) Investigation Of Compiler Options To Reduce Code Runtime; and 4) Development/Enhancement of Users Manuals for Aardvark and Phantom

Description: Specific impulse is defined in words in many ways. Very early in any text on rocket propulsion a phrase similar to .specific impulse is the thrust force per unit propellant weight flow per second. will be found.(2) It is only after seeing the mathematics written down does the definition mean something physically to scientists and engineers responsible for either measuring it or using someone.s value for it.

Description: The present paper deals with an experimental study of the stagnation-point heat transfer to a cooled copper surface with gas injection under subsonic conditions. Test were made with a probe that combined a steady-state water-cooled calorimeter that allows the capability to study convective blockage and to perform heat transfer measurements in presence of gas injection in the stagnation region. The copper probe was pierced by 52 holes, representing 2.4% of the total probe surface. The 1.2 MW high enthalpy plasma wind tunnel was operated at anode powers between 130 and 230 kW and a static pressures from 35 hPa up to 200 hPa. Air, carbon dioxide and argon were injected in the mass flow range 0-0.4 g/s in the boundary layer developed around the 50 mm diameter probe. The measured stagnation-point heat transfer rates are reported and discussed.

Description: A new all-electronic Particle Image Velocimetry technique that can efficiently map high speed gas flows has been developed in-house at the NASA Lewis Research Center. Particle Image Velocimetry is an optical technique for measuring the instantaneous two component velocity field across a planar region of a seeded flow field. A pulsed laser light sheet is used to illuminate the seed particles entrained in the flow field at two instances in time. One or more charged coupled device (CCD) cameras can be used to record the instantaneous positions of particles. Using the time between light sheet pulses and determining either the individual particle displacements or the average displacement of particles over a small subregion of the recorded image enables the calculation of the fluid velocity. Fuzzy logic minimizes the required operator intervention in identifying particles and computing velocity. Using two cameras that have the same view of the illumination plane yields two single exposure image frames. Two competing techniques that yield unambiguous velocity vector direction information have been widely used for reducing the single-exposure, multiple image frame data: (1) cross-correlation and (2) particle tracking. Correlation techniques yield averaged velocity estimates over subregions of the flow, whereas particle tracking techniques give individual particle velocity estimates. For the correlation technique, the correlation peak corresponding to the average displacement of particles across the subregion must be identified. Noise on the images and particle dropout result in misidentification of the true correlation peak. The subsequent velocity vector maps contain spurious vectors where the displacement peaks have been improperly identified. Typically these spurious vectors are replaced by a weighted average of the neighboring vectors, thereby decreasing the independence of the measurements. In this work, fuzzy logic techniques are used to determine the true correlation displacement peak even when it is not the maximum peak, hence maximizing the information recovery from the correlation operation, maintaining the number of independent measurements, and minimizing the number of spurious velocity vectors. Correlation peaks are correctly identified in both high and low seed density cases. The correlation velocity vector map can then be used as a guide for the particle-tracking operation. Again fuzzy logic techniques are used, this time to identify the correct particle image pairings between exposures to determine particle displacements, and thus the velocity. Combining these two techniques makes use of the higher spatial resolution available from the particle tracking. Particle tracking alone may not be possible in the high seed density images typically required for achieving good results from the correlation technique. This two-staged velocimetric technique can measure particle velocities with high spatial resolution over a broad range of seeding densities.

Description: Nuclear electric propulsion has been identified as an enabling technology for future NASA space science missions, such as the Jupiter Icy Moons Orbiter (JIMO) now under study. An important element of the nuclear electric propulsion spacecraft is the power conversion system, which converts the reactor heat to electrical power for use by the ion propulsion system and other spacecraft loads. The electrical integration of the power converter and ion thruster represents a key technical challenge in making nuclear electric propulsion technology possible. This technical hurdle was addressed extensively on December 1, 2003, when a closed- Brayton-cycle power-conversion unit was tested with a gridded ion thruster at the NASA Glenn Research Center. The test demonstrated end-to-end power throughput and marked the first-ever coupling of a Brayton turbo alternator and a gridded ion thruster, both of which are candidates for use on JIMO-type missions. The testing was conducted at Glenn's Vacuum Facility 6, where the Brayton unit was installed in the 3-m-diameter vacuum test port and the ion thruster was installed in the 7.6-m-diameter main chamber.

Description: NASA Glenn Research Center s Capillary Flow Experiments (CFE) program is developing experiment payloads to explore fluid interfaces in microgravity on the International Space Station. The information to be gained from the CFE is relevant to the design of fluid-bearing systems in which capillary forces predominate, for example in the passive positioning of liquids in spacecraft fuel tanks. To achieve the science goals of CFE, Glenn researchers constructed several types of experiment vessels. One type of vessel, known as the interior corner flow (ICF), will be used to determine important transients for low-gravity liquid management in a two-phase system. Each vessel has a cylindrical fluid reservoir connected to each end of the test chamber by internal transport tubes, each with a quarter-turn shutoff valve (see the following photograph). These multipiece vessels are made from polymethylmethacrylate (PMMA) because of its excellent optical properties (i.e., the fluids can be observed easily in the vessel). Because of the complexity of certain vessels, the test chamber had to be manufactured in pieces and welded chemically. Some past experience with adhesive bonded plastic showed that the experiment fluid degraded the adhesive to the point of failure. Therefore, it was necessary to see if the fluid also degraded the chemically welded PMMA joints.

Description: Experimental data were obtained to help validate analytical and computational fluid dynamics (CFD) codes used to compute unsteady cascade aerodynamics in a supersonicaxial- flow regime. Results from two analytical codes and one CFD code were compared with experimental data. One analytical code did not account for airfoil thickness or camber; another, using piston theory (piston code), accounted for thickness and camber upstream of the first shockwave/airfoil impingement locations. The Euler CFD code accounted fully for airfoil shape.

Description: The NPARC (National Project for Application-oriented Research in CFD) Alliance has released Version 1.0 of Wind-US, the latest in its line of general-purpose, multizone, compressible-flow Navier-Stokes solvers. The NPARC Alliance is a formal partnership between the NASA Glenn Research Center and the Air Force Arnold Engineering Development Center, with additional significant involvement by the Boeing Company s Phantom Works Group, whose mission is to provide an applications-oriented computational fluid dynamics (CFD) system primarily for aerospace flow simulation. The alliance is committed to the long-range maintenance and improvement of this capability, with teams focused on user support, code development, and validation.

Description: Working in concert with the NASA Technology Transfer and Partnership Office, the Great Lakes Industrial Technology Center, and Alchemitron Corporation of Elgin, Illinois, the NASA Glenn Research Center has applied nonlinear acoustic principles to industrial applications. High-intensity ultrasonic beam techniques employ the effects of acoustic radiation pressure and acoustic streaming to manipulate the behavior of liquids. This includes propelling liquids, moving bubbles, and ejecting liquids as droplets and fountains. Since these effects can be accomplished without mechanical pumps or moving parts, we are exploring how these techniques could be used to manipulate liquids in space applications. Some of these acoustic techniques could be used both in normal Earth gravity and in the microgravity of space.

Description: The Microgravity Observation of Bubble Interactions (MOBI) experiment is working to better understand the physics of gas-liquid suspensions. To study such suspensions, researchers generate bubbles in a large cylindrical flow channel. Then, they use various types of instrumentation, including video imaging, to study the bubbly suspension. Scientists will need a camera view of the majority of the gas-liquid suspension inside of the couette in order to gather the information needed from the MOBI experiment. This will provide the scientists with a qualitative picture of the flow that may indicate flow instabilities or imperfect axial mixing inside the couette. These requirements pose a significant challenge because the imaging and lighting system must be confined to a very tight space since the space available on the International Space Station experiment racks is very limited. In addition, because of the large field of view needed and the detail needed to see the gas-liquid suspension behavior in the image, a digital video camera with high resolution (1024 by 1024 pixels) had to be used. Although the high-resolution camera will provide scientists with the image quality they need, it left little space on the experiment rack for the lighting system. Many configurations were considered for the lighting system, including front-lighting and back-lighting, but because of mechanical design limitations with the couette, back-lighting was not an option.

Description: NASA s vision for exploration will once again expand the bounds of human presence in the universe with planned missions to the Moon and Mars. To attain the numerous goals of this vision, NASA will need to develop technologies in several areas, including advanced power-generation and thermal-control systems for spacecraft and life support. The development of these systems will have to be demonstrated prior to implementation to ensure safe and reliable operation in reduced-gravity environments. The Two-Phase Flow Facility (T(PHI) FFy) Project will provide the path to these enabling technologies for critical multiphase fluid products. The safety and reliability of future systems will be enhanced by addressing focused microgravity fluid physics issues associated with flow boiling, condensation, phase separation, and system stability, all of which are essential to exploration technology. The project--a multiyear effort initiated in 2004--will include concept development, normal-gravity testing (laboratories), reduced gravity aircraft flight campaigns (NASA s KC-135 and C-9 aircraft), space-flight experimentation (International Space Station), and model development. This project will be implemented by a team from the NASA Glenn Research Center, QSS Group, Inc., ZIN Technologies, Inc., and the Extramural Strategic Research Team composed of experts from academia.

Description: Poor understanding of flow boiling in microgravity has recently emerged as a key obstacle to the development of many types of power generation and advanced life support systems intended for space exploration. The critical heat flux (CHF) is perhaps the most important thermal design parameter for boiling systems involving both heatflux-controlled devices and intense heat removal. Exceeding the CHF limit can lead to permanent damage, including physical burnout of the heat-dissipating device. The importance of the CHF limit creates an urgent need to develop predictive design tools to ensure both the safe and reliable operation of a two-phase thermal management system under the reduced-gravity (like that on the Moon and Mars) and microgravity environments of space. At present, very limited information is available on flow-boiling heat transfer and the CHF under these conditions.

Description: The volatilization of silica (SiO2) to silicon hydroxides and oxyhydroxides because of reaction with water vapor is important in a variety of high-temperature corrosion processes. For example, the lifetimes of silicon carbide (SiC) and silicon nitride (Si3N4) - based components in combustion environments are limited by silica volatility. To understand and model this process, it is essential to have accurate thermodynamic data for the formation of volatile silicon hydroxides and oxyhydroxides.

Description: In 2004, President Bush outlined a new space exploration vision for NASA. The exploration programs will seek profound answers to questions of our origins, whether life exists beyond Earth, and how we could live in other worlds. In response, research projects from NASA s Fluid Physics Research Program were moved into the Exploration Systems Mission Directorate and realigned to support the major milestones of this directorate. A new goal of this research is to obtain an understanding of the physical phenomena that are important in the design of the many space-based and ground-based fluids systems that utilize multiphase flow, such as life support, propulsion, and power systems.

Description: Early work within the Aqua validation activity revealed there to be large differences in water vapor measurement accuracy among the various technologies in use for providing validation data. The validation measurements were made at globally distributed sites making it difficult to isolate the sources of the apparent measurement differences among the various sensors, which included both Raman lidar and radiosonde. Because of this, the AIRS Water Vapor Experiment-Ground (AWEX-G) was held in October - November, 2003 with the goal of bringing validation technologies to a common site for intercomparison and resolution of the measurement discrepancies. Using the University of Colorado Cryogenic Frostpoint Hygrometer (CFH) as the water vapor reference, the AWEX-G field campaign resulted in new correction techniques for both Raman lidar, Vaisala RS80-H and RS90/92 measurements that significantly improve the absolute accuracy of those measurement systems particularly in the upper troposphere. Mean comparisons of radiosondes and lidar are performed demonstrating agreement between corrected sensors and the CFH to generally within 5% thereby providing data of sufficient accuracy for Aqua validation purposes. Examples of the use of the correction techniques in radiance and retrieval comparisons are provided and discussed.

Description: Wind tunnel experiments will continue to be a primary source of validation data for many types of mathematical and computational models in the aerospace industry. The increased emphasis on accuracy of data acquired from these facilities requires understanding of the uncertainty of not only the measurement data but also any correction applied to the data. One of the largest and most critical corrections made to these data is due to wall interference. In an effort to understand the accuracy and suitability of these corrections, a statistical validation process for wall interference correction methods has been developed. This process is based on the use of independent cases which, after correction, are expected to produce the same result. Comparison of these independent cases with respect to the uncertainty in the correction process establishes a domain of applicability based on the capability of the method to provide reasonable corrections with respect to customer accuracy requirements. The statistical validation method was applied to the version of the Transonic Wall Interference Correction System (TWICS) recently implemented in the National Transonic Facility at NASA Langley Research Center. The TWICS code generates corrections for solid and slotted wall interference in the model pitch plane based on boundary pressure measurements. Before validation could be performed on this method, it was necessary to calibrate the ventilated wall boundary condition parameters. Discrimination comparisons are used to determine the most representative of three linear boundary condition models which have historically been used to represent longitudinally slotted test section walls. Of the three linear boundary condition models implemented for ventilated walls, the general slotted wall model was the most representative of the data. The TWICS code using the calibrated general slotted wall model was found to be valid to within the process uncertainty for test section Mach numbers less than or equal to 0.60. The scatter among the mean corrected results of the bodies of revolution validation cases was within one count of drag on a typical transport aircraft configuration for Mach numbers at or below 0.80 and two counts of drag for Mach numbers at or below 0.90.

Description: The ultimate goal of this project is to integrate microfluidic devices with NASA's space bioreactor systems. In such a system, the microfluidic device would provide realtime feedback control of the bioreactor by monitoring pH, glucose, and lactate levels in the cell media; and would provide an analytical capability to the bioreactor in exterrestrial environments for monitoring bioengineered cell products and health changes in cells due to environmental stressors. Such integrated systems could be used as biosentinels both in space and on planet surfaces. The objective is to demonstrate the ability of microfabricated devices to repeatedly and reproducibly perform bead cytometry experiments in micro, lunar, martian, and hypergravity (1.8g).

Description: The mass-averaged compressible Navier-Stokes equations are solved for circulation control airfoils. Numerical solutions are computed with a multigrid method that uses an implicit approximate factorization smoother. The effects of flow conditions (e.g., free-stream Mach number, angle of attack, momentum coefficient) and mesh on the prediction of circulation control airfoil flows are considered. In addition, the impact of turbulence modeling, including curvature effects and modifications to reduce eddy viscosity levels in the wall jet (i.e., Coanda flow), is discussed. Computed pressure distributions are compared with available experimental data.

Description: In an effort to validate computational fluid dynamics procedures for calculating flows around circulation control airfoils, the commercial flow solver FLUENT was utilized to study the flow around a general aviation circulation control airfoil. The results were compared to experimental and computational fluid dynamics results conducted at the NASA Langley Research Center. The current effort was conducted in three stages: 1. A comparison of the results for free-air conditions to those from experiments. 2. A study of wind-tunnel wall effects. and 3. A study of the stagnation-point behavior.

Description: NASA is pursuing innovative technologies and concepts as part of America's Vision for Space Exploration. The rapidly emerging field of nanotechnology has led to new concepts for multipurpose shields to prevent catastrophic loss of vehicles and crew against the triple threats of aeroheating during atmospheric entry, radiation (Solar and galactic cosmic rays) and Micrometorid/Orbital Debris (MMOD) strikes. One proposed concept is the Thermal Radiation Impact Protection System (TRIPS) using carbon nanotubes, hydrogenated carbon nanotubes, and ceramic coatings as a multi-use TPS. The Thermophysics Facilities Branch of the Space Technology Division at NASA Ames Research Center provides testing services for the development and validation of the present and future concepts being developed by NASA and national and International research firms. The Branch operates two key facilities - the Range Complex and the Arc Jets. The Ranges include both the Ames Vertical Gun Range (AVGR) and the Hypervelocity Free Flight (HFF) gas guns best suited for MMOD investigations. Test coupons can be installed in the AVGR or HFF and subjected to particle impacts from glass or metal particles from micron to _ inch (6.35-mm) diameters and at velocities from 5 to 8 kilometers per second. The facility can record high-speed data on film and provide damage assessment for analysis by the Principle Investigator or Ames personnel. Damaged articles can be installed in the Arc Jet facility for further testing to quantify the effects of damage on the heat shield s performance upon entry into atmospheric environments.

Description: Over the past 30 years, numerical methods and simulation tools for fluid dynamic problems have advanced as a new discipline, namely, computational fluid dynamics (CFD). Although a wide spectrum of flow regimes are encountered in many areas of science and engineering, simulation of compressible flow has been the major driver for developing computational algorithms and tools. This is probably due to a large demand for predicting the aerodynamic performance characteristics of flight vehicles, such as commercial, military, and space vehicles. As flow analysis is required to be more accurate and computationally efficient for both commercial and mission-oriented applications (such as those encountered in meteorology, aerospace vehicle development, general fluid engineering and biofluid analysis) CFD tools for engineering become increasingly important for predicting safety, performance and cost. This paper presents the author's perspective on the maturity of CFD, especially from an aerospace engineering point of view.

Description: An investigation of the aerothermodynamic environment of the Huygens entry probe has been conducted. A Monte Carlo simulation of the trajectory of the probe during entry into Titan's atmosphere was performed to identify a worst-case heating rate trajectory. Flowfield and radiation transport computations were performed at points along this trajectory to obtain convective and radiative heat-transfer distributions on the probe's heat shield. This investigation identified important physical and numerical factors, including atmospheric CH4 concentration, transition to turbulence, numerical diffusion modeling, and radiation modeling, which strongly influenced the aerothermodynamic environment.

Description: Recent experimental investigations and physical modeling studies have indicated that turbulence behaviors within a liquid jet have considerable effects on the atomization process. This study aims to model the turbulence effect in the atomization process of a cylindrical liquid jet. Two widely used models, the Kelvin-Helmholtz (KH) instability of Reitz (blob model) and the Taylor-Analogy-Breakup (TAB) secondary droplet breakup by O Rourke et al, are further extended to include turbulence effects. In the primary breakup model, the level of the turbulence effect on the liquid breakup depends on the characteristic scales and the initial flow conditions. For the secondary breakup, an additional turbulence force acted on parent drops is modeled and integrated into the TAB governing equation. The drop size formed from this breakup regime is estimated based on the energy balance before and after the breakup occurrence. This paper describes theoretical development of the current models, called "T-blob" and "T-TAB", for primary and secondary breakup respectivety. Several assessment studies are also presented in this paper.

Description: Experiments Along the Coexistence near Tricriticality (EXACT) will test exact predictions made by Renormalization Group theory by mapping the phase diagram of liquid He-3 and He-4 mixtures at the tricritical point, Tcp=0.867 K. A PdMn based High Resolution Thermometer (HRT) will be utilized by EXACT to make accurate measurements with a resolution that has never been attained for the temperature range 0.7-1 K. The basic design of this mini high resolution thermometer comprises a sensing element whose magnetic susceptibility changes with temperature, a thermal connector, magnetic shielding, and some permanent magnets to apply a constant magnetic field. In this study, we will quantitatively determine the resolution of possible sensing element candidates of 0.15%, 0.20%, and 0.25% ppm Mn and compare them with an annealed group of PdMn with the corresponding concentrations to see how this processing technique affects sensitivity.

Description: In 1821 German Physicist Thomas J. Seebeck discovered that heat could be converted into electricity when a temperature difference was applied across two points on a material. Theoretically, the generated voltage has a directly proportional relationship with the temperature difference. This relationship is the Seebeck coefficient that scientists always referred to when determining the efficiency of a thermoelectricity convention. In our experiments, however, hysteresis loops appeared when we plotted voltage against temperature difference, and the measured Seebeck appeared differently when the measurements were run under vacuum, air, and helium gas. Measurements were done by using a low-frequency AC measuring method. By simulating the experimental setup into a; thermal circuit, we found that the loop and inconsistency in measuring Seebeck coefficient could be explained by studying the behaviors of a RC circuit in a thermal sense. Under vacuum, the gap of the hysteresis loop can be largely eliminated if the time period of the temperature difference increased up to 4800s. The trend of the variations in measuring Seebeck coefficients in different environments can also be predicted by using different thermal circuit models.

Description: High performance aircraft typically use hydrocarbon fuel to regeneratively cool the airframe and engine components. As the coolant temperatures increase, the fuel may react with dissolved oxygen forming deposits that limit the regenerative cooling system performance. This study investigates the deposition of Jet A using a thermal stability experiment and computational fluid dynamics (CFD) modeling. The experimental portion of this study is performed with a high Reynolds number thermal stability (HiRets) tester in which fuel passes though an electrically heated tube and the fuel outlet temperature is held constant. If the thermal stability temperature of the fuel is exceeded, deposits form and adhere to the inside of the tube creating an insulating layer between the tube and the fuel. The HiRets tester measures the tube outer wall temperatures near the fuel outlet to report the effect of deposition occurring inside the tube. Final deposits are also estimated with a carbon burn off analysis. The CFD model was developed and used to simulate the fluid dynamics, heat transfer, chemistry, and transport of the deposit precursors. The model is calibrated to the experiment temperature results and carbon burn-off deposition results. The model results show that the dominant factor in deposition is the heated wall temperature and that most of the deposits are formed in the laminar sublayer. The models predicted a 7.0E-6 kilograms per square meter-sec deposition rate, which compared well to the carbon burn-off analysis deposition rate of 1.0E-6 kilograms per square meter-sec.

Description: An ideal microcalorimeter is characterized by a constant energy resolution across the sensor's dynamic range. Any dependence of pulse shape on the position within the absorber where an event occurs leads to a degradation in resolution that is linear with event s energy (excess broadening). In this paper we present a numerical simulation that was developed to model the variation in pulse shape with position based on the thermal conductivity within the absorber and between the absorber, sensor, and heat bath, for arbitrarily shaped absorbers and sensors. All the parameters required for the simulation can be measured from actual devices. We describe how the thermal conductivity of the absorber material is determined by comparing the results of this model with data taken from a position sensitive detector in which any position dependent effect is purposely emphasized by making a long, narrow absorber that is read out by sensors on both end. Finally, we present the implications for excess broadening given the measured parameters of our X-ray microcalorimeters.

Description: In 1995, the NASA Lewis Research Center and the Kirby Company (a major vacuum cleaner company) began negotiations for a Space Act Agreement to conduct research, technology development, and testing involving the flow behavior of airborne particulate flow behavior. Through these research efforts, we hope to identify ways to improve suction, flow rate, and surface agitation characteristics of nozzles used in vacuum cleaner nozzles. We plan to apply an advanced visualization technology, known as Stereoscopic Imaging Velocimetry (SIV), to a Kirby G-4 vacuum cleaner. Resultant data will be analyzed with a high-speed digital motion analysis system. We also plan to evaluate alternative vacuum cleaner nozzle designs. The overall goal of this project is to quantify both velocity fields and particle trajectories throughout the vacuum cleaner nozzle to optimize its "cleanability"--its ability to disturb and remove embedded dirt and other particulates from carpeting or hard surfaces. Reference

Description: The next generation of aircraft fighters requires higher engine performance and enhanced stealth characteristics for air superiority. A Lockheed-Martin/Boeing aircraft with a Pratt & Whitney F119 engine was selected by the Air Force for the next advanced tactical fighter (ATF). As part of this program, the NASA Lewis Research Center entered into a cooperative test program with Pratt & Whitney along with the Air Force to study the performance for various advanced nozzle concepts for the F119 engine. The area of interest was to measure the internal performance (both thrust and flow coefficients) of nozzle flaps redesigned for low observability with minimal performance loss. The experimental program was successfully completed May 1995 in Lewis' CE-22 facility. The models were tested over a wide range of geometric variations and nozzle pressure ratios. Results confirmed that the redesigned nozzle flaps had an insignificant effect on the thrust performance and that the resulting flow patterns should not be a problem in the cooling flow design. The results also agreed well with Pratt & Whitney's computational fluid dynamics analysis. The data obtained from this test were added to the current data base to help validate other performance prediction methodology.

Description: An operational change made recently in the drive motor system for the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT)/9- by 15-Foot Low-Speed Wind Tunnel (9x15 LSWT) complex resulted in dramatic power savings and expanded operating range. The 8x6 SWT/9x15 LSWT complex offers a unique combination of wind tunnel conditions for both high- and low-speed testing. Prior to the work discussed in this article, the 8- by 6-ft test section offered airflows ranging from Mach 0.36 to 2.0. Subsonic testing was done in the 9-ft high, 15-ft wide test area in the return leg of the facility. The air speed in this test section can range from 0 to 175 mph (Mach 0.23). In the past, we varied the air speed by using a combination of the compressor speed and the position of the tunnel flow-control doors. When very slow speeds were required in the 9x15 LSWT, these large tunnel flow control doors might be very nearly full open, bleeding off large quantities of air, even with the drive system operating at its previous minimum speed of about 510 rpm. Power drawn during this mode of operation varied between 15 and 18 MW/hr, but clearly much of this power was not being used to provide air that would be used for testing in the test section. The air exiting these large doors represented wasted power. Early this year, the facility's tunnel drive system was run on one motor instead of three to see if lower drive speeds could be achieved that would, in turn, result in large power savings because unnecessary air would not be blown out of the flow-control doors unnecessarily. In addition, if the drive could be run slower, then slower speeds would also be possible in the 8x6 SWT test section as an added benefit. Results of the first tests performed early last year showed that in fact the drive, when operating on only one motor, actually reached a steady-state speed of only 337 rpm and drew an amazingly small 6 MW/hr of electrical power. During daytime operation of the drive, this meant that it would be possible to save as much as 10 MW/hr, or nearly $600 per hour of operation, for many of the 9x15 LSWT's testing regimes. An added benefit of this power-saving venture was that since the 8x6 SWT and 9x15 LSWT are indeed on a common loop, if the compressor is slowed down to benefit the 9x15 LSWT, then the air moving through the 8x6 SWT is also moving slower than ever before. In fact, testing has proven that the 8x6 SWT can now achieve Mach 0.25, whereas its previous lower limit was Mach 0.36. This added benefit has attracted additional customers

Description: Computational Fluid Dynamics (CFD) is increasingly being used to both augment and create an aerodynamic performance database for aircraft configurations. This aerodynamic database contains the response of the aircraft to varying flight conditions and control surface deflections. The current work presents a novel method for calculating dynamic stability derivatives which reduces the computational cost over traditional unsteady CFD approaches by an order of magnitude, while still being applicable to arbitrarily complex geometries over a wide range of flow regimes. The primary thesis of this work is that the response to a forced motion can often be represented with a small, predictable number of frequency components without loss of accuracy. By resolving only those frequencies of interest, the computational effort is significantly reduced so that the routine calculation of dynamic derivatives becomes practical. The current implementation uses this same non-linear, frequency-domain approach and extends the application to the 3-D Euler equations. The current work uses a Cartesian, embedded-boundary method to automate the generation of dynamic stability derivatives.

Description: After video analysis of space shuttle flight STS-107's ascent showed that an object shed from the bipod-ramp region impacted the left wing, a transport analysis was initiated to determine a credible flight path and impact velocity for the piece of debris. This debris transport analysis was performed both during orbit, and after the subsequent re-entry accident. The analysis provided an accurate prediction of the velocity a large piece of foam bipod ramp would have as it impacted the wing leading edge. This prediction was corroborated by video analysis and fully-coupled CFD/six degree of freedom (DOF) simulations. While the prediction of impact velocity was accurate enough to predict critical damage in this case, one of the recommendations of the Columbia Accident Investigation Board (CAIB) for return-to-flight (RTF) was to analyze the complete debris environment experienced by the shuttle stack on ascent. This includes categorizing all possible debris sources, their probable geometric and aerodynamic characteristics, and their potential for damage. This paper is chiefly concerned with predicting the aerodynamic characteristics of a variety of potential debris sources (insulating foam and cork, nose-cone ablator, ice, ...) for the shuttle ascent configuration using CFD methods. These aerodynamic characteristics are used in the debris transport analysis to predict flight path, impact velocity and angle, and provide statistical variation to perform risk analyses where appropriate. The debris aerodynamic characteristics are difficult to determine using traditional methods, such as static or dynamic test data, due to the scaling requirements of simulating a typical debris event. The use of CFD methods has been a critical element for building confidence in the accuracy of the debris transport code by bridging the gap between existing aerodynamic data and the dynamics of full-scale, in-flight events.

Description: The scaling and similarity laws concerning the propagation of isolated spherical blast waves are briefly reviewed. Both point source explosions and high pressure gas explosions are considered. Test data on blast overpressure from the interaction and coalescence of spherical blast waves emanating from explosives in the form of shaped charges of different strength placed in the vicinity of a solid propellant stack are presented. These data are discussed with regard to the scaling laws concerning the decay of blast overpressure. The results point out the possibility of detecting source explosions from far-field pressure measurements.

Description: The National Aeronautics and Space Administration (NASA) seeks to license its Thermal Insulation Test Apparatuses. Designed by the Cryogenics Test Laboratory at the John F. Kennedy Space Center (KSC) in Florida, these patented technologies (U.S. Patent Numbers: Cryostat 1 - 6,742,926, Cryostat 2 - 6,487,866, and Cryostat 4 - 6,824,306) allow manufacturers to fabricate and test cryogenic insulation at their production and/or laboratory facilities. These new inventions allow for the thermal performance characterization of cylindrical and flat specimens (e.g., bulk-fill, flat-panel, multilayer, or continuously rolled) over the full range of pressures, from high vacuum to no vacuum, and over the full range of temperatures from 77K to 300K. In today's world, efficient, low-maintenance, low-temperature refrigeration is taking a more significant role, from the food industry, transportation, energy, and medical applications to the Space Shuttle. Most countries (including the United States) have laws requiring commercially available insulation materials to be tested and rated by an accepted methodology. The new Cryostat methods go beyond the formal capabilities of the ASTM methods to provide testing for real systems, including full-temperature differences plus full-range vacuum conditions.

Description: A method for rapidly estimating the aeroheating, shear stress, and other properties of hypersonic flow about a three-dimensional (3D) blunt body has been devised. First, the geometry of the body is specified in Cartesian coordinates. The surface of the body is then described by its derivatives, coordinates, and principal curvatures. Next, previously relatively simple equations are used to find, for each desired combination of angle of attack and meridional angle, a scaling factor and the shape of an equivalent axisymmetric body. These factors and equivalent shapes are entered as inputs into a previously developed computer program that solves the two-dimensional (2D) equations of flow in a non-equilibrium viscous shock layer (VSL) about an axisymmetric body. The coordinates in the output of the VSL code are transformed back to the Cartesian coordinates of the 3D body, so that computed flow quantities can be registered with locations in the 3D flow field of interest. In tests in which the 3D bodies were elliptic paraboloids, the estimates obtained by use of this method were found to agree well with solutions of 3D, finite-rate-chemistry, thin-VSL equations for a catalytic body.

Description: The Cryogenic Tank Analysis Program (CTAP) predicts the time-varying thermodynamic state of a cryogenic fluid in a tank or a Dewar flask. CTAP is designed to be compatible with EASY5x, which is a commercial software package that can be used to simulate a variety of processes and equipment systems. The mathematical model implemented in CTAP is a first-order differential equation for the pressure as a function of time.

Description: A capability for real-time computational simulation of aeroheating has been developed in support of the Hyper-X program, which is directed toward demonstrating the feasibility of operating an air-breathing ramjet/scramjet engine at mach 5, mach 7, and mach 10. The simulation software will serve as a valuable design tool for initial trajectory studies in which aerodynamic heating is expected to exert a major influence in the design of the Hyper-X airplane; this tool will aid in the selection of materials, sizing of structural skin thicknesses, and selection of components of a thermal-protection system (TPS) for structures that must be insulated against aeroheating.

Description: In designing systems for the long-term storage of cryogens in low gravity space environments, one must consider the effects of thermal stratification on excessive tank pressure that will occur due to environmental heat leakage. During low gravity operations, a Thermodynamic Venting System (TVS) concept is expected to maintain tank pressure without propellant resettling. The TVS consists of a recirculation pump, Joule-Thomson (J-T) expansion valve, and a parallel flow concentric tube heat exchanger combined with a longitudinal spray bar. Using a small amount of liquid extracted by the pump and passing it though the J-T valve, then through the heat exchanger, the bulk liquid and ullage are cooled, resulting in lower tank pressure. A series of TVS tests were conducted at the Marshall Space Flight Center using liquid nitrogen as a liquid oxygen simulant. The tests were performed at fill levels of 90%, 50%, and 25% with gaseous nitrogen and helium pressurants, and with a tank pressure control band of 7 kPa. A transient one-dimensional model of the TVS is used to analyze the data. The code is comprised of four models for the heat exchanger, the spray manifold and injector tubes, the recirculation pump, and the tank. The TVS model predicted ullage pressure and temperature and bulk liquid saturation pressure and temperature are compared with data. Details of predictions and comparisons with test data regarding pressure rise and collapse rates will be presented in the final paper.

Description: Capillary phenomena plays an important role in the management of product water in PEM fuel cells because of the length scales associated with the porous layers and the gas flow channels. The distribution of liquid water within the network of gas flow channels can be dramatically altered by capillary flow. We experimentally demonstrate the rapid movement of significant volumes of liquid via capillarity through thin liquid films which connect orthogonal channels. The microfluidic experiments discussed provide a good benchmark against which the proper modeling of capillarity by computational models may be tested. The effect of surface wettability, as expressed through the contact angle, on capillary flow will also be discussed.

Description: When surface tension forces dominate, and regardless of whether the situation is static or dynamic, the contact angle (the angle the interface between two immiscible fluids makes when it contacts a solid) is the key parameter that determines the shape of a fluid-fluid interface. The static contact angle is easy to measure and implement in models predicting static capillary surface shapes and such associated quantities as pressure drops. By contrast, when the interface moves relative to the solid (as in dynamic wetting processes) the dynamic contact angle is not identified unambiguously because it depends on the geometry of the system Consequently, its determination becomes problematic and measurements in one geometry cannot be applied in another for prediction purposes. However, knowing how to measure and use the dynamic contact angle is crucial to determine such dynamics as a microsystem throughput reliably. In this talk we will present experimental and analytical efforts aimed at resolving modeling issues present in dynamic wetting. We will review experiments that show the inadequacy of the usual hydrodynamic model when a fluid-fluid meniscus moves over a solid surface such as the wall of a small tube or duct. We will then present analytical results that show how to parametrize these problems in a predictive manner. We will illustrate these ideas by showing how to implement the method in numerical fluid mechanical calculations.

Description: Estimating how free energy changes with the state of a system is a central goal in applications of statistical mechanics to problems of chemical or biological interest. From these free energy changes it is possible, for example, to establish which states of the system are stable, what are their probabilities and how the equilibria between these states are influenced by external conditions. Free energies are also of great utility in determining kinetics of transitions between different states. A variety of methods have been developed to compute free energies of condensed phase systems. Here, I will focus on one class of methods - those that allow for calculating free energy changes along one or several generalized coordinates in the system, often called reaction coordinates or order parameters . Considering that in almost all cases of practical interest a significant computational effort is required to determine free energy changes along such coordinates it is hardly surprising that efficiencies of different methods are of great concern. In most cases, the main difficulty is associated with its shape along the reaction coordinate. If the free energy changes markedly along this coordinate Boltzmann sampling of its different values becomes highly non-uniform. This, in turn, may have considerable, detrimental effect on the performance of many methods for calculating free energies.

Description: The ability to estimate short-term fluxes of water vapor from the land surface is important for validating latent heat flux estimates from high resolution remote sensing techniques. A new, relatively inexpensive method is presented for estimating t h e ground-based values of the surface latent heat flux or evapotranspiration.

Description: Two-phase flow systems play vital roles in the design of some current and anticipated space applications of two-phase systems which include: thermal management systems, transfer line flow in cryogenic storage, space nuclear power facilities, design and operation of thermal bus, life support systems, propulsion systems, In Situ Resource Utilization (ISRU), and space processes for pharmaceutical applications. The design of two-phase flow systems for space applications requires a clear knowledge of the behaviors of the dispersed phase (bubble), its interaction with the continuous phase (liquid) and its effect on heat and mass transfer processes, The need to understand the bubble generation process arises from the fact that for all space applications, the size and distribution of bubbles are extremely crucial for heat and mass transfer control. One important force in two-phase flow systems is the lift force on a bubble or particle in a liquid shear flow. The shear lift is usually overwhelmed by buoyancy in normal gravity, but it becomes an important force in reduced gravity. Since the liquid flow is usually sheared because of the confining wall, the trajectories of bubbles and particles injected into the liquid flow are affected by the shear lift in reduced gravity. A series of experiments are performed to investigate the lift force on a bubble in a liquid shear flow and its effect on the detachment of a bubble from a wall under low gravity conditions. Experiments are executed in a Poiseuille flow in a channel. An air-water system is used in these experiments that are performed in the 2.2 second drop tower. A bubble is injected into the shear flow from a small injector and the shear lift is measured while the bubble is held stationary relative to the fluid. The trajectory of the bubble prior, during and after its detachment from the injector is investigated. The measured shear lift force is calculated from the trajectory of the bubble at the detachment point. These values for the shear lift are then compared with the theoretical predictions from various published works on shear lift in the open literature, which include asymptotic solutions at low bubble Reynolds number, potential flow predictions and numerical studies that deal with intermediate bubble Reynolds numbers.

Description: A rotor blade system with reduced blade-vortex interaction noise includes a plurality of tube members embedded in proximity to a tip of each rotor blade. The inlets of the tube members are arrayed at the leading edge of the blade slightly above the chord plane, while the outlets are arrayed at the blade tip face. Such a design rapidly diffuses the vorticity contained within the concentrated tip vortex because of enhanced flow mixing in the inner core, which prevents the development of a laminar core region.

Description: The adjoint parabold stability equations (PSE) formulation is used to calculate the boundary layer receptivity to localized surface roughness and suction for compressible boundary layers. Receptivity efficiency functions predicted by the adjoint PSE approach agree well with results based on other nonparallel methods including linearized Navier-Stokes equations for both Tollmien-Schlichting waves and crossflow instability in swept wing boundary layers. The receptivity efficiency function can be regarded as the Green's function to the disturbance amplitude evolution in a nonparallel (growing) boundary layer. Given the Fourier transformed geometry factor distribution along the chordwise direction, the linear disturbance amplitude evolution for a finite size, distributed nonuniformity can be computed by evaluating the integral effects of both disturbance generation and linear amplification. The synergistic approach via the linear adjoint PSE for receptivity and nonlinear PSE for disturbance evolution downstream of the leading edge forms the basis for an integrated transition prediction tool. Eventually, such physics-based, high fidelity prediction methods could simulate the transition process from the disturbance generation through the nonlinear breakdown in a holistic manner.

Description: A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76 deg/40 deg double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 30 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M = 0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Description: The dynamics of an ensemble of linear disturbances in boundary-layer flows at various Reynolds numbers is studied through an analysis of the transport equations for the mean disturbance kinetic energy and energy dissipation rate. Effects of adverse and favorable pressure-gradients on the disturbance dynamics are also included in the analysis Unlike the fully turbulent regime where nonlinear phase scrambling of the fluctuations affects the flow field even in proximity to the wall, the early stage transition regime fluctuations studied here are influenced cross the boundary layer by the solid boundary. The dominating dynamics in the disturbance kinetic energy and dissipation rate equations are described. These results are then used to formulate transition-sensitized turbulent transport equations, which are solved in a two-step process and applied to zero-pressure-gradient flow over a flat plate. Computed results are in good agreement with experimental data.

Description: The effect of discrete contour bumps on reducing the transonic drag at off-design conditions on an airfoil have been examined. The research focused on fully-turbulent flow conditions, at a realistic flight chord Reynolds number of 30 million. State-of-the-art computational fluid dynamics methods were used to design a new baseline airfoil, and a family of fixed contour bumps. The new configurations were experimentally evaluated in the 0.3-m Transonic Cryogenic Tunnel at the NASA Langley Research center, which utilizes an adaptive wall test section to minimize wall interference. The computational study showed that transonic drag reduction, on the order of 12% - 15%, was possible using a surface contour bump to spread a normal shock wave. The computational study also indicated that the divergence drag Mach number was increased for the contour bump applications. Preliminary analysis of the experimental data showed a similar contour bump effect, but this data needed to be further analyzed for residual wall interference corrections.

Description: The object of this paper is to investigate the feasibility of applying CFD methods to aerodynamic analyses for aircraft stability and control. The integrated aerodynamic parameters used in stability and control, however, are not necessarily those extensively validated in the state of the art CFD technology. Hence, an exploratory study of such applications and the comparison of the solutions to available experimental data will help to assess the validity of the current computation methods. In addition, this study will also examine issues related to wind tunnel measurements such as measurement uncertainty and support interference effects. Several sets of experimental data from the NASA Langley 14x22-Foot Subsonic Tunnel and the National Transonic Facility are presented. Two Navier-Stokes flow solvers, one using structured meshes and the other unstructured meshes, were used to compute longitudinal static stability derivatives for an advanced Blended Wing Body configuration over a wide range of angles of attack. The computations were performed for two different Reynolds numbers and the resulting forces and moments are compared with the above mentioned wind tunnel data.

Description: An airfoil having a fore airfoil element, an aft airfoil element, and a slot region in between them. These elements induce laminar flow over substantially all of the fore airfoil element and also provide for laminar flow in at least a portion of the slot region. The method of the invention is one for inducing natural laminar flow over an airfoil. In the method, a fore airfoil element, having a leading and trailing edge, and an aft airfoil element define a slot region. Natural laminar flow is induced over substantially all of the fore airfoil element, by inducing the pressures on both surfaces of the fore airfoil element to decrease to a location proximate the trailing edge of the fore airfoil element using pressures created by the aft airfoil element.

Description: The present invention relates to a heat exchanger panel which has broad utility in high temperature environments. The heat exchanger panel has a first panel, a second panel, and at least one fluid containment device positioned intermediate the first and second panels. At least one of the first panel and the second panel have at least one feature on an interior surface to accommodate the at least one fluid containment device. In a preferred embodiment, each of the first and second panels is formed from a high conductivity, high temperature composite material. Also, in a preferred embodiment, the first and second panels are joined together by one or more composite fasteners.

Description: A multi-element rotor blade includes an individually controllable main element and fixed aerodynamic surface in an aerodynamically efficient location relative to the main element. The main element is controlled to locate the fixed aerodynamic surface in a position to increase lift and/or reduce drag upon the main element at various azimuthal positions during rotation.

Description: A gas-liquid separator uses a helical passageway to impart a spiral motion to a fluid passing therethrough. The centrifugal fore generated by the spiraling motion urges the liquid component of the fluid radially outward which forces the gas component radially inward. The gas component is then filtered through a gas-permeable, liquid-impervious membrane and discharged through a central passageway.

Description: If a CFD code/model developer is asked what experimental data he wants to validate his code or numerical model, his answer will be: "Everything, everywhere, at all times." Since this is not possible, practical, or even reasonable, the developer must understand what can be measured within the limits imposed by the test article, the test location, the test environment and the available diagnostic equipment. At the same time, it is important for the expermentalist/diagnostician to understand what the CFD developer needs (as opposed to wants) in order to conduct a useful CFD validation experiment. If these needs are not known, it is possible to neglect easily measured quantities at locations needed by the developer, rendering the data set useless for validation purposes. It is also important for the experimentalist/diagnostician to understand what the developer is trying to validate so that the experiment can be designed to isolate (as much as possible) the effects of a particular physical phenomena that is associated with the model to be validated. The probability of a successful validation experiment can be greatly increased if the two groups work together, each understanding the needs and limitations of the other.

Description: A fluid-control electro-active device includes a piezo-diaphragm made from a ferroelectric material sandwiched by first and second electrode patterns configured to introduce an electric field into the ferroelectric material when voltage is applied thereto. The electric field originates at a region of the ferroelectric material between the first and second electrode patterns, and extends radially outward from this region of the ferroelectric material and substantially parallel to the plane of the ferroelectric material. The piezo-diaphragm deflects symmetrically about this region in a direction substantially perpendicular to the electric field. An annular region coupled to and extending radially outward from the piezo-diaphragm perimetrically borders the piezo-diaphragm, A housing is connected to the region and at least one fluid flow path with piezo-diaphragm disposed therein.

Description: The Wind-US Computational Fluid Dynamics flow solver computed flow solutions for a transonic diffusing duct. The calculations used an unstructured (hexahedral) grid. The Spalart-Allmaras turbulence model was used. Static pressures along the upper and lower wall agreed well with experiment, as did velocity profiles. The effect of the smoothing input parameters on convergence and solution accuracy was investigated. The meaning and proper use of these parameters are discussed for the benefit of Wind-US users. Finally, the unstructured solver is compared to the structured solver in terms of run times and solution accuracy.

Description: An experimental study has been performed to develop a large force and moment aerodynamic data set on a slender axisymmetric missile configuration having cruciform strakes and in-line control tail fins. The data include six-component balance measurements of the configuration aerodynamics and three-component measurements on all four tail fins. The test variables include angle of attack, roll angle, Mach number, model buildup, strake length, nose size, and tail fin deflection angles to provide pitch, yaw, and roll control. Test Mach numbers ranged from 0.60 to 4.63. The entire data set is presented on a CD-ROM that is attached to this paper. The CD-ROM also includes extensive plots of both the six-component configuration data and the three-component tail fin data. Selected samples of these plots are presented in this paper to illustrate the features of the data and to investigate the effects of the test variables.

Description: A wind tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) on a six percent thick slightly cambered elliptical circulation control airfoil with both upper and lower surface blowing capability. Parametric evaluations of jet slot heights and Coanda surface shapes were conducted at momentum coefficients (Cm) from 0.0 to 0.12. Test data were acquired at Mach numbers of 0.3, 0.5, 0.7, 0.8, and 0.84 at Reynolds numbers per foot of 2.43 x 105 to 1.05 x 106. For a transonic condition, (Mach = 0.8 at alpha = 3 degrees), it was generally found the smaller slot and larger Coanda surface combination was overall more effective than other slot/Coanda surface combinations. Lower surface blowing was not as effective as the upper surface blowing over the same range of momentum coefficients. No appreciable Coanda surface, slot height, or slot blowing position preference was indicated transonically with the dual slot blowing.

Description: During Phase I of this project, Raytheon Aircraft Company (RAC) has analytically and experimentally evaluated key components of a system that could be implemented for active tailoring of wing lift distribution using low-drag, trailing-edge modifications. Simple systems such as those studied by RAC could be used to enhance the cruise performance of a business jet configuration over a range of typical flight conditions. The trailing-edge modifications focus on simple, deployable mechanisms comprised of extendable small flap panels over portions of the span that could be used to subtly but positively optimize the lift and drag characteristics. The report includes results from low speed wind tunnel testing of the trailing-edge devices, descriptions of potential mechanisms for automation, and an assessment of the technology.

Description: This research project focuses on the dynamics of interfacial regions between miscible fluids. While much attention has focused on immiscible interfaces in the past, miscible interfaces have been explored to a much lesser degree, so that there are many open questions regarding their dynamics at this time. Among the more pressing issues is the role that nonconventional stresses can play in such interfacial regions. Such stresses are typically not accounted for in efforts to model the dynamics of miscible flows. Our research aims to clarify under which circumstances these stresses do have to be taken into account, and what quantitative approaches are most suitable in this regard. In order to address these issues, we have focused on conducting linear stability analyses and nonlinear simulations for capillary tube and Hele-Shaw flows, and to compare the results with corresponding experiments performed in the labs of our co-investigators Prof. Maxworthy at USC, and Dr. Balasubramaniam at NASA. Over the duration of the project we have, among other things, focused on the effects of variable diffusion coefficients in such flows, and specifically on their influence in the growth of instabilities. Furthermore, our three-dimensional spectral element simulations have made good progress, so that we have come to a point where we can conduct more detailed comparisons with experimental observations. We are currently focusing our efforts on reproducing the tip-splitting instability observed by Maxworthy. Finally, we have discovered a new core-annular flow instability in the Stokes flow regime during the last year. This represents a significant finding, as this instability does not have an immiscible counterpart.

Description: Method and system for analyzing aircraft data, including multiple selected flight parameters for a selected phase of a selected flight, and for determining when the selected phase of the selected flight is atypical, when compared with corresponding data for the same phase for other similar flights. A flight signature is computed using continuous-valued and discrete-valued flight parameters for the selected flight parameters and is optionally compared with a statistical distribution of other observed flight signatures, yielding atypicality scores for the same phase for other similar flights. A cluster analysis is optionally applied to the flight signatures to define an optimal collection of clusters. A level of atypicality for a selected flight is estimated, based upon an index associated with the cluster analysis.

Description: In an attempt to explain the high loss of panels from the south face of the Vehicle Assembly Building (VAB) during Hurricane Frances, a three-dimensional computational fluid dynamics (3-D CFD) model was developed to simulate local velocity and pressure distributions resulting from such a storm. A preconditioned compressible Navier-Stokes flow solver 1 was used to compute the flow field around the VAB complex, including the Launch Control Center, the Low and High Bays of the VAB, and several outbuildings in the immediate LC-39 area. The mapping of the forces and velocities on and along the affected faces of the VAB correlated surprisingly well with the extensive damage areas realized on both on the south face and on the southeast section of the roof. The model results were also consistent with the minimal damage seen on the east, north, and west faces of the structure.

Description: Flowfield testing of a 1/5th scale model of a payload/fairing configuration, typical of an expendable launch vehicle, has been performed. Two-dimensional (planar) velocity measurements were carried out in four planes with the aid of Laser Doppler Velocimetry (LDV). Computational Fluid Dynamics (CFD) analysis results for the scale model flowfleld are compared with the test data. The CFD results are in general agreement with the test data. The ability of the CFD methodology in identifying the global flow features (including critical points such as vortex, saddle point, etc.) has been demonstrated. Practical problems and difficulties associated with the LDV method applied to the complex geometry under consideration have been summarized.

Description: The aerodynamic and aerothermal heating at the leeward surfaces of the Apollo capsule flying high angle of attack were found difficult to simulate using conventional tools. Due to the large subsonic region in the proximity of the shoulder of the base shield, correlation-based tools were found to be inadequate. CFD tools are too time consuming for conceptual design purposes, and cannot account for the transient effects of material response such as wall temperature and blowing. An accurate and timely simulation is essential to effectively size the thermal protection system (TPS), to enhance its performance, and to ensure the safety of the crew. Northrop Grumman Corporation applied a modified version of the MASCC/ATAC program, an inviscid flowfield code with boundary layer solver to simulate this scenario. The MASCC/ATAC code is believed to be the only non-CFD code that can rigorously perform the simulations on the windward surface. The predictions on the windward side of the conical surface were found to be in good agreement with flight data over a wide range of environments. The results are presented in the paper.

Description: This paper describes model structures and parameter estimation algorithms suitable for the identification of unsteady aerodynamic models from input-output data. The model structures presented are state space models and include linear time-invariant (LTI) models and linear parameter-varying (LPV) models. They cover a wide range of local and parameter dependent identification problems arising in unsteady aerodynamics and nonlinear flight dynamics. We present a residue algorithm for estimating model parameters from data. The algorithm can incorporate apriori information and is described in detail. The algorithms are evaluated on the F-16XL wind-tunnel test data from NAS Langley Research Center. Results of numerical evaluation are presented. The paper concludes with a discussion major issues and directions for future work.

Description: Multistage, explicit time stepping can be tailored to accelerate convergence for scalar advection-diffusion problems by using optimized multistage coefficients that vary with the local cell Reynolds number. And, when combined with local preconditioning, variable-coefficient multistage schemes for computational fluid dynamics codes can also provide an order of magnitude faster convergence, relative to standard, fixed-coefficient schemes, for the Navier-Stokes system of equations.

Description: Planar laser-induced fluorescence (PLIF) is demonstrated to be a valuable tool for studying the onset of transition to turbulence. For this study, we have used PLIF of nitric oxide (NO) to image underexpanded axisymmetric free jets issuing into a low-pressure chamber through a smooth converging nozzle with a sonic orifice. Flows were studied over a range of Reynolds numbers and nozzle-exit-to-ambient pressure ratios with the aim of empirically determining criteria governing the onset of turbulence. We have developed an image processing technique, involving calculation of the standard deviation of the intensity in PLIF images, in order to aid in the identification of turbulence. We have used the resulting images to identify laminar, transitional and turbulent flow regimes. Jet scaling parameters were used to define a rescaled Reynolds number that incorporates the influence of a varying pressure ratio. An empirical correlation was found between transition length and this rescaled Reynolds number for highly underexpanded jets.

Description: Significant advances have been made to non-intrusive flow field diagnostics in the past decade. Camera based techniques are now capable of determining physical qualities such as surface deformation, surface pressure and temperature, flow velocities, and molecular species concentration. In each case, extracting the pertinent information from the large volume of acquired data requires powerful and efficient data visualization tools. The additional requirement for real time visualization is fueled by an increased emphasis on minimizing test time in expensive facilities. This paper will address a capability titled LiveView3D, which is the first step in the development phase of an in depth, real time data visualization and analysis tool for use in aerospace testing facilities.

Description: The flowfield characteristics in a rocket engine coolant channels are analyzed by use of a commercial CFD and multiphysics software developed by the CFD Research Corp. called CFD-ACE+. The channels are characterized by high Reynolds number flow, varying aspect ratio, varying curvature, asymmetric and symmetric heating. The supercritical hydrogen coolant introduces large property variations that have a strong influence on the developing flow and the resulting heat transfer. This paper only shows the effect of aspect ratio and curvature for constant properties.

Description: Wall heat flux measurements in a 1.5 in. diameter circular cross-section rocket chamber for a uni-element shear coaxial injector element operating on gaseous oxygen (GOz)/gaseous hydrogen (GH,) propellants are presented. The wall heat flux measurements were made using arrays of Gardon type heat flux gauges and coaxial thermocouple instrumentation. Wall heat flux measurements were made for two cases. For the first case, GOZ/GHz oxidizer-rich (O/F=l65) and fuel-rich preburners (O/F=1.09) integrated with the main chamber were utilized to provide vitiated hot fuel and oxidizer to the study shear coaxial injector element. For the second case, the preburners were removed and ambient temperature gaseous oxygen/gaseous hydrogen propellants were supplied to the study injector. Experiments were conducted at four chamber pressures of 750, 600, 450 and 300psia for each case. The overall mixture ratio for the preburner case was 6.6, whereas for the ambient propellant case, the mixture ratio was 6.0. Total propellant flow was nominally 0.27-0.29 Ibm/s for the 750 psia case with flowrates scaled down linearly for lower chamber pressures. The axial heat flux profile results for both the preburner and ambient propellant cases show peak heat flux levels a t axial locations between 2.0 and 3.0 in. from the injector face. The maximum heat flux level was about two times greater for the preburner case. This is attributed to the higher injector fuel-to-oxidizer momentum flux ratio that promotes mixing and higher initial propellant temperature for the preburner case which results in a shorter reaction zone. The axial heat flux profiles were also scaled with respect to the chamber pressure to the power 0.8. The results at the four chamber pressures for both cases collapsed to a single profile indicating that at least to first approximation, the basic fluid dynamic structures in the flow field are pressure independent as long as the chamber/njector/nozzle geometry and injection velocities remain the same.

Description: The fluid in the Internal Active Thermal Control System (IATCS) of the International Space Station (ISS) is water based. The fluid in the ISS Laboratory Module and Node 1 initially contained a mix of water, phosphate (corrosion control), borate (pH buffer), and silver sulfate (Ag2SO4) (microbial control) at a pH of 9.5+/-0.5. Over time, the chemistry of the fluid changed. Fluid changes included a pH drop from 9.5 to 8.3 due to diffusion of carbon dioxide (CO2) through Teflon(reistered Trademark) (DuPont) hoses, increases in dissolved nickel (Ni) levels, deposition of silver (Ag) to metal surfaces, and precipitation of the phosphate (PO4) as nickel phosphate (NiPO4). The drop in pH and unavailability of a antimicrobial has provided an environment conducive to microbial growth. Microbial levels in the fluid have increased from 〉10 colony-forming units (CFUs)/100 ml to 10(exp 6) CFUs/100 ml. The heat exchangers in the IATCS loops are considered the weakest point in the loop because of the material thickness (=7 mil). It is made of a Ni-based braze filler/CRES 347. Results of a preliminary test performed at Hamilton Sundstrand indicated the possibility of pitting on this material at locations where Ag deposits were found. Later, tests have confirmed that chemical corrosion of the materials is a concern for this system. Accumulation of micro-organisms on surfaces (biofilm) can also result in material degradation and can amplify the damage caused by the chemical corrosion, known as microbiologically influenced corrosion (MIC). This paper will discuss the results of a 6-mo test performed to characterize and quantify the damage from microbial accumulation on the surface of the ISS/ATCS heat exchanger materials. The test was designed to quantify the damage to the materials under worst-case conditions with and without micro-organisms present at pH 8.3 and 9.5.

Description: This paper describes an application of a general purpose computer program, GFSSP (Generalized Fluid System Simulation Program) for calculating flow distribution in a network of micro-channels. GFSSP employs a finite volume formulation of mass and momentum conservation equations in a network consisting of nodes and branches. Mass conservation equation is solved for pressures at the nodes while the momentum conservation equation is solved at the branches to calculate flowrate. The system of equations describing the fluid network is solved by a numerical method that is a combination of the Newton-Raphson and successive substitution methods. The numerical results have been compared with test data and detailed CFD (computational Fluid Dynamics) calculations. The agreement between test data and predictions is satisfactory. The discrepancies between the predictions and test data can be attributed to the frictional correlation which does not include the effect of surface tension or electro-kinetic effect.

Description: Shadowgraph techniques were applied to the cold flow ejector facility at the Propulsion Research Center at the University of Alabama in Huntsville. The setup for the experiments was relatively simple and was accomplished at very little cost. Series of shadowgraph images were taken of both dual nozzle and single nozzle strut based ejectors operating over a range of chamber pressures. The density gradient patterns in the shadowgraphs were compared to pressure data measured along the top and side walls of the mixing duct. The shadowgraph images showed the presence of barrel shocks emanating from the nozzles which at low pressures terminated in Mach disks and at higher pressures extended beyond the barrel shape and reflected off the walls of the duct. Based on pressure data from previous testing, reflected shocks were expected on the walls of the duct. The shadowgraph images confirmed the locations of these reflected shocks on the top wall of the duct. The shadowgraph images also showed the structure change which correlated to a change in pitch of the ejector noise, and corresponded to a change in trend of the duct wall pressure ratio distributions. The images produced from the setup provided insight into the complex flow behavior inside the ejector duct. In addition, the techniques were a valuable tool as an educational device for students.

Description: The SAFE-I00a test article at the NASA Marshall Space Flight Center was used to simulate a variety of potential reactor transients; the SAFEl00a is a resistively heated, stainless-steel heat-pipe (HP)-reactor core segment, coupled to a gas-flow heat exchanger (HX). For these transients the core power was controlled by a point kinetics model with reactivity feedback based on core average temperature; the neutron generation time and the temperature feedback coefficient are provided as model inputs. This type of non-nuclear test is expected to provide reasonable approximation of reactor transient behavior because reactivity feedback is very simple in a compact fast reactor (simple, negative, and relatively monotonic temperature feedback, caused mostly by thermal expansion) and calculations show there are no significant reactivity effects associated with fluid in the HP (the worth of the entire inventory of Na in the core is .〈el, so fluid movement and temperature changes will cause very minor effects). In previous SAFE-100 tests, the point kinetics model was based on core thermal expansion via deflection measurements. It was found that core deflection was a strung function of how the SAFE-100 modules were fabricated and assembled (in terms of straightness, gaps, and other tolerances). To remove the added variable of how this particular core expands as compared to a different concept, it was decided to use a temperature based feedback model (based on several thermocouples placed throughout the core).

Description: 1. Kink dynamics. The first segment of a polygomized dislocation spiral step measured by AFM demonstrates up to 60% scattering in the critical length l*- the length when the segment starts to propagate. On orthorhombic lysozyme, this length is shorter than that the observed interkink distance. Step energy from the critical segment length based on the Gibbs-Thomson law (GTL), l* = 20(omega)alpha/(Delta)mu is several times larger than the energy from 2D nucleation rate. Here o is tine building block specific voiume, a is the step riser specific free energy, Delta(mu) is the crystallization driving force. These new data support our earlier assumption that the classical Frenkel, Burton -Cabrera-Frank concept of the abundant kink supply by fluctuations is not applicable for strongly polygonized steps. Step rate measurements on brushite confirms that statement. This is the1D nucleation of kinks that control step propagation. The GTL is valid only if l* 〈Dk/vk, the diffusion path of a kink that has diffusivity Dk and average growth velocity vk. This is equivalent to supersaturations sigma less than approx. alpha/2l*, where alpha is the building block size. For lysozyme, sigma much less than (1%). Conventionally used interstep distance generated by screw dislocation, 19(omega)alpha/Delta(mu) should be replaced by the very different real one, approx.4l*. 2. Stoichiometry. Kink, and thus step and face rates of a non-Kossel complex molecular monocomponent or any binary, AB, lattice was found theoretically to be proportional to 1/(zeta(sup 1/2) + zeta(sup - 1/2)), where zeta = [B]/[A] is the stoichiometry ratio in solution. The velocities reach maxima at zeta = 1. AFM studies of step rates on CaOxalate monohydrate (kidney stones) from aqueous solution was found to obey the law mentioned above. Generalization for more complex lattice will be discussed. 3. Turbulence. In agreement with theory, high precision in-situ laser interferometry of the (101) KDP crystal face shows step bunching if solution flows parallel to the step flow. The bunch height increases with the distance the bunch travels, i.e. with the face size. However, when the flow rate, u, increases, at u greater than approx. 1 m / s , the average step bunch height decreases as 1/u. The pheonomenon is attributed to the turbulent rather than laminar viscous boundary layer where diffusivity Dt = 0.5u(sub tau),y, i.e. increases linearly with the distance y from the solid face. Friction velocity, u(sub tau) approx. u(sup 7/8). Dramatically larger rate of the mass/heat transport within the turbulent, as compared to the laminar, viscous layer will be discussed.

Description: This paper details the author experiences with the validation of computer models to predict low gravity fluid behavior. It reviews the literature of low gravity fluid behavior as a starting point for developing a baseline set of test cases. It examines authors attempts to validate their models against these cases and the issues they encountered. The main issues seem to be that: Most of the data is described by empirical correlation rather than fundamental relation; Detailed measurements of the flow field have not been made; Free surface shapes are observed but through thick plastic cylinders, and therefore subject to a great deal of optical distortion; and Heat transfer process time constants are on the order of minutes to days but the zero-gravity time available has been only seconds.

Description: Packed bed reactors are well known for their vast and diverse applications in the chemical industry; from gas absorption, to stripping, to catalytic conversion. Use of this type of reactor in terrestrial applications has been rather extensive because of its simplicity and relative ease of operation. Developing similar reactors for use in microgravity is critical to many space-based advanced life support systems. However, the hydrodynamics of two-phase flow packed bed reactors in this new environment and the effects of one physiochemical process on another has not been adequately assessed. Surface tension or capillary forces play a much greater role which results in a shifting in flow regime transitions and pressure drop. Results from low gravity experiments related to flow regimes and two-phase pressure drop models are presented in this paper along with a description of plans for a flight experiment on the International Space Station (ISS). Understanding the packed bed hydrodynamics and its effects on mass transfer processes in microgravity is crucial for the design of packed bed chemical or biological reactors to be used for water reclamation and other life support processes involving water purification.

Description: To support development of a microgravity pressure control capability for liquid oxygen, thermodynamic vent system (TVS) testing was conducted at Marshall Space Flight Center (MSFC) using liquid nitrogen (LN2) as a LOX simulant. The spray bar TVS hardware used was originally designed by the Boeing Company for testing in liquid hydrogen (LH2). With this concept, a small portion of the tank fluid is passed through a Joule-Thomson (J-T) device, and then through a longitudinal spray bar mixed-heat exchanger in order to cool the bulk fluid. To accommodate the larger mass flow rates associated with LN2, the TVS hardware was modified by replacing the recirculation pump with an LN2 compatible pump and replacing the J-T valve. The primary advantage of the spray-bar configuration is that tank pressure control can be achieved independent of liquid and vapor location, enhancing the applicability of ground test data to microgravity conditions. Performance testing revealed that the spray-bar TVS was effective in controlling tank pressure within a 6.89 kPa band for fill levels of 90%, 50%, and 25%. Tests were also conducted with gaseous helium (GHe) in the ullage. The TVS operated nominally with GHe in the ullage, with performance similar to the tests with gaseous nitrogen (GN2). Testing demonstrated that the spray-bar TVS design was flexible enough for use in two different propellants with minimal hardware modifications.

Description: This paper presents a comparison of a single nozzle and a dual nozzle strut based ejector. The results are focused on the fluid properties in the ejector duct. The research focused on choking mechanisms, mass flow entrainment, and mixing duct pressure distributions. The two ejectors were tests at equivalent primary mass flow rates. This corresponds to chamber pressures ranging from 100 psi to 900 psi in the single nozzle strut and 50 psi to 450 psi in the dual nozzle strut. Secondary flow was drawn from the lab at atmospheric pressure, and was not controlled. The secondary flow was found to choke at a value of 2.3 lb/s for a primary mass flow rate at approximately 2.1 lb/s for both ejectors. This choke was believed to be a mass addition choke rather than a traditional aerodynamic choke. The mixing duct pressure distribution exhibited two distinct trends at "low pressure" trend and at "high pressure" trend. For the low pressure trend, the mixing length for the ejectors remained fixed around 20 inches, regardless of the chamber pressure. For the higher pressure trend, the mixing length was considerably longer and increased with increasing chamber pressure. At high chamber pressures (high mass flow rates), a supersonic core flow was present at the exit of the duct. For these cases, the two streams did not have time to mix by the end of the duct.

Description: Candidate passive flow control devices were chosen from a NASA flow visualization study to investigate their effectiveness at improving flow quality over a flat-top carrier model. Flow over the deck was analyzed using a particle image velocimeter and a 1/120th scaled carrier model in a low-speed wind tunnel. Baseline (no devices) flow quality was compared to flow quality from combinations of bow and deck-edge devices at both zero and 20 degrees yaw. Devices included plain flaps and spiral cross-section columnar vortex generators attached in various combinations to the front and sides of the deck. Centerline and cross plane measurements were made with velocity and average turbulence measurements reported. Results show that the bow/deck-edge flap and bow/deck-edge columnar vortex generator pairs reduce flight deck turbulence both at zero yaw and at 20 degrees yaw by a factor of approximately 20. Of the devices tested, the most effective bow-only device appears to be the plain flap.

Description: The pulsed flow emitted from a shrouded Hartmann-Sprenger tube was sampled with high-frequency pressure transducers and with laser particle imaging velocimetry, and found to consist of a train of vortices. Thrust and mass flow were also monitored using a thrust plate and orifice, respectively. The tube and shroud lengths were altered to give four different operating frequencies. From the data, the radius, velocity, and circulation of the vortex rings was obtained. Each frequency corresponded to a different length to diameter ratio of the pulse of air leaving the driver shroud. Two of the frequencies had length to diameter ratios below the formation number, and two above. The formation number is the value of length to diameter ratio below which the pulse converts to a vortex ring only, and above which the pulse becomes a vortex ring plus a trailing jet. A modified version of the slug model of vortex ring formation was used to compare the observations with calculated values. Because the flow exit area is an annulus, vorticity is shed at both the inner and outer edge of the jet. This results in a reduced circulation compared with the value calculated from slug theory accounting only for the outer edge. If the value of circulation obtained from laser particle imaging velocimetry is used in the slug model calculation of vortex ring velocity, the agreement is quite good. The vortex ring radius, which does not depend on the circulation, agrees well with predictions from the slug model.

Description: The program objectives are fully defined in the original proposal entitled Program of Research in Flight Dynamics in GW at NASA Langley Research Center, which was originated March 20, 1975, and in the renewals of the research program from January 1, 2003 to September 30, 2005. The program in its present form includes three major topics: 1. the improvement of existing methods and development of new methods for wind tunnel and flight data analysis, 2. the application of these methods to wind tunnel and flight test data obtained from advanced airplanes, 3. the correlation of flight results with wind tunnel measurements, and theoretical predictions.

Description: A computational investigation of a two-dimensional nozzle was completed to assess the use of fluidic injection to manipulate flow separation and cause thrust vectoring of the primary jet thrust. The nozzle was designed with a recessed cavity to enhance the throat shifting method of fluidic thrust vectoring. Several design cycles with the structured-grid, computational fluid dynamics code PAB3D and with experiments in the NASA Langley Research Center Jet Exit Test Facility have been completed to guide the nozzle design and analyze performance. This paper presents computational results on potential design improvements for best experimental configuration tested to date. Nozzle design variables included cavity divergence angle, cavity convergence angle and upstream throat height. Pulsed fluidic injection was also investigated for its ability to decrease mass flow requirements. Internal nozzle performance (wind-off conditions) and thrust vector angles were computed for several configurations over a range of nozzle pressure ratios from 2 to 7, with the fluidic injection flow rate equal to 3 percent of the primary flow rate. Computational results indicate that increasing cavity divergence angle beyond 10 is detrimental to thrust vectoring efficiency, while increasing cavity convergence angle from 20 to 30 improves thrust vectoring efficiency at nozzle pressure ratios greater than 2, albeit at the expense of discharge coefficient. Pulsed injection was no more efficient than steady injection for the Dual Throat Nozzle concept.

Description: A new porous condition has been implemented in the PAB3D solver for simulating the flow over porous surfaces. The newly-added boundary condition is utilized to compute the flow field of a non-axisymmetric, convergent-divergent nozzle incorporating porous cavities for shock-boundary layer interaction control. The nozzle has an expansion ratio (exit area/throat area) of 1.797 and a design nozzle pressure ratio of 8.78. The flow fields for a baseline nozzle (no porosity) and for a nozzle with porous surfaces (10% porosity ratio) are computed for NPR varying from 2.01 to 9.54. Computational model results indicate that the over-expanded nozzle flow was dominated by shock-induced boundary-layer separation. Porous configurations were capable of controlling off-design separation in the nozzle by encouraging stable separation of the exhaust flow. Computational simulation results, wall centerline pressure, mach contours, and thrust efficiency ratio are presented and discussed. Computed results are in excellent agreement with experimental data.

Description: Computational aerothermodynamic simulations of Orbiter windside tile damage in flight were performed in support of the Space Shuttle Return-to-Flight effort. The simulations were performed for both hypervelocity flight and low-enthalpy wind tunnel conditions and contributed to the Return-to-Flight program by providing information to support a variety of damage scenario analyses. Computations at flight conditions were performed at or very near the peak heating trajectory point for multiple damage scenarios involving damage windside acreage reaction cured glass (RCG) coated silica tile(s). The cavities formed by the missing tile examined in this study were relatively short leading to flow features which indicated open cavity behavior. Results of the computations indicated elevated heating bump factor levels predicted for flight over the predictions for wind tunnel conditions. The peak heating bump factors, defined as the local heating to a reference value upstream of the cavity, on the cavity floor for flight simulation were 67% larger than the peak wind tunnel simulation value. On the downstream face of the cavity the flight simulation values were 60% larger than the wind tunnel simulation values. On the outer mold line (OML) downstream of the cavity, the flight values are about 20% larger than the wind tunnel simulation values. The higher heating bump factors observed in the flight simulations were due to the larger driving potential in terms of energy entering the cavity for the flight simulations. This is evidenced by the larger rate of increase in the total enthalpy through the boundary layer prior to the cavity for the flight simulation.

Description: An axisymmetric phase field model is developed and used to model surface tension forces on liquid jets in microgravity. The previous work in this area is reviewed and a baseline drop tower experiment selected for model comparison. This paper uses the model to parametrically investigate the influence of key parameters on the geysers formed by jets in microgravity. Investigation of the contact angle showed the expected trend of increasing contact angle increasing geyser height. Investigation of the tank radius showed some interesting effects and demonstrated the zone of free surface deformation is quite large. Variation of the surface tension with a laminar jet showed clearly the evolution of free surface shape with Weber number. It predicted a breakthrough Weber number of 1.

Description: A detailed analysis of the liquid film characteristics and the accompanying heat transfer of a free surface controlled liquid impinging jet onto a rotating disk are presented. The computations were run on a two-dimensional axi-symmetric Eulerian mesh while the free surface was calculated with the volume of fluid method. Flow rates between 3 and 15 1pm with rotational speeds between 50 and 200 rpm are analyzed. The effects of inlet temperature on the film thickness and heat transfer are characterized as well as evaporative effects. The conjugate heating effect is modeled, and was found to effect the heat transfer results the most at both the inner and outer edges of the heated surface. The heat transfer was enhanced with both increasing flow rate and increasing rotational speeds. When evaporative effects were modeled, the evaporation was found to increase the heat transfer at the lower flow rates the most because of a fully developed thermal field that was achieved. The evaporative effects did not significantly enhance the heat transfer at the higher flow rates.

Description: Arc jet tests of candidate tile repair materials and baseline Orbiter uncoated reusable surface insulation (RSI) were performed in the Johnson Space Center's (JSC) Atmospheric Reentry Materials and Structures Evaluation Facility (ARMSEF) from June 23, 2003, through August 19, 2003. These tests were performed to screen candidate tile repair materials by verifying the high temperature performance and determining the thermal stability. In addition, tests to determine the surface emissivity at high temperatures and the geometric shrinkage of bare RSI were performed. In addition, tests were performed to determine the surface emissivity at high temperatures and the geometric shrinkage of uncoated RSI.

Description: Results from unsteady Reynolds-averaged Navier-Stokes computations are described for two different synthetic jet flows issuing into a turbulent boundary layer crossflow through a circular orifice. In one case the jet effect is mostly contained within the boundary layer, while in the other case the jet effect extends beyond the boundary layer edge. Both cases have momentum flux ratios less than 2. Several numerical parameters are investigated, and some lessons learned regarding the CFD methods for computing these types of flow fields are outlined. Results in both cases are compared to experiment.

Description: An analysis of the flow state on a trapezoidal wing model from the NASA 3-D High Lift Flow Physics Experiment is presented. The objective of the experiment was to characterize the flow over a non-proprietary semi-span three-element high-lift configuration to aid in assessing the state of the art in the computation of three-dimensional high-lift flows. Surface pressures and hot-film sensors are used to determine the flow conditions on the slat, main, and flap. The locations of the attachments lines and the values of the attachment line Reynolds number are estimated based on the model surface pressures. Data from the hot-films are used to determine if the flow is laminar, transitional, or turbulent by examining the hot-film time histories, statistics, and frequency spectra.

Description: Boundary layer receptivity to two-dimensional slow and fast acoustic waves is investigated by solving Navier-Stokes equations for Mach 4.5 flow over a flat plate with a finite-thickness leading edge. Higher order spatial and temporal schemes are employed to obtain the solution whereby the flat-plate leading edge region is resolved by providing a sufficiently refined grid. The results show that the instability waves are generated in the leading edge region and that the boundary-layer is much more receptive to slow acoustic waves (by almost a factor of 20) as compared to the fast waves. Hence, this leading-edge receptivity mechanism is expected to be more relevant in the transition process for high Mach number flows. The effect of acoustic wave incidence angle is also studied and it is found that the receptivity of the boundary layer on the windward side (with respect to the acoustic forcing) decreases by more than a factor of 4 when the incidence angle is increased from 0 to 45 deg. However, the receptivity coefficient for the leeward side is found to vary relatively weakly with the incidence angle. The effect of leading-edge thickness is also studied and bluntness is found to stabilize the boundary layer. The relative significance of fast acoustic waves is enhanced in the presence of bluntness.

Description: A static and dynamic Chimera overset-grid capability is added to an established NASA tetrahedral unstructured parallel Navier-Stokes flow solver, USM3D. Modifications to the solver primarily consist of a few strategic calls to the Donor interpolation Receptor Transaction library (DiRTlib) to facilitate communication of solution information between various grids. The assembly of multiple overlapping grids into a single-zone composite grid is performed by the Structured, Unstructured and Generalized Grid AssembleR (SUGGAR) code. Several test cases are presented to verify the implementation, assess overset-grid solution accuracy and convergence relative to single-grid solutions, and demonstrate the prescribed relative grid motion capability.

Description: We apply an unsteady Reynolds-averaged Navier-Stokes (URANS) solver for unstructured grids to unsteady flows on moving and stationary grids. Example problems considered are relevant to active flow control and stability and control. Computational results are presented using the Spalart-Allmaras turbulence model and are compared to experimental data. The effect of grid and time-step refinement are examined.

Description: Discrepancies between experiment and computation for shuttle leeside flow separation, which came to light in the Columbia accident investigation, are resolved. Tests were run in the Langley Research Center 20-Inch Hypersonic CF4 Tunnel with a baseline orbiter model and two extended trailing edge models. The extended trailing edges altered the wing leeside separation lines, moving the lines toward the fuselage, proving that wing trailing edge modeling does affect the orbiter leeside flow. Computations were then made with a wake grid. These calculations more closely matched baseline experiments. Thus, the present findings demonstrate that it is imperative to include the wake flow domain in CFD calculations in order to accurately predict leeside flow separation for hypersonic vehicles at high angles of attack.

Description: WIND, Fluent, and PowerFLOW were used to predict the lift, drag, and moment coefficients of a business-jet airfoil with a rime ice (rough and jagged, but no protruding horns) and with a glaze ice (rough and jagged end has two or more protruding horns) for angles of attack from zero to and after stall. The performance of the following turbulence models were examined by comparing predictions with available experimental data. Spalart-Allmaras (S-A), RNG k-epsilon, shear-stress transport, v(sup 2)-f, and a differential Reynolds stress model with and without non-equilibrium wall functions. For steady RANS simulations, WIND and FLUENT were found to give nearly identical results if the grid about the iced airfoil, the turbulence model, and the order of accuracy of the numerical schemes used are the same. The use of wall functions was found to be acceptable for the rime ice configuration and the flow conditions examined. For rime ice, the S-A model was found to predict accurately until near the stall angle. For glaze ice, the CFD predictions were much less satisfactory for all turbulence models and codes investigated because of the large separated region produced by the horns. For unsteady RANS, WIND and FLUENT did not provide better results. PowerFLOW, based on the Lattice Boltzmann method, gave excellent results for the lift coefficient at and near stall for the rime ice, where the flow is inherently unsteady.