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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS1990

Rapid mix concepts for low emission combustors in gas turbine engines

NASA LeRC has identified the Rich burn/Quick mix/Lean burn (RQL) combustor as a potential gas turbine combustor concept to reduce NOx emissions in High Speed Civil Transport (HSCT) aircraft. To demonstrate reduced NOx levels, NASA LeRC soon will test a flametube version of an RQL combustor. The critical technology needed for the RQL combustor is a method of quickly mixing combustion air with rich burn gases. Two concepts were proposed to enhance jet mixing in a circular cross-section: the Asymmetric Jet Penetration (AJP) concept; and the Lobed Mixer (LM) concept. In Phase 1, two preliminary configurations of the AJP concept were compared with a conventional 12-jet radial-inflow slot design. The configurations were screened using an advanced 3-D Computational Fluid Dynamics (CFD) code named REFLEQS. Both non-reacting and reacting analyses were performed. For an objective comparison, the conventional design was optimized by parametric variation of the jet-to-mainstream momentum flux (J) ratio. The optimum J was then employed in the AJP simulations. Results showed that the three-jet AJP configuration was superior in overall mixedness compared to the conventional design. However, in regards to NOx emissions, the AJP configuration was inferior. The higher emission level for AJP was caused by a single hot spot located in the wake of the central jet as it entered the combustor. Ways of maintaining good mixedness while eliminating the hot spot were identified for Phase 2 study. Overall, Phase 1 showed the viability of using CFD analyses to evaluate quick-mix concepts. A high probability exists that advancing mixing concepts will reduce NOx emissions in RQL combustors, and should be explored in Phase 2, by parallel numerical and experimental work.

Talpallikar, Milind V.↗
NASA NTRS2014

Assessment of Reduced-Kinetics Mechanisms for Combustion of Jet Fuel in CFD Applications

A computational effort was undertaken to analyze the details of fluid flow in Lean-Direct Injection (LDI) combustors for next-generation LDI design. The National Combustor Code (NCC) was used to perform reacting flow computations on single-element LDI injector configurations. The feasibility of using a reduced chemical-kinetics approach, which optimizes the reaction rates and species to model the emissions characteristics typical of lean-burning gas-turbine combustors, was assessed. The assessments were performed with Reynolds- Averaged Navier-Stokes (RANS) and Time-Filtered Navier Stokes (TFNS) time-integration, with a Lagrangian spray model with the NCC code. The NCC predictions for EINOx and combustor exit temperature were compared with experimental data for two different single-element LDI injector configurations, with 60deg and 45deg axially swept swirler vanes. The effects of turbulence-chemistry interaction on the predicted flow in a typical LDI combustor were studied with detailed comparisons of NCC TFNS with experimental data.

Ajmani, Kumud↗
NASA NTRS2012

Magnetic Field Effects on Plasma Plumes

Here, we will discuss our numerical studies of plasma jets and loops, of basic interest for plasma propulsion and plasma astrophysics. Space plasma propulsion systems require strong guiding magnetic fields known as magnetic nozzles to control plasma flow and produce thrust. Propulsion methods currently being developed that require magnetic nozzles include the VAriable Specific Impulse Magnetoplasma Rocket (VASIMR) [1] and magnetoplasmadynamic thrusters. Magnetic nozzles are functionally similar to de Laval nozzles, but are inherently more complex due to electromagnetic field interactions. The two crucial physical phenomenon are thrust production and plasma detachment. Thrust production encompasses the energy conversion within the nozzle and momentum transfer to a spacecraft. Plasma detachment through magnetic reconnection addresses the problem of the fluid separating efficiently from the magnetic field lines to produce maximum thrust. Plasma jets similar to those of VASIMR will be studied with particular interest in dual jet configurations, which begin as a plasma loops between two nozzles. This research strives to fulfill a need for computational study of these systems and should culminate with a greater understanding of the crucial physics of magnetic nozzles with dual jet plasma thrusters, as well as astrophysics problems such as magnetic reconnection and dynamics of coronal loops.[2] To study this problem a novel, hybrid kinetic theory and single fluid magnetohydrodynamic (MHD) solver known as the Magneto-Gas Kinetic Method is used.[3] The solver is comprised of a "hydrodynamic" portion based on the Gas Kinetic Method and a "magnetic" portion that accounts for the electromagnetic behaviour of the fluid through source terms based on the resistive MHD equations. This method is being further developed to include additional physics such as the Hall effect. Here, we will discuss the current level of code development, as well as numerical simulation results

Ebersohn, F.↗
NASA NTRS2017

Computational Simulations of the NASA Langley HyMETS Arc-Jet Facility

The Hypersonic Materials Environmental Test System (HyMETS) arc-jet facility located at the NASA Langley Research Center in Hampton, Virginia, is primarily used for the research, development, and evaluation of high-temperature thermal protection systems for hypersonic vehicles and reentry systems. In order to improve testing capabilities and knowledge of the test article environment, an effort is underway to computationally simulate the flow-field using computational fluid dynamics (CFD). A detailed three-dimensional model of the arc-jet nozzle and free-jet portion of the flow-field has been developed and compared to calibration probe Pitot pressure and stagnation-point heat flux for three test conditions at low, medium, and high enthalpy. The CFD model takes into account uniform pressure and non-uniform enthalpy profiles at the nozzle inlet as well as catalytic recombination efficiency effects at the probe surface. Comparing the CFD results and test data indicates an effectively fully-catalytic copper surface on the heat flux probe of about 10% efficiency and a 2-3 kpa pressure drop from the arc heater bore, where the pressure is measured, to the plenum section, prior to the nozzle. With these assumptions, the CFD results are well within the uncertainty of the stagnation pressure and heat flux measurements. The conditions at the nozzle exit were also compared with radial and axial velocimetry. This simulation capability will be used to evaluate various three-dimensional models that are tested in the HyMETS facility. An end-to-end aerothermal and thermal simulation of HyMETS test articles will follow this work to provide a better understanding of the test environment, test results, and to aid in test planning. Additional flow-field diagnostic measurements will also be considered to improve the modeling capability.

Brune, A. J.↗
NASA NTRS2002

Thermal Protection System Evaluation Using Arc-jet Flows: Flight Simulation or Research Tool?

The arc-jet has been used to evaluate thermal protection systems (TPS) and materials for the past forty years. Systems that have been studied in this environmerd include ablators, active, and passive TPS concepts designed for vehicles entering planetary and Earth atmospheres. The question of whether arc-jet flow can simulate a flight environment or is it a research tool that provides an aero-thermodynamic heating environment to obtain critical material properties will be addressed. Stagnation point tests in arc-jets are commonly used to obtain material properties such as mass loss rates, thermal chemical stability data, optical properties, and surface catalytic efficiency. These properties are required in computational fluid dynamic codes to accurately predict the performance of a TPS during flight. Special facilities have been developed at NASA Ames Research Center to approximate the flow environment over the mid-fuselage and body flap regions of proposed space-planes type vehicles. This paper compares flow environments generated in flight over a vehicle with those created over an arc-jet test articles in terms of scale, chemistry, and fluid dynamic properties. Flight experiments are essential in order to validate the material properties obtained from arc-jet tests and used to predict flight performance of any TPS being considered for use on a vehicle entering the Earth atmosphere at hypersonic speed.

Stewart, David A.↗
NASA NTRS2009

Dynamic Measurement of Temperature, Velocity, and Density in Hot Jets Using Rayleigh Scattering

A molecular Rayleigh scattering technique is utilized to measure gas temperature, velocity, and density in unseeded gas flows at sampling rates up to 10 kHz, providing fluctuation information up to 5 kHz based on the Nyquist theorem. A high-power continuous-wave laser beam is focused at a point in an air flow field and Rayleigh scattered light is collected and fiber-optically transmitted to a Fabry-Perot interferometer for spectral analysis. Photomultiplier tubes operated in the photon counting mode allow high-frequency sampling of the total signal level and the circular interference pattern to provide dynamic density, temperature, and velocity measurements. Mean and root mean square velocity, temperature, and density, as well as power spectral density calculations, are presented for measurements in a hydrogen-combustor heated jet facility with a 50.8-mm diameter nozzle at NASA John H. Glenn Research Center at Lewis Field. The Rayleigh measurements are compared with particle image velocimetry data and computational fluid dynamics predictions. This technique is aimed at aeronautics research related to identifying noise sources in free jets, as well as applications in supersonic and hypersonic flows where measurement of flow properties, including mass flux, is required in the presence of shocks and ionization occurrence.

Mielke, Amy F.↗
NASA NTRS1994

Experimental Investigation of Jet-Induced Mixing of a Large Liquid Hydrogen Storage Tank

Experiments have been conducted to investigate the effect of fluid mixing on the depressurization of a large liquid hydrogen storage tank. The test tank is approximately ellipsoidal, having a volume of 4.89 m(exp 3) and an average wall heat flux of 4.2 W/m(exp 2) due to external heat input. A mixer unit was installed near the bottom of the tank to generate an upward directed axial jet flow normal to the liquid-vapor interface. Mixing tests were initiated after achieving thermally stratified conditions in the tank either by the introduction of hydrogen gas into the tank or by self-pressurization due to ambient heat leak through the tank wall. The subcooled liquid jet directed towards the liquid-vapor interface by the mixer induced vapor condensation and caused a reduction in tank pressure. Tests were conducted at two jet submergence depths for jet Reynolds numbers from 80,000 to 495,000 and Richardson numbers from 0.014 to 0.52. Results show that the rate of tank pressure change is controlled by the competing effects of subcooled jet flow and the free convection boundary layer flow due to external tank wall heating. It is shown that existing correlations for mixing time and vapor condensation rate based on small scale tanks may not be applicable to large scale liquid hydrogen systems.

Lin, C. S.↗
NASA NTRS1978

Some flow phenomena in a constant area duct with a Borda type inlet including the critical region

Mass limiting flow characteristics for a 55 L/D tube with a Borda type inlet were assessed over large ranges of temperature and pressure, using fluid nitrogen. Under certain conditions, separation and pressure drop at the inlet was sufficiently strong to permit partial vaporization and the remaining fluid flowed through the tube as if it were a free jet. An empirical relation was determined which defines conditions under which this type of flow can occur. A flow coefficient is presented which enables estimations of flow rates over the experimental range. A flow rate stagnation pressure map for selected stagnation isotherms and pressure profiles document these flow phenomena.

Hendricks, R. C.↗
NASA NTRS1978

Some flow phenomena in a constant area duct with a Borda type inlet including the critical region

Mass limiting flow characteristics for a 55 L/D tube with a Borda type inlet have been assessed over large ranges of temperature and pressure using fluid nitrogen. Under certain conditions, separation and pressure drop at the inlet was sufficiently strong to permit partial vaporization and the remaining fluid flowed through the tube as if it were a free jet. An empirical relation was determined which defines conditions under which this type of flow can occur. A flow coefficient is presented which enables one to estimate flow rates over the experimental range. A flow rate stagnation pressure map for selected stagnation isotherms and pressure profiles document these flow phenomena.

Hendricks, R. C.↗
NASA NTRS1986

Detonation wave compression in gas turbines

A study was made of the concept of augmenting the performance of low pressure ratio gas turbines by detonation wave compression of part of the flow. The concept exploits the constant volume heat release of detonation waves to increase the efficiency of the Brayton cycle. In the models studied, a fraction of the compressor output was channeled into detonation ducts where it was processed by transient transverse detonation waves. Gas dynamic studies determined the maximum cycling frequency of detonation ducts, proved that upstream propagation of pressure pulses represented no problems and determined the variations of detonation duct output with time. Mixing and wave compression were used to recombine the combustor and detonation duct flows and a concept for a spiral collector to further smooth the pressure and temperature pulses was presented as an optional component. The best performance was obtained with a single firing of the ducts so that the flow could be re-established before the next detonation was initiated. At the optimum conditions of maximum frequency of the detonation ducts, the gas turbine efficiency was found to be 45 percent while that of a corresponding pressure ratio 5 conventional gas turbine was only 26%. Comparable improvements in specific fuel consumption data were found for gas turbines operating as jet engines, turbofans, and shaft output machines. Direct use of the detonation duct output for jet propulsion proved unsatisfactory. Careful analysis of the models of the fluid flow phenomena led to the conclusion that even more elaborate calculations would not diminish the uncertainties in the analysis of the system. Feasibility of the concept to work as an engine now requires validation in an engineering laboratory experiment.

Wortman, A.↗
NASA NTRS1994

Measurement of a surface heat flux and temperature

The Heat Flux Microsensor is a new sensor which was recently patented by Virginia Tech and is just starting to be marketed by Vatell Corp. The sensor is made using the thin-film microfabrication techniques directly on the material that is to be measured. It consists of several thin-film layers forming a differential thermopile across a thermal resistance layer. The measured heat flux q is proportional to the temperature difference across the resistance layer q= k(sub g)/delta(sub g) x (t(sub 1) - T(sub 2)), where k(sub g) is the thermal conductivity and delta (sub g) is the thickness of the thermal resistance layer. Because the gages are sputter coated directly onto the surface, their total thickness is less than 2 micrometers, which is two orders of magnitude thinner than previous gages. The resulting temperature difference across the thermal resistance layer (delta is less than 1 micrometer) is very small even at high heat fluxes. To generate a measurable signal many thermocouple pairs are put in series to form a differential thermopile. The combination of series thermocouple junctions and thin-film design creates a gage with very attractive characteristics. It is not only physically non-intrusive to the flow, but also causes minimal disruption of the surface temperature. Because it is so thin, the response time is less than 20 microsec. Consequently, the frequency response is flat from 0 to over 50 kHz. Moreover, the signal of the Heat Flux Microsensor is directly proportional to the heat flux. Therefore, it can easily be used in both steady and transient flows, and it measures both the steady and unsteady components of the surface heat flux. A version of the Heat Flux Microsensor has been developed to meet the harsh demands of combustion environments. These gages use platinum and platinum-10 percent rhodium as the thermoelectric materials. The thermal resistance layer is silicon monoxide and a protective coating of Al2O3 is deposited on top of the sensor. The superimposed thin-film pattern of all six layers is presented. The large pads are for connection with pins used to bring the signal out the back of the ceramic. In addition to the heat flux measurement, the surface temperature is measured with a platinum resistance layer (RTS). The resistance of this layer increases with increasing temperature. Therefore, these gages simultaneously measure the surface temperature and heat flux. The demonstrated applications include rocket nozzles, SCRAM jet engines, gas turbine engines, boiling heat transfer, flame experiments, basic fluid heat transfer, hypersonic flight, and shock tube testing. The laboratory involves using one of these sensors in a small combustion flame. The sensor is made on a 2.5 cm diameter piece of aluminum nitride ceramic.

Davis, R. M.↗
NASA NTRS1992

Precision Heating Process

A heat sealing process was developed by SEBRA based on technology that originated in work with NASA's Jet Propulsion Laboratory. The project involved connecting and transferring blood and fluids between sterile plastic containers while maintaining a closed system. SEBRA markets the PIRF Process to manufacturers of medical catheters. It is a precisely controlled method of heating thermoplastic materials in a mold to form or weld catheters and other products. The process offers advantages in fast, precise welding or shape forming of catheters as well as applications in a variety of other industries.

Source record↗
NASA NTRS2017

Elastomeric Thermal Insulation Design Considerations in Long, Aluminized Solid Rocket Motors

An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.

Martin, Heath T.↗
NASA NTRS2017

A Comparison of System Architectures for a Mechanically Pumped Two-Phase Thermal Control System

The NASA Jet Propulsion Laboratory is developing a mechanically pumped two-phase fluid loop thermal control system to enable novel mission designs and greater science return. Pumped two-phase fluid loops have the potential to provide robust and effective thermal control that combine the best aspects of passive two-phase systems (heat pipes) and mechanically pumped single-phase fluid loops. The current program requirements include the development of a system with multiple 1 sq.m evaporators, each of which is capable of remaining spatially and temporally isothermal while accommodating heat loads of up to 500 W and local fluxes of up to 5 W/sq.cm. The goal is to attain this using less than 5 W of power. Such a system would be able to accommodate the next generation of payload and bus electronics while using minimal resources. This paper compares two different mechanically pumped two-phase fluid loop architectures in the context of these requirements. A mixed flow and separated flow architecture are compared on a theoretical and experimental basis. Test data from sub-scale, single evaporator/single condenser, mixed flow and separated flow testbeds are presented. In addition, a model is introduced to better understand separated flow systems and some expressions for the theoretical performance limits of such systems are developed. To date, the investigation suggests that a separated flow architecture is better suited to the program requirements. Separated flow systems have the potential to accommodate an isothermalizing two-phase evaporator while using lower levels of power than would be required for a mixed flow system. In addition, it is argued that separated flow systems are more robust and amenable to analysis than mixed flow systems, since they significantly reduce the occurrence of two-phase flow by separating phases in the evaporator. Future work will include developing a full-scale testbed that includes multiple evaporators and condensers in a representative flight configuration.

Furst, Benjamin↗
NASA NTRS2012

Measurements and Simulations of Surface Dielectric Barrier Discharges Used as Plasma Actuators

This report is a Ph.D. dissertation performed under NRA cooperative agreement and submitted as part of the final report. Asymmetric surface dielectric barrier discharges (DBDs) have shown promise for use as aerodynamic actuators for active flow control. In this project we studied DBD actuators experimentally and numerically. Our DBDs used a symmetric triangular high voltage waveform to generate plasma in atmospheric pressure air. Time-averaged measurements indicated that the induced force of a single barrier actuator design (one electrode insulated from the plasma) can be increased exponentially above the results of previous studies by decreasing both the length and thickness of the electrode exposed to the plasma. This increased force may allow these devices to control flow separation in a wider range of flow environments. Experiments using an intensified digital camera to examine the plasma on time scales of a few nanoseconds showed that, in addition to the previously-observed filamentary and jet-like plasma structures, discharges with very thin exposed electrodes exhibited a weak but constant plasma immediately adjacent to those electrodes. In double-barrier actuators (both electrodes insulated), decreasing the diameter of the narrower electrode lead to increasing forces, and recorded images showed the simultaneous existence of both filamentary and jet-like plasma structures. The development and application of a time-dependent, two-dimensional computational fluid plasma model has aided in understanding the detailed physics of surface DBDs at all-time scales. For simulated single-barrier discharges, the model qualitatively reproduced the filamentary and jet-like micro-discharge structures. The model was somewhat successful in reproducing the observed characteristics of double-barrier actuators. For both actuator geometries, the model indicated that the majority of the forces induced on the neutral gas occur in between micro-discharges as the plasmas decay.

Hoskinson, Alan R.↗
NASA NTRS2008

Jet Penetration into a Scaled Microfabricated Stirling Cycle Regenerator

The cooler and heater adjacent to the regenerator of a Stirling cycle engine have tubes or channels which form jets that pass into the regenerator while diffusing within the matrix. An inactive part of the matrix, beyond the cores of these jets, does not participate fully in the heat transfer between the flow of working fluid and the regenerator matrix material, weakening the regenerator s ability to exchange heat with the working fluid. The objective of the present program is to document this effect on the performance of the regenerator and to develop a model for generalizing the results. However, the small scales of actual Stirling regenerator matrices (on the order of tens of microns) make direct measurements of this effect very difficult. As a result, jet spreading within a regenerator matrix has not been characterized well and is poorly understood. Also, modeling is lacking experimental verification. To address this, a large-scale mockup of thirty times actual scale was constructed and operated under conditions that are dynamically similar to the engine operation. Jet penetration with round jets and slot jets into the microfabricated regenerator geometry are then measured by conventional means. The results are compared with those from a study of spreading of round jets within woven screen regenerator for further documentation of the comparative performance of the microfabricated regenerator geometry.

Sun, Liyong↗
NASA NTRS2000

Advanced Methods for Aircraft Engine Thrust and Noise Benefits: Nozzle-Inlet Flow Analysis

In this project, we continue to develop the previous joint research between the Fluid Mechanics and Acoustics Laboratory (FM&AL) at Hampton University (HU) and the Jet Noise Team (JNT) at the NASA Langley Research Center (NASA LaRC). The FM&AL was established at Hampton University in June of 1996 and has conducted research under two NASA grants: NAG-1-1835 (1996-99), and NAG-1-1936 (1997-00). In addition, the FM&AL has jointly conducted research with the Central AeroHydrodynamics Institute (TsAGI, Moscow) in Russia under a Civilian Research and Development Foundation (CRDF) grant #RE2-136 (1996-99). The goals of the FM&AL programs are twofold: (1) to improve the working efficiency of the FM&AUs team in generating new innovative ideas and in conducting research in the field of fluid dynamics and acoustics, basically for improvement of supersonic and subsonic aircraft engines, and (2) to attract promising minority students to this research and training and, in cooperation with other HU departments, to teach them basic knowledge in Aerodynamics, Gas Dynamics, and Theoretical and Experimental Methods in Aeroacoustics and Computational Fluid Dynamics (CFD). The research at the HU FM&AL supports reduction schemes associated with the emission of engine pollutants for commercial aircraft and concepts for reduction of observables for military aircraft. These research endeavors relate to the goals of the NASA Strategic Enterprise in Aeronautics concerning the development of environmentally acceptable aircraft. It is in this precise area, where the US aircraft industry, academia, and Government are in great need of trained professionals and which is a high priority goal of the Minority University Research and Education (MUREP) Program, that the HU FM&AL can make its most important contribution. The main achievements for the reporting period in the development of concepts for noise reduction and improvement in efficiency for jet exhaust nozzles and inlets for aircraft engines are as follows: (1) Publications- The AIAA Paper #99-1924 has been presented at the 5th AIAA/CEAS Aeroacoustics Conference, May 10-12, 1999, Seattle, WA; the AIAA Paper #00-3315 has been accepted for the 36th AIAA/ASME/ SAE/ASEE Joint Propulsion Conference, 17-19 July, 2000, Huntsville, AL; and another paper has been accepted for the International Environmental Congress, 14-16 June, 2000, St.-Petersburg, Russia. (2) Two patents were granted on July 20, 1999, and January 12, 2000. (3) Three reports/presentations at the NASA LaRC and GRC (06/22199, 09/26/ 99, and 06/25/00). (4) Grants and Proposals: Four proposals were submitted to the NASA and CRDF; a NASA Faculty Award was granted on January, 2000. A CRDF Young Investigator Program Award was granted for a 3 months visit of the Russian scientist to the HU FM&AL (03/99-05/99). (5) Theory and Numerical Simulations- Analytical theory, numerical simulation, comparison of theoretical with experimental results, and modification of theoretical approaches, models, grids etc. have been conducted for several complicated 2D and 3D nozzle and inlet designs using NASA codes based on full Euler and Navier-Stokes solvers: CFL3D, CRAFT, GODUNOV, and others. New approach for environmental monitoring via infrasound. (6) Experimental Tests: Experimental acoustic tests at the TsAGI, Moscow, with nozzles having Screwdriver or Axisymmetric Plug and Permeable Shells. A small scale working model of the NASA Low Speed Wind Tunnel (LSWT) has been installed in the Experimental Hall of the HU FM&AL (June, 1999). Preliminary preparations for experimental tests were made. (7) Students Research Activity: Involvement of the two graduate students as research assistants in the current research project.

Gilinsky, Mikhail↗
NASA NTRS2002

Progress Toward Generation of a Navier-Stokes Database for a Harrier in Ground Effect

The Harrier YAV-8B aircraft is capable of vertical and short-field take-off and landing (V/STOL) by directing its four exhaust nozzles toward the ground, or conventional flight by rotating its nozzles into a horizontal position. The British Royal Air Force and the United States Marine Corps have used this aircraft for more than 30 years to provide a quick reaction time for troop support, and reduce the need for long runways. The success of this powered-lift (PL) vehicle has also prompted the more recent design of the Joint Strike Fighter (JSF). However there are significant safety issues that must be addressed when operating a PL vehicle in close proximity to the ground. Hot Gas Ingestion (HGI) by the inlets can result in a rapid loss of powered lift; and high-speed jet flows along the ground plane can induce low pressures underneath the vehicle, causing a 'suck-down' effect. Under these conditions, departure from controlled flight may occur. Moreover, unsteady ground vortices and jet fountains can affect the aircraft,s controllability and its proximity to ground troops. The viscous, time-dependent flow fields of PL vehicles are difficult to accurately and efficiently predict using Computational Fluid Dynamics (CFD). A number of researchers have used the time-dependent Reynolds-averaged Navier-Stokes (RANS) equations to compute flows for single and multiple jets in a cross-flow. A few have added some geometric complexity to the problem by computing flows for jet-augmented delta wings near a ground plane. Smith et.al. computed for the first time a single RANS solution about a simplified Harrier. This geometry included a fuselage, wing, leading edge root extension (LERX), inlets, and exhaust nozzles. All of these investigations cite two practical problems with computing these flows: 1) the need for improved solution accuracy; and, 2) the need for faster solution methods. We view the need for faster solution methods as key to improving the solution accuracy and making this class of computation more routine. One can hardly refine grids, explore the use of advanced turbulence models, and generate databases when it takes weeks of dedicated computer time for a single solution. Chaderjian, Ahmad, Pandya, and Murman have focused on reducing the time-to-solution for this very difficult and complex problem through process automation and exploitation of parallel computing. They began with the Harrier geometry reported, and added a deflected wing flap and empennage for greater realism. To date more than 80 solutions have been carried out. This paper will describe this process and progress made in reducing the time required to generate a simple longitudinal force and moment database for a Harrier in ground effect. It shows a typical snap-shot from an unsteady streakline animation, where fluid particles are colored by temperature. The ground vortex and a jet-fountain vortex are highlighted. It also shows a similar streakline image, where HGI occurs due to the vehicle in close proximity to the ground. It is show the mean lift coefficient as a function of angle of attack and height. The angle of attack range was 4 deg less than or = alpha less than or = 10 deg with an increment of 1 degree, and the height range was 10 ft less than or = h less than or = 30ft with an increment of 5 feet. This 35 solution database was extended to over 2500 cases using a monotone cubic-spline interpolation procedure. The suck-down effect (reduction of lift near the ground) is highlighted in the figure. The "cushion effect," the conventional reduction of lift as the vehicle moves out of ground effect, is also indicated. All 35 RANS solutions were obtained using 952 Silicon Graphics Origin 2000 and 3000 processors in dedicated mode for one week. Typically, 112 processors were assigned to each case. Some other cases used fewer processors to utilize all available CPUS. The final paper will report on the automation of the solution process, including: grid generation, job monitoring, solution completion criteria, and post processing. Moreover, improvements in parallel efficiency for a dual time-step algorithm for the RANS equations will also be presented. Results will be discussed in detail using unsteady streakline flow visualization to correlate unsteady flow structures with dominant aerodynamic frequencies. The stability derivatives, CL, and CL, will also be presented.

Chaderjian, Neal M.↗