ALBERT

Description: Physical constraints of any real system can have a drastic effect on its performance. Some of the more recognized constraints are actuator and sensor saturation and bandwidth, power consumption, sampling rate (sensor and control-loop) and computation limits. These constraints can degrade system s performance, such as settling time, overshoot, rising time, and stability margins. In order to address these issues, researchers have investigated the use of robust and nonlinear controllers that can incorporate uncertainty and constraints into a controller design. For instance, uncertainties can be addressed in the synthesis model used in such algorithms as H(sub infinity), or mu. There is a significant amount of literature addressing this type of problem. However, there is one constraint that has not often been considered; that is, actuator authority resolution. In this work, thruster resolution and controller schemes to compensate for this effect are investigated for position and attitude control of a Low Earth Orbit formation flight system In many academic problems, actuators are assumed to have infinite resolution. In real system applications, such as formation flight systems, the system actuators will not have infinite resolution. High-precision formation flying requires the relative position and the relative attitude to be controlled on the order of millimeters and arc-seconds, respectively. Therefore, the minimum force resolution is a significant concern in this application. Without the sufficient actuator resolution, the system may be unable to attain the required pointing and position precision control. Furthermore, fuel may be wasted due to high-frequency chattering phenomena when attempting to provide a fine control with inadequate actuators. To address this issue, a Sliding Mode Controller is developed along with the boundary Layer Control to provide the best control resolution constraints. A Genetic algorithm is used to optimize the controller parameters according to the states error and fuel consumption criterion. The tradeoffs and effects of the minimum force limitation on performance are studied and compared to the case without the limitation. Furthermore, two methods are proposed to reduce chattering and improve precision.

Description: A Hele-Shaw flow apparatus constructed at Michigan State University (MSU) produces conditions that reduce influences of buoyancy-driven flows. In addition, in the MSU Hele-Shaw apparatus it is possible to adjust the heat losses from the fuel sample (0.001 in. thick cellulose) and the flow speed of the approaching oxidizer flow (air) so that the "flamelet regime of flame spread" is entered. In this regime various features of the flame-to-smolder (and vice versa) transition can be studied. For the relatively wide (approx. 17.5 cm) and long (approx. 20 cm) samples used, approximately ten flamelets existed at all times. The flamelet behavior was studied mechanistically and statistically. A heat transfer analysis of the dominant heat transfer mechanisms was conducted. Results indicate that radiation and conduction processes are important, and that a simple 1-D model using the Broido-Shafizadeh model for cellulose decomposition chemistry can describe aspects of the flamelet spread process. Introduction

Description: Combustion experiments using arrays of droplets seek to provide a link between single droplet combustion phenomena and the behavior of complex spray combustion systems. Both single droplet and droplet array studies have been conducted in microgravity to better isolate the droplet interaction phenomena and eliminate or reduce the effects of buoyancy-induced convection. In most experiments involving droplet arrays, the droplets are supported on fibers to keep them stationary and close together before the combustion event. The presence of the fiber, however, disturbs the combustion process by introducing a source of heat transfer and asymmetry into the configuration. As the number of drops in a droplet array increases, supporting the drops on fibers becomes less practical because of the cumulative effect of the fibers on the combustion process. To eliminate the effect of the fiber, several researchers have conducted microgravity experiments using unsupported droplets. Jackson and Avedisian investigated single, unsupported drops while Nomura et al. studied droplet clouds formed by a condensation technique. The overall objective of this research is to extend the study of unsupported drops by investigating the combustion of well-characterized drop clusters in a microgravity environment. Direct experimental observations and measurements of the combustion of droplet clusters would provide unique experimental data for the verification and improvement of spray combustion models. In this work, the formation of drop clusters is precisely controlled using an acoustic levitation system so that dilute, as well as dense clusters can be created and stabilized before combustion in microgravity is begun. While the low-gravity test facility is being completed, tests have been conducted in 1-g to characterize the effect of the acoustic field on the vaporization of single and multiple droplets. This is important because in the combustion experiment, the droplets will be formed and levitated prior to ignition. Therefore, the droplets will begin to vaporize in the acoustic field thus forming the "initial conditions" for the combustion process. Understanding droplet vaporization in the acoustic field of this levitator is a necessary step that will help to interpret the experimental results obtained in low-gravity.

Description: A viewgraph presentation on the concept of compliant casing for transonic axial compressors is shown. The topics include: 1) Concept for compliant casing; 2) Rig and facility details; 3) Experimental results; and 4) Numerical results.

Description: Model the interactions between the structural dynamics and the performance dynamics of a gas turbine engine. Generally these two aspects are considered separate, unrelated phenomena and are studied independently. For diagnostic purposes, it is desirable to bring together as much information as possible, and that involves understanding how performance is affected by structural dynamics (if it is) and vice versa. This can involve the relationship between thrust response and the excitation of structural modes, for instance. The job will involve investigating and characterizing these dynamical relationships, generating a model that incorporates them, and suggesting and/or developing diagnostic and prognostic techniques that can be incorporated in a data fusion system. If no coupling is found, at the least a vibration model should be generated that can be used for diagnostics and prognostics related to blade loss, for instance.

Description: The fellowship research project continued last-year work on implementing implicit time marching concepts in the Broadband Aeroacoustic System Simulator (BASS) code. This code is being developed at NASA Glenn for analysis of unsteady flow and sources of noise in propulsion systems, including jet noise and fan noise.

Description: Modern fan designs have blades with forward sweep; a lean, thin cross section; and a wide chord to improve performance and reduce noise. These geometric features coupled with the presence of a shock wave can lead to flutter instability. Flutter is a self-excited dynamic instability arising because of fluid-structure interaction, which causes the energy from the surrounding fluid to be extracted by the vibrating structure. An in-flight occurrence of flutter could be catastrophic and is a significant design issue for rotor blades in gas turbines. Understanding the flutter behavior and the influence of flow features on flutter will lead to a better and safer design. An aeroelastic analysis code, TURBO, has been developed and validated for flutter calculations at the NASA Glenn Research Center. The code has been used to understand the occurrence of flutter in a forward-swept fan design. The forward-swept fan, which consists of 22 inserted blades, encountered flutter during wind tunnel tests at part speed conditions.

Description: This work is motivated by the need to accurately predict heat transfer in turbomachinery. For efficient gas turbine operation, flow temperatures in the hot gas path exceed acceptable metal temperatures in many regions of the engine. So that the integrity of the parts can be maintained for an acceptable engine life, the parts must be cooled. Efficient cooling schemes require accurate heat transfer prediction to minimize regions that are overcooled and, even more importantly, to ensure adequate cooling in high-heat-flux regions.

Description: Future aeropropulsion gas turbine engines must be affordable in addition to being energy efficient and environmentally benign. Progress in aerodynamic design capability is required not only to maximize the specific thrust of next-generation engines without sacrificing fuel consumption, but also to reduce parts count by increasing the aerodynamic loading of the compression system. To meet future compressor requirements, the NASA Glenn Research Center is investigating advanced aerodynamic design concepts that will lead to more compact, higher efficiency, and wider operability configurations than are currently in operation.

Description: Noise sources in high-speed jets were identified by directly correlating flow density fluctuation (cause) to far-field sound pressure fluctuation (effect). The experimental study was performed in a nozzle facility at the NASA Glenn Research Center in support of NASA s initiative to reduce the noise emitted by commercial airplanes. Previous efforts to use this correlation method have failed because the tools for measuring jet turbulence were intrusive. In the present experiment, a molecular Rayleigh-scattering technique was used that depended on laser light scattering by gas molecules in air. The technique allowed accurate measurement of air density fluctuations from different points in the plume. The study was conducted in shock-free, unheated jets of Mach numbers 0.95, 1.4, and 1.8. The turbulent motion, as evident from density fluctuation spectra was remarkably similar in all three jets, whereas the noise sources were significantly different. The correlation study was conducted by keeping a microphone at a fixed location (at the peak noise emission angle of 30 to the jet axis and 50 nozzle diameters away) while moving the laser probe volume from point to point in the flow. The following figure shows maps of the nondimensional coherence value measured at different Strouhal frequencies ([frequency diameter]/jet speed) in the supersonic Mach 1.8 and subsonic Mach 0.95 jets. The higher the coherence, the stronger the source was.

Description: Forced response, or resonant vibrations, in turbomachinery components can cause blades to crack or fail because of the large vibratory blade stresses and subsequent high-cycle fatigue. Forced-response vibrations occur when turbomachinery blades are subjected to periodic excitation at a frequency close to their natural frequency. Rotor blades in a turbine are constantly subjected to periodic excitations when they pass through the spatially nonuniform flowfield created by upstream vanes. Accurate numerical prediction of the unsteady aerodynamics phenomena that cause forced-response vibrations can lead to an improved understanding of the problem and offer potential approaches to reduce or eliminate specific forced-response problems. The objective of the current work was to validate an unsteady aerodynamics code (named TURBO) for the modeling of the unsteady blade row interactions that can cause forced response vibrations. The three-dimensional, unsteady, multi-blade-row, Reynolds-averaged Navier-Stokes turbomachinery code named TURBO was used to model a high-pressure turbine stage for which benchmark data were recently acquired under a NASA contract by researchers at the Ohio State University. The test article was an initial design for a high-pressure turbine stage that experienced forced-response vibrations which were eliminated by increasing the axial gap. The data, acquired in a short duration or shock tunnel test facility, included unsteady blade surface pressures and vibratory strains.

Description: The High-Flow Jet Exit Rig at the NASA Glenn Research Center is designed to test single flow jet nozzles and to measure the appropriate thrust and noise levels. The rig has been designed for the maximum hot condition of 16 lbm/sec of combustion air at 1960 R (maximum) and to produce a maximum thrust of 2000 lb. It was designed for cold flow of 29.1 lbm/sec of air at 530 R. In addition, it can test dual-flow nozzles (nozzles with bypass flow in addition to core flow) with independent control of each flow. The High- Flow Jet Exit Rig was successfully fabricated in late 2001 and is being readied for checkout tests. The rig will be installed in Glenn's Aeroacoustic Propulsion Laboratory. The High-Flow Jet Exit Rig consists of the following major components: a single component force balance, the natural-gas-fueled J-79 combustor assembly, the plenum and manifold assembly, an acoustic/instrumentation/seeding (A/I/S) section, a table, and the research nozzles. The rig will be unique in that it is designed to operate uncooled. The structure survives the 1960 R test condition because it uses carefully selected high temperature alloy materials such as Hastelloy-X. The lower plenum assembly was designed to operate at pressures to 450 psig at 1960 R, in accordance with the ASME B31.3 piping code. The natural gas-fueled combustor fires directly into the lower manifold. The hot air is directed through eight 1-1/2-in. supply pipes that supply the upper plenum. The flow is conditioned in the upper plenum prior to flowing to the research nozzle. The 1-1/2-in. supply lines are arranged in a U-shaped design to provide for a flexible piping system. The combustor assembly checkout was successfully conducted in Glenn's Engine Component Research Laboratory in the spring of 2001. The combustor is a low-smoke version of the J79 combustor used to power the F4 Phantom military aircraft. The natural gas-fueled combustor demonstrated high-efficiency combustion over a wide range of operating conditions. This wide operating envelope is required to support the testing of both single- and dual-flow nozzles. Key research goals include providing simultaneous, highly accurate acoustic, flow, and thrust measurements on jet nozzle models in realistic flight conditions, as well as providing scaleable acoustic results. The High-Flow Jet Exit Rig is a second-generation high-flow test rig. Improvements include cleaner flow with reduced levels of particulate, soot, and odor. Choked-flow metering is required with plus or minus 0.25-percent accuracy. Thrust measurements from 0 to 2000 lbf are required with plus or minus 0.25-percent accuracy. Improved acoustics will be achieved by minimizing noise through large pipe bend radii, lower internal flow velocities, and microdrilled choke plates with thousands of 0.040-in.- diameter holes.

Description: A new test stand, the Small Hot Jet Acoustic Rig, was commissioned into service at NASA Glenn Research Center's Aeroacoustic Propulsion Laboratory. This new rig provides researchers with an all-in-one platform with which to economically evaluate the thrust performance, acoustic performance, and plume turbulence characteristics of new nozzle concepts. It features an integral force balance, exceptionally low internal flownoise, and provisions to conduct laser-based plume turbulence studies with Particle Imaging Velocimetry, shadowgraphs, schlieren photography, and other techniques. The rig also features an integral combustor and can deliver air to the test nozzle at temperatures ranging from ambient to 1300 F. The Small Hot Jet Acoustic Rig is the fourth semipermanent rig now residing in the Aeroacoustic Propulsion Laboratory. It will add to the facility's substantial list of acoustic research capabilities and improve its already impressive productivity.

Description: A major source of fan noise in commercial turbofan engines is the interaction of the wake from the fan blades with the stationary vanes (stators) directly behind them. The Trailing Edge Blowing (TEB) project team at the NASA Glenn Research Center designed and fabricated new fan blades to study the effects of fan trailing edge blowing as a potential noise-reduction concept. The intent is to fill the rotor wake by supplying air to the rotor blade trailing edge at the proper conditions to minimize the wake deficit, and thus generate less noise. The TEB hardware is designed for the Active Noise Control Fan (ANCF) test rig in Glenn's Aeroacoustic Propulsion Laboratory. For this test, the air is fed from an external supply through the shaft of the rig. It is distributed to the base of each blade through an impeller, where it is forced into a plenum at the core of each blade. In actual engine configuration, air would most likely be bled from the compressor, but only at times when noise is an issue, such as takeoffs and landings. Glenn researchers designed and manufactured the blades in-house, using new techniques and concepts. The skins, which were designed for maximum strength in the directions of highest stress, were molded from multiple layers of carbon fiber. Considerable use was made of rapid prototyping techniques, such as laser sintering. The core was sintered from a lightweight polymer, and the retainer was CNC-machined (computer numerical control machined) from aluminum. All the components were joined with a cold-cure aerospace adhesive. These techniques and processes reduced the overall cost and allowed the new concept to be studied much sooner than would be possible using traditional fabrication methods. Since this test rig did not support the use of blade-monitoring techniques such as strain gauges, extensive bench testing was required to qualify the design. The blades were examined using a variety of methods including holography, pull tests (cyclic and failure), shake tests, rap tests, and nondestructive inspection. Acoustic testing of the ANCF fan using TEB has been ongoing since January of 2001. The fan has completed about 100 hr of testing with no structural, vibrational, or fatigue problems. Far-field acoustic measurements, in-duct mode measurements, precise hot wire surveys, and detailed performance measurements are providing data for evaluating the concept. The far-field noise data show that tone noise was reduced significantly with the initial ANCF TEB fan design. In addition, a significant reduction in unsteady stator loading has been measured, indicating the potential for stator broadband noise reduction. The acoustic benefits will be assessed and important design parameters identified to improve the ability to fully exploit any benefit provided by this technique. On the basis of the success of trailing edge blowing, Glenn plans to continue this research with a higher speed, higher pressure ratio fan operating in an acoustic wind tunnel to simulate flight conditions.

Description: The NASA Glenn Research Center was the major contributor of 2-kW-class ion thruster technology to the Deep Space 1 mission, which was successfully completed in early 2002. Recently, NASA s Office of Space Science awarded approximately $21 million to Glenn to develop higher power xenon ion propulsion systems for large flagship missions such as outer planet explorers and sample return missions. The project, referred to as NASA's Evolutionary Xenon Thruster (NEXT), is a logical follow-on to the ion propulsion system demonstrated on Deep Space 1. The propulsion system power level for NEXT is expected to be as high as 25 kW, incorporating multiple ion thrusters, each capable of being throttled over a 1- to 6-kW power range. To date, engineering model thrusters have been developed, and performance and plume diagnostics are now being documented. The project team-Glenn, the Jet Propulsion Laboratory, General Dynamics, Boeing Electron Dynamic Devices, the Applied Physics Laboratory, the University of Michigan, and Colorado State University-is in the process of developing hardware for a ground demonstration of the NEXT propulsion system, which comprises a xenon feed system, controllers, multiple thrusters, and power processors. The development program also will include life assessments by tests and analyses, single-string tests of ion thrusters and power systems, and finally, multistring thruster system tests in calendar year 2005. In addition, NASA's Office of Space Science selected Glenn to lead the development of a 25-kW xenon thruster to enable NASA to conduct future missions to the outer planets of Jupiter and beyond, under the High Power Electric Propulsion (HiPEP) program. The development of a 100-kW-class ion propulsion system and power conversion systems are critical components to enable future nuclear-electric propulsion systems. In fiscal year 2003, a team composed of Glenn, the Boeing Company, General Dynamics, the Applied Physics Laboratory, the Naval Research Laboratory, the University of Wisconsin, the University of Michigan, and Colorado State University will perform a 6-month study that will result in the design of a 25-kW ion thruster, a propellant feed system, and a power processing architecture. The following 2 years will involve hardware development, wear tests, single-string tests of the thruster-power circuits and the xenon feed system, and subsystem service life analyses. The 2-kW-class ion propulsion technology developed for the Deep Space 1 mission will be used for NASA's discovery mission Dawn, which involves maneuvering a spacecraft to survey the asteroids Ceres and Vesta. The 6-kW-class ion thruster subsystem technology under NEXT is scheduled to be flight ready by calendar year 2006. The less mature 25- kW ion thruster system under HiPEP is expected to be ready for a flight advanced development program in calendar year 2006.

Description: Twenty-first-century aeropropulsion and power research will enable new transport engine and aircraft systems including: 1) Emerging ultralow noise and emissions with the use of intelligent turbofans; 2) Future distributed vectored propulsion with 24-hour operations and greater community mobility; 3) Research in hybrid combustion and electric propulsion systems leading to silent aircraft with near-zero emissions; and 4) The culmination of these revolutions will deliver an all-electric- powered propulsion system with zero-impact emissions and noise and high-capacity, on-demand operation

Description: The objective is to develop the capability to numerically model the performance of gas turbine engines used for aircraft propulsion. This capability will provide turbine engine designers with a means of accurately predicting the performance of new engines in a system environment prior to building and testing. The 'numerical test cell' developed under this project will reduce the number of component and engine tests required during development. As a result, the project will help to reduce the design cycle time and cost of gas turbine engines. This capability will be distributed to U.S. turbine engine manufacturers and air framers. This project focuses on goals of maintaining U.S. superiority in commercial gas turbine engine development for the aeronautics industry.

Description: In this work, we have considered an annular cascade configuration subjected to unsteady inflow conditions. The unsteady response calculation has been implemented into the time marching CFD code, MSUTURBO. The computed steady state results for the pressure distribution demonstrated good agreement with experimental data. We have computed results for the amplitudes of the unsteady pressure over the blade surfaces. With the increase in gas turbine engine structural complexity and performance over the past 50 years, structural engineers have created an array of safety nets to ensure against component failures in turbine engines. In order to reduce what is now considered to be excessive conservatism and yet maintain the same adequate margins of safety, there is a pressing need to explore methods of incorporating probabilistic design procedures into engine development. Probabilistic methods combine and prioritize the statistical distributions of each design variable, generate an interactive distribution and offer the designer a quantified relationship between robustness, endurance and performance. The designer can therefore iterate between weight reduction, life increase, engine size reduction, speed increase etc.

Description: The purpose of this article is to show that the Navier-Stokes equations can be rewritten as a set of linearized inhomogeneous Euler equations (in convective form) with source terms that are exactly the same as those that would result from externally imposed shear stress and energy flux perturbations. These results are used to develop a mathematical basis for some existing and potential new jet noise models by appropriately choosing the base flow about which the linearization is carried out.

Description: The low-emissions combustor development at the NASA Glenn Research Center is directed toward advanced high-pressure aircraft gas turbine applications. The emphasis of this research is to reduce nitrogen oxides (NOx) at high-power conditions and to maintain carbon monoxide and unburned hydrocarbons at their current low levels at low-power conditions. Low-NOx combustors can be classified into rich burn and lean burn concepts. Lean burn combustors can be further classified into lean-premixed-prevaporized (LPP) and lean direct injection (LDI) combustors. In both concepts, all the combustor air, except for liner cooling flow, enters through the combustor dome so that the combustion occurs at the lowest possible flame temperature. The LPP concept has been shown to have the lowest NOx emissions, but for advanced high-pressure-ratio engines, the possibly of autoignition or flashback precludes its use. LDI differs from LPP in that the fuel is injected directly into the flame zone and, thus, does not have the potential for autoignition or flashback and should have greater stability. However, since it is not premixed and prevaporized, the key is good atomization and mixing of the fuel quickly and uniformly so that flame temperatures are low and NOx formation levels are comparable to those of LPP.

Description: It shows the variation in compressor mass flow with time as the mass flow is throttled to drive the compressor into surge. Surge begins where wide variations in mass flow occur. Air injection is then turned on to bring about a recovery from the initial surge condition and stabilize the compressor. The throttle is closed further until surge is again initiated. Air injection is increased to again recover from the surge condition and stabilize the compressor.

Description: The performance of compressors and the sophistication of analysis tools have reached a level such that less well understood flow mechanisms are gaining importance to designers. In current design systems, the effect on performance of many of these mechanisms, such as blade row interactions, is not typically addressed rigorously. A detailed set of Laser Doppler Velocimetry data was used to confirm the fidelity of an unsteady model of a transonic compressor stage (rotor-stator) simulated with the TURBO unsteady multistage turbomachinery solver. The kinematics of the velocity field were accurately simulated, and the unsteady simulation was then used to assess changes in loss production due to unsteady blade-row-interaction mechanisms. This work was done at the NASA Glenn Research Center in support of the Ultra-Efficient Engine Technology Program.

Description: Unsteady ejectors are currently under investigation for use in some pulse detonation engine (PDE) propulsion systems. This is due primarily to their potential high performance in comparison to steady ejectors of similar dimensions relative to the source or driver jet. Although some experimental work has been done in the past to study thrust augmentation with unsteady ejectors, there is no proven theory by which optimal design parameters can be selected and an effective ejector constructed for a given pulsed flow. Therefore, an experimental facility was developed at the NASA Glenn Research Center to study the correlation between ejector design and performance, and to get a better understanding of the flow phenomena that result in thrust augmentation. A commercially available pulsejet was used for the unsteady driving jet. This was paired with a basic, yet flexible, ejector design that allowed parametric evaluation of the effects that length, diameter, and inlet radius have on performance.

Description: To support the Revolutionary Aeropropulsion Concept Program, NASA Glenn Research Center' s Structural Mechanics and Dynamics Branch is developing a compact, nonpolluting, bearingless electric machine with electric power supplied by fuel cells for future more-electric aircraft. The use of such electric drives for propulsive fans or propellers depends on the successful development of ultra-high-power-density machines that can generate power densities of 50 hp/lb or more, whereas conventional electric machines generate usually 0.2 hp/lb. One possible candidate for such ultra-high-power-density machines, a round-rotor synchronous machine with an engineering current density as high as 20 000 A/cm2 was selected to investigate how much torque and power can be produced. A simple synchronous machine model that consists of rotor and stator windings and back-irons was considered first. The model had a sinusoidally distributed winding that produces a sinusoidal distribution of flux P poles. Excitation of the rotor winding produced P poles of rotor flux, which interacted with the P stator poles to produce torque.

Description: A new, molecular Rayleigh-scattering-based flow diagnostic developed at the NASA Glenn Research Center has been used for the first time to measure the power spectrum of both gas density and radial velocity components in the plumes of high-speed jets. The objective of the work is to develop an unseeded, nonintrusive dynamic measurement technique for studying turbulent flows in NASA test facilities. This technique provides aerothermodynamic data not previously obtainable. It is particularly important for supersonic flows, where hot wire and pitot probes are difficult to use and disturb the flow under study. The effort is part of the nonintrusive instrumentation development program supporting propulsion research at the NASA Glenn Research Center. In particular, this work is measuring fluctuations in flow velocity, density, and temperature for jet noise studies. These data are valuable to researchers studying the correlation of flow fluctuations with far-field noise. One of the main objectives in jet noise research is to identify noise sources in the jet and to determine their contribution to noise generation. The technique is based on analyzing light scattered from molecules within the jet using a Fabry-Perot interferometer operating in a static imaging mode. The PC-based data acquisition system can simultaneously sample velocity and density data at rates to about 100 kHz and can handle up to 10 million data records. We used this system to interrogate three different jet nozzle designs in a Glenn free-jet facility. Each nozzle had a 25.4-mm exit diameter. One was convergent, used for subsonic flow measurements and to produce a screeching underexpanded jet with a fully expanded Mach number of 1.42. The other nozzles (Mach 1.4 and 1.8) were convergent-divergent types. The radial component of velocity and gas density were simultaneously measured in this work.

Description: High-power electric propulsion systems have been shown to be enabling for a number of NASA concepts, including piloted missions to Mars and Earth-orbiting solar electric power generation for terrestrial use (refs. 1 and 2). These types of missions require moderate transfer times and sizable thrust levels, resulting in an optimized propulsion system with greater specific impulse than conventional chemical systems and greater thrust than ion thruster systems. Hall thruster technology will offer a favorable combination of performance, reliability, and lifetime for such applications if input power can be scaled by more than an order of magnitude from the kilowatt level of the current state-of-the-art systems. As a result, the NASA Glenn Research Center conducted strategic technology research and development into high-power Hall thruster technology. During program year 2002, an in-house fabricated thruster, designated the NASA-457M, was experimentally evaluated at input powers up to 72 kW. These tests demonstrated the efficacy of scaling Hall thrusters to high power suitable for a range of future missions. Thrust up to nearly 3 N was measured. Discharge specific impulses ranged from 1750 to 3250 sec, with discharge efficiencies between 46 and 65 percent. This thruster is the highest power, highest thrust Hall thruster ever tested.

Description: Integration of entire system includes: Fuel cells, motors, propulsors, thermal/power management, compressors, etc. Use of existing, pre-developed NPSS capabilities includes: 1) Optimization tools; 2) Gas turbine models for hybrid systems; 3) Increased interplay between subsystems; 4) Off-design modeling capabilities; 5) Altitude effects; and 6) Existing transient modeling architecture. Other factors inclde: 1) Easier transfer between users and groups of users; 2) General aerospace industry acceptance and familiarity; and 3) Flexible analysis tool that can also be used for ground power applications.

Description: The objective is to develop high fidelity tools that can influence ISTAR design In particular, tools for coupling Fluid-Thermal-Structural simulations RBCC/TBCC designers carefully balance aerodynamic, thermal, weight, & structural considerations; consistent multidisciplinary solutions reveal details (at modest cost) At Scram mode design point, simulations give details of inlet & combustor performance, thermal loads, structural deflections.

Description: The goal of this research is to develop an advanced engineering analysis system that enables high-fidelity, multi-disciplinary, full propulsion system simulations to be performed early in the design process (a virtual test cell that integrates propulsion and information technologies). This will enable rapid, high-confidence, cost-effective design of revolutionary systems.

Description: The objective is to develop a coupled fluid/structure analysis tool for rocket turbopumps, advance hardware concepts and designs, and improve safety, reliability, and cost of space transportation.

Description: A multidisciplinary effort is underway at the NASA Glenn Research Center to develop concepts for revolutionary, nontraditional fuel cell power and propulsion systems for aircraft applications. There is a growing interest in the use of fuel cells as a power source for electric propulsion as well as an auxiliary power unit to substantially reduce or eliminate environmentally harmful emissions. A systems analysis effort was initiated to assess potential concepts in an effort to identify those configurations with the highest payoff potential. Among the technologies under consideration are advanced proton exchange membrane (PEM) and solid oxide fuel cells, alternative fuels and fuel processing, and fuel storage. Prior to this effort, the majority of fuel cell analysis done at Glenn was done for space applications. Because of this, a new suite of models was developed. These models include the hydrogen-air PEM fuel cell; internal reforming solid oxide fuel cell; balance-of-plant components (compressor, humidifier, separator, and heat exchangers); compressed gas, cryogenic, and liquid fuel storage tanks; and gas turbine/generator models for hybrid system applications. Initial mass, volume, and performance estimates of a variety of PEM systems operating on hydrogen and reformate have been completed for a baseline general aviation aircraft. Solid oxide/turbine hybrid systems are being analyzed. In conjunction with the analysis efforts, a joint effort has been initiated with Glenn s Computer Services Division to integrate fuel cell stack and component models with the visualization environment that supports the GRUVE lab, Glenn s virtual reality facility. The objective of this work is to provide an environment to assist engineers in the integration of fuel cell propulsion systems into aircraft and provide a better understanding of the interaction between system components and the resulting effect on the overall design and performance of the aircraft. Initially, three-dimensional computer-aided design (CAD) models of representative PEM fuel cell stack and components were developed and integrated into the virtual reality environment along with an Excel-based model used to calculate fuel cell electrical performance on the basis of cell dimensions (see the figure). CAD models of a representative general aviation aircraft were also developed and added to the environment. With the use of special headgear, users will be able to virtually manipulate the fuel cell s physical characteristics and its placement within the aircraft while receiving information on the resultant fuel cell output power and performance. As the systems analysis effort progresses, we will add more component models to the GRUVE environment to help us more fully understand the effect of various system configurations on the aircraft.

Description: Alan Powell has made significant contributions to the understanding of many aeroacoustic problems, in particular, the problems of broadband noise from jets and boundary layers. In this paper, some analytic results are presented for the calculation of the correlation function of the broadband noise radiated from a wing, a propeller, and a jet in uniform forward motion. It is shown that, when the observer (or microphone) motion is suitably chosen, the geometric terms of the radiation formula become time independent. The time independence of these terms leads to a significant simplification of the statistical analysis of the radiated noise, even when the near field terms are included. For a wing in forward motion, if the observer is in the moving reference frame, then the correlation function of the near and far field noise can be related to a space-time cross-correlation function of the pressure on the wing surface. A similar result holds for a propeller in forward flight if the observer is in a reference frame that is attached to the propeller and rotates at the shaft speed. For a jet in motion, it is shown that the correlation function of the radiated noise can be related to the space-time crosscorrelation of the Lighthill stress tensor in the jet. Exact analytical results are derived for all three cases. For the cases under present consideration, the inclusion of the near field terms does not introduce additional complexity, as compared to existing formulations that are limited to the far field.

Description: Combustor liners fabricated from a SiC/SiC composite were nondestructively interrogated before and after combustion rig testing by X-ray, ultrasonic and thermographic techniques. In addition, mechanical test results were obtained from witness coupons, representing the as-manufactured liners, and from coupons machined from the components after combustion exposure. Thermography indications were found to correlate with reduced material properties obtained after rig testing. The thermography indications in the SiC/SiC liners were delaminations and damaged fiber tows, as determined through microstructural examinations. [copyright] 2003 American Institute of Physics

Description: NASA Glenn Research Center is currently evaluating the possibility of using high- temperature polymer matrix composites to reinforce the combustion chamber of a rocket engine. One potential design utilizes a honeycomb structure composed of a PMR-II- 50/M40J 4HS composite facesheet and titanium honeycomb core to reinforce a stainless steel shell. In order to properly fabricate this structure, adhesive bond PMR-II-50 composite. Proper prebond surface preparation is critical in order to obtain an acceptable adhesive bond. Improperly treated surfaces will exhibit decreased bond strength and durability, especially in metallic bonds where interface are susceptible to degradation due to heat and moisture. Most treatments for titanium and stainless steel alloys require the use of strong chemicals to etch and clean the surface. This processes are difficult to perform due to limited processing facilities as well as safety and environmental risks and they do not consistently yield optimum bond durability. Boeing Phantom Works previously developed sol-gel surface preparations for titanium alloys using a PETI-5 based polyimide adhesive. In support of part of NASA Glenn Research Center, UDRI and Boeing Phantom Works evaluated variations of this high temperature sol-gel surface preparation, primer type, and primer cure conditions on the adhesion performance of titanium and stainless steel using Cytec FM 680-1 polyimide adhesive. It was also found that a modified cure cycle of the FM 680-1 adhesive, i.e., 4 hrs at 370 F in vacuum + post cure, significantly increased the adhesion strength compared to the manufacturer's suggested cure cycle. In addition, the surface preparation of the PMR-II-50 composite was evaluated in terms of surface cleanness and roughness. This presentation will discuss the results of strength and durability testing conducted on titanium, stainless steel, and PMR-II-50 composite adherends to evaluate possible bonding processes.

Description: Applying binaural simulation techniques to structural acoustic data can be very computationally intensive as the number of discrete noise sources can be very large. Typically, Head Related Transfer Functions (HRTFs) are used to individually filter the signals from each of the sources in the acoustic field. Therefore, creating a binaural simulation implies the use of potentially hundreds of real time filters. This paper details two methods of reducing the number of real-time computations required by: (i) using the singular value decomposition (SVD) to reduce the complexity of the HRTFs by breaking them into dominant singular values and vectors and (ii) by using equivalent source reduction (ESR) to reduce the number of sources to be analyzed in real-time by replacing sources on the scale of a structural wavelength with sources on the scale of an acoustic wavelength. The ESR and SVD reduction methods can be combined to provide an estimated computation time reduction of 99.4% for the structural acoustic data tested. In addition, preliminary tests have shown that there is a 97% correlation between the results of the combined reduction methods and the results found with the current binaural simulation techniques

Description: A conical resonator (having a dissonant acoustic design) was tested in four configurations: (1) baseline resonator with closed ends and no blockage, (2) closed resonator with internal blockage, (3) ventilated resonator with no blockage, and (4) ventilated resonator with an applied pressure differential. These tests were conducted to investigate the effects of blockage and ventilation holes on dynamic pressurization. Additionally, the investigation was to determine the ability of acoustic pressurization to impede flow through the resonator. In each of the configurations studied, the entire resonator was oscillated at the gas resonant frequency while dynamic pressure, static pressure, and temperature of the fluid were measured. In the final configuration, flow through the resonator was recorded for three oscillation conditions. Ambient condition air was used as the working fluid.

Description: This paper describes active tip clearance control research being conducted by NASA to improve turbine engine systems. The target application for this effort is commercial aircraft engines. However, technologies developed for clearance control can benefit a broad spectrum of current and future turbomachinery. The first portion of the paper addresses the research from a programmatic viewpoint. Recent studies that provide motivation for the work, identification of key technologies, and NASA's plan for addressing deficiencies in the technologies are discussed. The later portion of the paper drills down into one of the key technologies by presenting equations and results for a preliminary dynamic model of the tip clearance phenomena.

Description: Space Shuttle Reusable Solid Rocket Motors (RSRM) are static tested at two ATK Thiokol Propulsion facilities in Utah, T-24 and T-97. The newer T-97 static test facility was recently upgraded to allow thrust measurement capability. All previous static test motor thrust measurements have been taken at T-24; data from these tests were used to characterize thrust parameters and requirement limits for flight motors. Validation of the new T-97 thrust measurement system is required prior to use for official RSRM performance assessments. Since thrust cannot be measured on RSRM flight motors, flight motor measured chamber pressure and a nominal thrust-to-pressure relationship (based on static test motor thrust and pressure measurements) are used to reconstruct flight motor performance. Historical static test and flight motor performance data are used in conjunction with production subscale test data to predict RSRM performance. The predicted motor performance is provided to support Space Shuttle trajectory and system loads analyses. Therefore, an accurate nominal thrust-to-pressure (F/P) relationship is critical for accurate RSRM flight motor performance and Space Shuttle analyses. Flight Support Motors (FSM) 7, 8, and 9 provided thrust data for the validation of the T-97 thrust measurement system. The T-97 thrust data were analyzed and compared to thrust previously measured at T-24 to verify measured thrust data and identify any test-stand bias. The T-97 FIP data were consistent and within the T-24 static test statistical family expectation. The FSMs 7-9 thrust data met all NASA contract requirements, and the test stand is now verified for future thrust measurements.

Description: Recent studies of xenon Hall thrusters have shown peak efficiencies at specific impulses of less than 3000 s. This was a consequence of modern Hall thruster magnetic field topographies, which have been optimized for 300 V discharges. On-going research at the NASA Glenn Research Center is investigating this behavior and methods to enhance thruster performance. To conduct these studies, a laboratory model Hall thruster that uses a pair of trim coils to tailor the magnetic field topography for high specific impulse operation has been developed. The thruster-the NASA-173Mv2 was tested to determine how current density and magnetic field topography affect performance, divergence, and plasma oscillations at voltages up to 1000 V. Test results showed there was a minimum current density and optimum magnetic field topography at which efficiency monotonically increased with voltage. At 1000 V, 10 milligrams per second the total specific impulse was 3390 s and the total efficiency was 60.8%. Plume divergence decreased at 400-1000 V, but increased at 300-400 V as the result of plasma oscillations. The dominant oscillation frequency steadily increased with voltage, from 14.5 kHz at 300 V, to 22 kHz at 1000 V. An additional oscillatory mode in the 80-90 kHz frequency range began to appear above 500 V. The use of trim coils to modify the magnetic field improved performance while decreasing plume divergence and the frequency and magnitude of plasma oscillations.

Description: Torque tension testing of a newly designed Reusable Solid Rocket Motor nozzle bolted assembly was successfully completed. Test results showed that the 3-sigma preload variation was as expected at the required input torque level and the preload relaxation were within the engineering limits. A shim installation technique was demonstrated as a simple process to fill a shear lip gap between nozzle housings in the joint region. A new automated torque system was successfully demonstrated in this test. This torque control tool was found to be very precise and accurate. The bolted assembly performance was further evaluated using the Nozzle Structural Test Bed. Both current socket head cap screw and proposed multiphase alloy bolt configurations were tested. Results indicated that joint skip and bolt bending were significantly reduced with the new multiphase alloy bolt design. This paper summarizes all the test results completed to date.

Description: This talk will provide an overview of the field of computational aero-acoustics and its use in fan noise prediction. After a brief history of computational fluid dynamics, some of the recent developments in computational aero-acoustics will be explored. Computational issues concerning sound wave production, propagation, and reflection in practical turbo-machinery applications will be discussed including: (a) High order/High Resolution Numerical Techniques. (b) High Resolution Boundary Conditions. [c] MIMD Parallel Computing. [d] Form of Governing Equations Useful for Simulations. In addition, the basic design of our Broadband Analysis Stator Simulator (BASS) code and its application to a 2 D rotor wake-stator interaction will be shown. An example of the noise produced by the wakes from a rotor impinging upon a stator cascade will be shown.

Description: The space-time conservation-element and solution-element method is employed to numerically study the near-field screech-tone noise of a typical underexpanded circular jet issuing from a sonic nozzle. Both axisymmetric and fully three-dimensional computations are carried out. The self-sustained feedback loop is properly simulated. The computed shock-cell structure, acoustic wave length, screech-tone frequency, and sound-pressure levels are in good agreement with existing experimental results.

Description: Nonlinear acoustic standing waves of high amplitude have been demonstrated by utilizing the effects of resonator shape to prevent the pressure waves from entering saturation. Experimentally, nonlinear acoustic standing waves have been generated by shaking an entire resonating cavity. While this promotes more efficient energy transfer than a piston-driven resonator, it also introduces complicated structural dynamics into the system. Experiments have shown that these dynamics result in resonator forcing functions comprised of a sum of several Fourier modes. However, previous numerical studies of the acoustics generated within the resonator assumed simple sinusoidal waves as the driving force. Using a previously developed numerical code, this paper demonstrates the effects of using a forcing function constructed with a series of harmonic sinusoidal waves on resonating cavities. From these results, a method will be demonstrated which allows the direct numerical analysis of experimentally generated nonlinear acoustic waves in resonators driven by harmonic forcing functions.

Description: The 2002 annual report of the Structural Mechanics and Dynamics Branch reflects the majority of the work performed by the branch staff during the 2002 calendar year. Its purpose is to give a brief review of the branch s technical accomplishments. The Structural Mechanics and Dynamics Branch develops innovative computational tools, benchmark experimental data, and solutions to long-term barrier problems in the areas of propulsion aeroelasticity, active and passive damping, engine vibration control, rotor dynamics, magnetic suspension, structural mechanics, probabilistics, smart structures, engine system dynamics, and engine containment. Furthermore, the branch is developing a compact, nonpolluting, bearingless electric machine with electric power supplied by fuel cells for future "more electric" aircraft. An ultra-high-power-density machine that can generate projected power densities of 50 hp/lb or more, in comparison to conventional electric machines, which generate usually 0.2 hp/lb, is under development for application to electric drives for propulsive fans or propellers. In the future, propulsion and power systems will need to be lighter, to operate at higher temperatures, and to be more reliable in order to achieve higher performance and economic viability. The Structural Mechanics and Dynamics Branch is working to achieve these complex, challenging goals.

Description: This paper summarizes major theoretical results for pulse detonation engine performance taking into account real gas chemistry, as well as significant performance differences resulting from the presence of ram and compression heating. An unsteady CFD analysis, as well as a thermodynamic cycle analysis, was conducted in order to determine the actual and the ideal performance for an air-breathing pulse detonation engine (PDE) using either a hydrogen-air or ethylene-air mixture over a flight Mach number range from 0 to 4. The results clearly elucidate the competitive regime of PDE application relative to ramjets and gas turbines.

Description: This effort extends into high frequency (〉500 Hz), an earlier developed adaptive control algorithm for the suppression of thermo-acoustic instabilities in a liquidfueled combustor. The earlier work covered the development of a controls algorithm for the suppression of a low frequency (~280 Hz) combustion instability based on simulations, with no hardware testing involved. The work described here includes changes to the simulation and controller design necessary to control the high frequency instability, augmentations to the control algorithm to improve its performance, and finally hardware testing and results with an experimental combustor rig developed for the high frequency case. The Adaptive Sliding Phasor Averaged Control (ASPAC) algorithm modulates the fuel flow in the combustor with a control phase that continuously slides back and forth within the phase region that reduces the amplitude of the instability. The results demonstrate the power of the method - that it can identify and suppress the instability even when the instability amplitude is buried in the noise of the combustor pressure. The successful testing of the ASPAC approach helped complete an important NASA milestone to demonstrate advanced technologies for low-emission combustors.

Description: This research program focuses on characterizing the effect of impeller-diffuser interactions in a centrifugal compressor stage on its performance using unsteady threedimensional Reynolds-averaged Navier-Stokes simulations. The computed results show that the interaction between the downstream diffuser pressure field and the impeller tip clearance flow can account for performance changes in the impeller. The magnitude of performance change due to this interaction was examined for an impeller with varying tip clearance followed by a vaned or vaneless diffuser. The impact of unsteady impeller-diffuser interaction, primarily through the impeller tip clearance flow, is reflected through a time-averaged change in impeller loss, blockage and slip. The results show that there exists a tip clearance where the beneficial effect of the impeller-diffuser interaction on the impeller performance is at a maximum. A flow feature that consists of tip flow back leakage was shown to occur at design speed for the centrifugal compressor stage. This flow phenomenon is described as tip flow that originates in one passage, flows downstream of the impeller trailing edge and then returns to upstream of the impeller trailing edge of a neighboring passage. Such a flow feature is a source of loss in the impeller. A hypothesis is put forth to show that changing the diffuser vane count and changing impeller-diffuser gap has an analogous effect on the impeller performance. The centrifugal compressor stage was analyzed using diffusers of different vane counts, producing an impeller performance trend similar to that when the impeller-diffuser gap was varied, thus supporting the hypothesis made. This has the implication that the effect impeller performance associated with changing the impeller-diffuser gap and changing diffuser vane count can be described by the non-dimensional ratio of impeller-diffuser gap to diffuser vane pitch. A procedure is proposed and developed for isolating impeller passage blockage change without the need to define the region of blockage generation (which may incur a certain degree of arbitrariness). This method has been assessed for its applicability and utility.

Description: Improved blade tip sealing in the high pressure compressor and high pressure turbine can provide dramatic improvements in specific fuel consumption, time-on-wing, compressor stall margin and engine efficiency as well as increased payload and mission range capabilities of both military and commercial gas turbine engines. The preliminary design of a mechanically actuated active clearance control (ACC) system for turbine blade tip clearance management is presented along with the design of a bench top test rig in which the system is to be evaluated. The ACC system utilizes mechanically actuated seal carrier segments and clearance measurement feedback to provide fast and precise active clearance control throughout engine operation. The purpose of this active clearance control system is to improve upon current case cooling methods. These systems have relatively slow response and do not use clearance measurement, thereby forcing cold build clearances to set the minimum clearances at extreme operating conditions (e.g., takeoff, re-burst) and not allowing cruise clearances to be minimized due to the possibility of throttle transients (e.g., step change in altitude). The active turbine blade tip clearance control system design presented herein will be evaluated to ensure that proper response and positional accuracy is achievable under simulated high-pressure turbine conditions. The test rig will simulate proper seal carrier pressure and temperature loading as well as the magnitudes and rates of blade tip clearance changes of an actual gas turbine engine. The results of these evaluations will be presented in future works.

Description: The objective of this study was to demonstrate the high-fidelity numerical simulation of a modern high-bypass turbofan engine. The simulation utilizes the Numerical Propulsion System Simulation (NPSS) thermodynamic cycle modeling system coupled to a high-fidelity full-engine model represented by a set of coupled three-dimensional computational fluid dynamic (CFD) component models. Boundary conditions from the balanced, steady-state cycle model are used to define component boundary conditions in the full-engine model. Operating characteristics of the three-dimensional component models are integrated into the cycle model via partial performance maps generated automatically from the CFD flow solutions using one-dimensional meanline turbomachinery programs. This paper reports on the progress made towards the full-engine simulation of the GE90-94B engine, highlighting the generation of the high-pressure compressor partial performance map. The ongoing work will provide a system to evaluate the steady and unsteady aerodynamic and mechanical interactions between engine components at design and off-design operating conditions.

Description: Planetary exploration may be enhanced by the use of aircraft for mobility. This paper reviews the development of aircraft for planetary exploration missions at NASA and reviews the power and propulsion options for planetary aircraft. Several advanced concepts for aircraft exploration, including the use of in situ resources, the possibility of a flexible all-solid-state aircraft, the use of entomopters on Mars, and the possibility of aerostat exploration of Titan, are presented.

Description: Sense lines are used in pressure measurements to passively transmit information from hostile environments to areas where transducers can be used. The transfer function of a sense line can be used to obtain information about the measured environment from the protected sensor. Several properties of this transfer function are examined, including frequency dependence, Helmholtz resonance, and time of flight delay.

Description: A split-fiber probe was used to acquire unsteady data in a research compressor. The probe has two thin films deposited on a quartz cylinder 200 microns in diameter. A split-fiber probe allows simultaneous measurement of velocity magnitude and direction in a plane that is perpendicular to the sensing cylinder, because it has its circumference divided into two independent parts. Local heat transfer considerations indicated that the probe direction characteristic is linear in the range of flow incidence angles of +/- 35. Calibration tests confirmed this assumption. Of course, the velocity characteristic is nonlinear as is typical in thermal anemometry. The probe was used extensively in the NASA Glenn Research Center (GRC) low-speed, multistage axial compressor, and worked reliably during a test program of several months duration. The velocity and direction characteristics of the probe showed only minute changes during the entire test program. An algorithm was developed to decompose the probe signals into velocity magnitude and velocity direction. The averaged unsteady data were compared with data acquired by pneumatic probes. An overall excellent agreement between the averaged data acquired by a split-fiber probe and a pneumatic probe boosts confidence in the reliability of the unsteady content of the split-fiber probe data. To investigate the features of unsteady data, two methods were used: ensemble averaging and frequency analysis. The velocity distribution in a rotor blade passage was retrieved using the ensemble averaging method. Frequencies of excitation forces that may contribute to high cycle fatigue problems were identified by applying a fast Fourier transform to the absolute velocity data.

Description: Investigations of unsteady pressure loadings on the blades of fans operating near the stall flutter boundary are carried out under simulated conditions in the NASA Transonic Flutter Cascade facility (TFC). It has been observed that for inlet Mach numbers of about 0.8, the cascade flowfield exhibits intense low-frequency pressure oscillations. The origins of these oscillations were not clear. It was speculated that this behavior was either caused by instabilities in the blade separated flow zone or that it was a tunnel resonance phenomenon. It has now been determined that the strong low-frequency oscillations, observed in the TFC facility, are not a cascade phenomenon contributing to blade flutter, but that they are solely caused by the tunnel resonance characteristics. Most likely, the self-induced oscillations originate in the system of exit duct resonators. For sure, the self-induced oscillations can be significantly suppressed for a narrow range of inlet Mach numbers by tuning one of the resonators. A considerable amount of flutter simulation data has been acquired in this facility to date, and therefore it is of interest to know how much this tunnel self-induced flow oscillation influences the experimental data at high subsonic Mach numbers since this facility is being used to simulate flutter in transonic fans. In short, can this body of experimental data still be used reliably to verify computer codes for blade flutter and blade life predictions? To answer this question a study on resonance effects in the NASA TFC facility was carried out. The results, based on spectral and ensemble averaging analysis of the cascade data, showed that the interaction between self-induced oscillations and forced blade motion oscillations is very weak and can generally be neglected. The forced motion data acquired with the mistuned tunnel, when strong self-induced oscillations were present, can be used as reliable forced pressure fluctuations provided that they are extracted from raw data sets by an ensemble averaging procedure.

Description: This final report has been prepared by Honeywell Engines & Systems, Phoenix, Arizona, a unit of Honeywell International Inc., documenting work performed during the period June 1999 through December 1999 for the National Aeronautics and Space Administration (NASA) Glenn Research Center, Cleveland, Ohio, under the Small Engine Technology (SET) Program, Contract No. NAS3-27483, Task Order 24, Business and Regional Aircraft System Studies. The work performed under SET Task 24 consisted of evaluating the noise reduction benefits compared to the baseline noise levels of representative 1992 technology aircraft, obtained by applying different combinations of noise reduction technologies to five business and regional aircraft configurations. This report focuses on the selection of the aircraft configurations and noise reduction technologies, the prediction of noise levels for those aircraft, and the comparison of the noise levels with those of the baseline aircraft.

Description: The idea of using mixing enhancement to reduce jet noise is not new. Lobed mixers have been around since shortly after jet noise became a problem. However, these designs were often a post-design fix that rarely was worth its weight and thrust loss from a system perspective. Recent advances in CFD and some inspired concepts involving chevrons have shown how mixing enhancement can be successfully employed in noise reduction by subtle manipulation of the nozzle geometry. At NASA Glenn Research Center, this recent success has provided an opportunity to explore our paradigms of jet noise understanding, prediction, and reduction. Recent advances in turbulence measurement technology for hot jets have also greatly aided our ability to explore the cause and effect relationships of nozzle geometry, plume turbulence, and acoustic far field. By studying the flow and sound fields of jets with various degrees of mixing enhancement and subsequent noise manipulation, we are able to explore our intuition regarding how jets make noise, test our prediction codes, and pursue advanced noise reduction concepts. The paper will cover some of the existing paradigms of jet noise as they relate to mixing enhancement for jet noise reduction, and present experimental and analytical observations that support these paradigms.

Description: A method to intended for measurement of the insertion loss of an acoustic treatment applied to an aircraft fuselage in-situ is documented in this paper. Using this method, the performance of a treatment applied to a limited portion of an aircraft fuselage can be assessed even though the untreated fuselage also radiates into the cabin, corrupting the intensity measurement. This corrupting noise in the intensity measurement incoherent with the panel vibration of interest is removed by correlating the intensity to reference transducers such as accelerometers. Insertion loss of the acoustic treatments is estimated from the ratio of correlated intensity measurements with and without a treatment applied. In the case of turbulent boundary layer excitation of the fuselage, this technique can be used to assess the performance of noise control methods without requiring treatment of the entire fuselage. Several experimental studies and numerical simulations have been conducted, and results from three case studies are documented in this paper. Conclusions are drawn about the use of this method to study aircraft sidewall treatments.

Description: As part of the NASA/Navy Abrupt Wing Stall Program, a relatively low-cost, rapid-access wind-tunnel free-to-roll rig was developed. This rig combines the use of conventional models and test apparatuses to evaluate both transonic performance and wing-drop/rock tendencies in a single tunnel entry. A description of the test hardware as well as a description of the experimental procedures is given. The free-to-roll test rig has been used successfully to assess the static and dynamic characteristics of three different configurations--two configurations that exhibit uncommanded lateral motions, (pre-production F/A-18E and AV-8B), and one that did not (F/A-18C).

Description: The LaRC investigative process for airframe noise has proven to be a useful guide for elucidation of the physics of flow-induced noise generation over the last five years. This process, relying on a close interplay between experiment and computation, is described and demonstrated here on the archetypal problem of flap-edge noise. Some detailed results from both experiment and computation are shown to illustrate the process, and a description of the multi-source physics seen in this problem is conjectured.

Description: Results from an experimental investigation of flow field generated by pitched and yawed jets discharging from a flat plate into a cross-flow are presented. The circular jet was pitched at alpha = 20deg and 45deg and yawed between Beta = 0deg and 90deg in increments of 15deg. The measurements were performed with two -wires providing all three components of velocity and turbulent stresses. These data were obtained at downstream locations of x = 3, 5, 10 and 20, where the distance x, normalized by the jet diameter, is measured from the center of the orifice. Data for all configurations were acquired at a momentum-flux ratio J = 8. Additionally, for selected angles and locations, surveys were conducted for J = 1.5, 4, and 20. As expected, the jet penetration is found to be higher at larger alpha. With increasing beta the jet spreads more. The rate of reduction of peak streamwise vorticity, max, with the downstream distance is significantly less at higher Beta but is found to be practically independent of alpha. Thus, at the farthest measurement station x = 20, xmax is about five times larger for Beta = 75deg compared to the levels at Beta = 0deg. Streamwise velocity within the jet-vortex structure is found to depend on the parameter J. At J = 1.5 and 4, 'wake-like' velocity profiles are observed. In comparison, a 'jet-like' overshoot is present at higher J. Distributions of turbulent stresses for various cases are documented. Peak normal stresses are found to occur within the core of the streamwise vortices. With yaw, at lower values of J, high turbulence is also observed in the boundary layer underneath the jet-vortex structure

Description: A multi grid solution procedure for the numerical simulation of turbulent flows in complex geometries has been developed. A Full Multigrid-Full Approximation Scheme (FMG-FAS) is incorporated into the continuity and momentum equations, while the scalars are decoupled from the multi grid V-cycle. A standard kappa-Epsilon turbulence model with wall functions has been used to close the governing equations. The numerical solution is accomplished by solving for the Cartesian velocity components either with a traditional grid staggering arrangement or with a multiple velocity grid staggering arrangement. The two solution methodologies are evaluated for relative computational efficiency. The solution procedure with traditional staggering arrangement is subsequently applied to calculate the flow and temperature fields around a model Short Take-off and Vertical Landing (STOVL) aircraft hovering in ground proximity.

Description: Processes of soot formation and oxidation must be understood in order to achieve reliable computational combustion calculations for nonpremixed (diffusion) flames involving hydrocarbon fuels. Motivated by this observation, the present investigation extended earlier work on soot formation and oxidation in laminar jet ethylene/air and methane/oxygen premixed and acetylene-nitrogen/air diffusion flames at atmospheric pressure in this laboratory, emphasizing soot surface growth and early soot surface oxidation in laminar diffusion flames fueled with a variety of hydrocarbons at pressures in the range 0.1 - 1.0 atm.

Description: The Cool Flame Experiment aims to address the role of diffusive transport on the structure and the stability of gas-phase, non-isothermal, hydrocarbon oxidation reactions, cool flames and auto-ignition fronts in an unstirred, static reactor. These reactions cannot be studied on Earth where natural convection due to self-heating during the course of slow reaction dominates diffusive transport and produces spatio-temporal variations in the thermal and thus species concentration profiles. On Earth, reactions with associated Rayleigh numbers (Ra) less than the critical Ra for onset of convection (Ra(sub cr) approx. 600) cannot be achieved in laboratory-scale vessels for conditions representative of nearly all low-temperature reactions. In fact, the Ra at 1g ranges from 10(exp 4) - 10(exp 5) (or larger), while at reduced-gravity, these values can be reduced two to six orders of magnitude (below Ra(sub cr)), depending on the reduced-gravity test facility. Currently, laboratory (1g) and NASA s KC-135 reduced-gravity (g) aircraft studies are being conducted in parallel with the development of a detailed chemical kinetic model that includes thermal and species diffusion. Select experiments have also been conducted at partial gravity (Martian, 0.3gearth) aboard the KC-135 aircraft. This paper discusses these preliminary results for propane-oxygen premixtures in the low to intermediate temperature range (310- 350 C) at reduced-gravity.

Description: The Electric Particulate Suspension (EPS) is a combustion ignition system being developed at Iowa State University for evaluating quenching effects of powders in microgravity (quenching distance, ignition energy, flammability limits). Because of the high cloud uniformity possible and its simplicity, the EPS method has potential for "benchmark" design of quenching flames that would provide NASA and the scientific community with a new fire standard. Microgravity is expected to increase suspension uniformity even further and extend combustion testing to higher concentrations (rich fuel limit) than is possible at normal gravity. Two new combustion parameters are being investigated with this new method: (1) the particle velocity distribution and (2) particle-oxidant slip velocity. Both walls and (inert) particles can be tested as quenching media. The EPS method supports combustion modeling by providing accurate measurement of flame-quenching distance as a parameter in laminar flame theory as it closely relates to characteristic flame thickness and flame structure. Because of its design simplicity, EPS is suitable for testing on the International Space Station (ISS). Laser scans showing stratification effects at 1-g have been studied for different materials, aluminum, glass, and copper. PTV/PIV and a leak hole sampling rig give particle velocity distribution with particle slip velocity evaluated using LDA. Sample quenching and ignition energy curves are given for aluminum powder. Testing is planned for the KC-135 and NASA s two second drop tower. Only 1-g ground-based data have been reported to date.

Description: The problem considered is that of a single-component liquid fuel (n-heptane) droplet undergoing evaporation and combustion in a hot, convective, low pressure, zero-gravity environment of infinite expanse. For a moving droplet, the relative velocity (U(sub infinity)) between the droplet and freestream is subject to change due to the influence of the drag force on the droplet. For a suspended droplet, the relative velocity is kept constant. The governing equations for the gas-phase and the liquid-phase consist of the unsteady, axisymmetric equations of mass, momentum, species (gas-phase only) and energy conservation. Interfacial conservation equations are employed to couple the two phases. Variable properties are used in the gas- and liquid-phase. Multicomponent diffusion in the gas-phase is accounted for by solving the Stefan-Maxwell equations for the species diffusion velocities. A one-step overall reaction is used to model the combustion. The governing equations are discretized using the finite volume and SIMPLEC methods. A colocated grid is adopted. Hyperbolic tangent stretching functions are used to concentrate grid points near the fore and aft lines of symmetry and at the droplet surface in both the gas- and liquid-phase. The discretization equations are solved using the ADI method with the TDMA used on each line of the two alternating directions. Iterations are performed within each time-step until convergence is achieved. The grid spacing, size of the computational domain and time-step were tested to ensure that all solutions are independent of these parameters. A detailed discussion of the numerical model is given.

Description: Oxygen-enhanced combustion permits certain benefits and flexibility that are not otherwise available in the design of practical combustors, as discussed by Baukal. The cost of pure and enriched oxygen has declined to the point that oxygen-enhanced combustion is preferable to combustion in air for many applications. Carbon sequestration is greatly facilitated by oxygen enrichment because nitrogen can be eliminated from the product stream. For example, when natural gas (or natural gas diluted with CO2) is burned in pure oxygen, the only significant products are water and CO2. Oxygen-enhanced combustion also has important implications for soot formation, as explored in this work. We propose that soot inception in nonpremixed flames requires a region where C/O ratio, temperature, and residence time are above certain critical values. Soot does not form at low temperatures, with the threshold in nonpremixed flames ranging from about 1250-1650 K, a temperature referred to here as the critical temperature for soot inception, Tc. Soot inception also can be suppressed when residence time is short (equivalently, when the strain rate in counterflow flames is high). Soot induction times of 0.8-15 ms were reported by Tesner and Shurupov for acetylene/nitrogen mixtures at 1473 K. Burner stabilized spherical microgravity flames are employed in this work for two main reasons. First, this configuration offers unrestricted control over convection direction. Second, in steady state these flames are strain-free and thus can yield intrinsic sooting limits in diffusion flames, similar to the way past work in premixed flames has provided intrinsic values of C/O ratio associated with soot inception limits.

Description: Combustion of solid fuel particles has many important applications, including power generation and space propulsion systems. The current models available for describing the combustion process of these particles, especially porous solid particles, include various simplifying approximations. One of the most limiting approximations is the lumping of the physical properties of the porous fuel with the heterogeneous chemical reaction rate constants [1]. The primary objective of the present work is to develop a rigorous modeling approach that could decouple such physical and chemical effects from the global heterogeneous reaction rates. For the purpose of validating this model, experiments with porous graphite particles of varying sizes and porosity are being performed under normal and micro gravity.

Description: Diffusive-thermal instabilities are well known features of premixed and diffusion flames. In one of its form the instability appears as spontaneous oscillations. In premixed systems oscillations are predicted to occur when the effective Lewis number, defined as the ratio of the thermal diffusivity of the mixture to the mass diffusivity of the deficient component, is sufficiently larger than one. Oscillations would therefore occur in mixtures that are deficient in the less mobile reactant, namely in lean hydrocarbon-air or rich hydrogen-air mixtures. The theoretical predictions summarized above are in general agreement with experimental results; see for example [5] where a jet configuration was used and experiments were conducted for various inert-diluted propane and methane flames burning in inert-diluted oxygen. Nitrogen, argon and SF6 were used as inert in order to produce conditions of substantially different Lewis numbers and mixture strength. In accord with the predicted trend, it was found that oscillations arise at near extinction conditions, that for oscillations to occur it suffices that one of the two Lewis numbers be sufficiently large, and that oscillations are more likely to be observed when is relatively large.

Description: The present experimental study of soot processes in hydrocarbon-fueled laminar nonbuoyant and nonpremixed (diffusion) flames at microgravity within a spacecraft was motivated by the relevance of soot to the performance of power and propulsion systems, to the hazards of unwanted fires, and to the emission of combustion-generated pollutants. Soot processes in turbulent flames are of greatest practical interest, however, direct study of turbulent flames is not tractable because the unsteadiness and distortion of turbulent flames limit available residence times and spatial resolution within regions where soot processes are important. Thus, laminar diffusion flames are generally used to provide more tractable model flame systems to study processes relevant to turbulent diffusion flames, justified by the known similarities of gas-phase processes in laminar and turbulent diffusion flames, based on the widely-accepted laminar flamelet concept of turbulent flames. Unfortunately, laminar diffusion flames at normal gravity are affected by buoyancy due to their relatively small flow velocities and, as discussed next, they do not have the same utility for simulating the soot processes as they do for simulating the gas phase processes of turbulent flames.

Description: Studies of soot oxidation have ranged from in situ flame studies to shock tubes to flow reactors. Each of these systems possesses particular advantages and limitations related to temperature, time and chemical environments. Despite the aforementioned differences, these soot oxidation investigations share three striking features. First and foremost is the wide variation in the rates of oxidation. Reported oxidation rates vary by factors of +6 to - 20 relative to the Nagle Strickland-Constable (NSC) rate for graphite oxidation [3]. Rate variations are not surprising, as the temperatures, residence times, types of oxidants and methods of oxidation differ from study to study. Nevertheless, a valid explanation for rate differences of this magnitude has yet to be presented.

Description: Pulse detonation engines (PDB) have generated considerable research interest in recent years as a chemical propulsion system potentially offering improved performance and reduced complexity compared to conventional gas turbines and rocket engines. The detonative mode of combustion employed by these devices offers a theoretical thermodynamic advantage over the constant-pressure deflagrative combustion mode used in conventional engines. However, the unsteady blowdown process intrinsic to all pulse detonation devices has made realistic estimates of the actual propulsive performance of PDES problematic. The recent review article by Kailasanath highlights some of the progress that has been made in comparing the available experimental measurements with analytical and numerical models.

Description: A study was conducted to identify engine cycle and technologies needed for a regional aircraft which could be capable of achieving a 10 EPNdB reduction in community noise level relative to current FAR36 Stage 3 limits. The study was directed toward 100-passenger regional aircraft with engine configurations in the 15,000 pound thrust class. The study focused on Ultra High Bypass Ratio (UHBR) cycles due to low exhaust jet velocities and reduced fan tip speeds. The baseline engine for this study employed a gear-driven, 1000 ft/sec tip speed fan and had a cruise bypass ratio of 14:1. A revised engine configuration employing fan and turbine design improvements are predicted to be 9.2 dB below current takeoff limits and 12.8 dB below current approach limits. An economic analysis was also done by estimating Direct Operating Cost (DOC).

Description: Our role in this project was to participate in the design of the signal processing suite to analyze plasma density measurements on board a small constellation (3 or 4) satellites in Low Earth Orbit. As we are new to space craft experiments, one of the challenges was to simply gain understanding of the quantity of data which would flow from the satellites, and possibly to interact with the design teams in generating optimal sampling patterns. For example, as the fleet of satellites were intended to fly through the same volume of space (displaced slightly in time and space), the bulk plasma structure should be common among the spacecraft. Therefore, an optimal, limited bandwidth data downlink would take advantage of this commonality. Also, motivated by techniques in ionospheric radar, we hoped to investigate the possibility of employing aperiodic sampling in order to gain access to a wider spatial spectrum without suffering aliasing in k-space.

Description: A study was conducted to identify and evaluate noise reduction technologies for advanced ducted prop propulsion systems that would allow increased capacity operation and result in an economically competitive commercial transport. The study investigated the aero/acoustic/structural advancements in fan and nacelle technology required to match or exceed the fuel burned and economic benefits of a constrained diameter large Advanced Ducted Propeller (ADP) compared to an unconstrained ADP propulsion system with a noise goal of 5 to 10 EPNDB reduction relative to FAR 36 Stage 3 at each of the three measuring stations namely, takeoff (cutback), approach and sideline. A second generation ADP was selected to operate within the maximum nacelle diameter constrain of 160 deg to allow installation under the wing. The impact of fan and nacelle technologies of the second generation ADP on fuel burn and direct operating costs for a typical 3000 nm mission was evaluated through use of a large, twin engine commercial airplane simulation model. The major emphasis of this study focused on fan blade aero/acoustic and structural technology evaluations and advanced nacelle designs. Results of this study have identified the testing required to verify the interactive performance of these components, along with noise characteristics, by wind tunnel testing utilizing and advanced interaction rig.

Description: The National Aeronautics and Space Administration (NASA), Langley Research Center (LaRC), the Environmental Measurement and Modeling Division of the United States Department of Transportation s John A. Volpe National Transportation Systems Center (Volpe), and several other organizations (see Appendix A for a complete list of participating organizations and individuals) conducted a noise measurement study at NASA s Wallops Flight Facility (Wallops) near Chincoteague, Virginia during September 2000. This test was intended to determine engine installation effects on four civil transport airplanes: a Boeing 767-400, a McDonnell-Douglas DC9, a Dassault Falcon 2000, and a Beechcraft King Air. Wallops was chosen for this study because of the relatively low ambient noise of the site and the degree of control over airplane operating procedures enabled by operating over a runway closed to other uses during the test period. Measurements were conducted using a twenty microphone U-shaped array oriented perpendicular to the flight path; microphones were mounted such that ground effects were minimized and low elevation angles were observed.

Description: A framework for an effective computational methodology for characterizing the stability and the impact of distortion in high-speed multi-stage compressor is being developed. The methodology consists of using a few isolated-blade row Navier-Stokes solutions for each blade row to construct a body force database. The purpose of the body force database is to replace each blade row in a multi-stage compressor by a body force distribution to produce same pressure rise and flow turning. To do this, each body force database is generated in such a way that it can respond to the changes in local flow conditions. Once the database is generated, no hrther Navier-Stokes computations are necessary. The process is repeated for every blade row in the multi-stage compressor. The body forces are then embedded as source terms in an Euler solver. The method is developed to have the capability to compute the performance in a flow that has radial as well as circumferential non-uniformity with a length scale larger than a blade pitch; thus it can potentially be used to characterize the stability of a compressor under design. It is these two latter features as well as the accompanying procedure to obtain the body force representation that distinguish the present methodology from the streamline curvature method. The overall computational procedures have been developed. A dimensional analysis was carried out to determine the local flow conditions for parameterizing the magnitudes of the local body force representation of blade rows. An Euler solver was modified to embed the body forces as source terms. The results from the dimensional analysis show that the body forces can be parameterized in terms of the two relative flow angles, the relative Mach number, and the Reynolds number. For flow in a high-speed transonic blade row, they can be parameterized in terms of the local relative Mach number alone.

Description: Three forms of the high-frequency asymptotic Green's function for Lilley's equation are reviewed and compared to the exact solution over wide range of Strouhal numbers. The asymmetric approximation, which applies to sources away form the jet axis, and the quasi-symmetric approximation, which is arrived at by making a near-axis source assumption, are both obtained for parallel round jets from a formal Fourier-transform solution. The ray-theory solution, which is the only high-frequency approximation that can be applied to more general mean flows, follows from a WKB ansatz and is shown to be closely related to the asymmetric approximation. The comparisons show that the best overall prediction of the exact Green's function is given by the asymmetric approximation which remains accurate down to a Strouhal number of 1/2. The close relationship between the asymmetric and ray-theory approximations suggests that the high-frequency asymptotic Green's function for more general mean flows would be similarly successful.

Description: Fan noise continues to be a significant issue for commercial aircraft engines and there still exists a requirement for improved understanding of the fundamental issues associated with fan noise source mechanisms. At the present time, most of the prediction methods identify the dominant acoustic sources to be associated with the stator vanes or blade trailing edges which are downstream of the fan face. However recent studies have shown that acoustic waves are significantly attenuated as they propagate upstream through a rotor, and if the appropriate corrections are applied, sound radiation from the engine inlet is significantly underpredicted. The prediction models can only be applied to fans with subsonic tip speeds. In contrast, most aircraft engines have fan tip speeds which are transonic and this implies an even higher attenuation for upstream propagating acoustic waves. Consequently understanding how sound propagates upstream through the fan is an important, and not well understood phenomena. The objective of this study is to provide improved insight into the upstream propagation effects through a rotor which are relevant to full scale engines. The focus of this study is on broadband fan noise generated by boundary layer turbulence interacting with the trailing edges of the fan blades. If this source mechanism is important upstream of the fan, the sound must propagate upstream through a transonic non uniform flow which includes large gradients and non linearities. Developing acoustic propagation models in this type of flow is challenging and currently limited to low frequency applications, where the frequency is of the same order as the blade passing frequency of the fan. For trailing edge noise, much higher frequencies are relevant and so a suitable approach needs to be developed, which is not limited by an unacceptably large computational effort. In this study we are in the process of developing a computational method which applies for the high frequencies of interest, and allows for any type of flow field associated with the fan. In this progress report the approach to be used and the basic equations will be presented. Some initial results will be given, but these are preliminary and need further verification.

Description: A major challenge in the design and development of turbomachine airfoils for gas turbine engines is high cycle fatigue failures due to flutter and aerodynamically induced forced vibrations. In order to predict the aeroelastic response of gas turbine airfoils early in the design phase, accurate unsteady aerodynamic models are required. However, accurate predictions of flutter and forced vibration stress at all operating conditions have remained elusive. The overall objectives of this research program are to develop a transition model suitable for unsteady separated flow and quantify the effects of transition on airfoil steady and unsteady aerodynamics for attached and separated flow using this model. Furthermore, the capability of current state-of-the-art unsteady aerodynamic models to predict the oscillating airfoil response of compressor airfoils over a range of realistic reduced frequencies, Mach numbers, and loading levels will be evaluated through correlation with benchmark data. This comprehensive evaluation will assess the assumptions used in unsteady aerodynamic models. The results of this evaluation can be used to direct improvement of current models and the development of future models. The transition modeling effort will also make strides in improving predictions of steady flow performance of fan and compressor blades at off-design conditions. This report summarizes the progress and results obtained in the first year of this program. These include: installation and verification of the operation of the parallel version of TURBO; the grid generation and initiation of steady flow simulations of the NASA/Pratt&Whitney airfoil at a Mach number of 0.5 and chordal incidence angles of 0 and 10 deg.; and the investigation of the prediction of laminar separation bubbles on a NACA 0012 airfoil.

Description: The overall objective of the current effort at NASA GRC is to evaluate, develop, and apply methodologies suitable for modeling intra-engine trace chemical changes over post combustor flow path relevant to the pollutant emissions from aircraft engines. At the present time, the focus is the high pressure turbine environment. At first, the trace chemistry model of CNEWT were implemented into GLENN-HT as well as NCC. Then, CNEWT, CGLENN-HT, and NCC were applied to the trace species evolution in a cascade of Cambridge University's No. 2 rotor and in a turbine vane passage. In general, the results from these different codes provide similar features. However, the details of some of the quantities of interest can be sensitive to the differences of these codes. This report summaries the implementation effort and presents the comparison of the No. 2 rotor results obtained from these different codes. The comparison of the turbine vane passage results is reported elsewhere. In addition to the implementation of trace chemistry model into existing CFD codes, several pre/post-processing tools that can handle the manipulations of the geometry, the unstructured and structured grids as well as the CFD solutions also have been enhanced and seamlessly tied with NCC, CGLENN-HT, and CNEWT. Thus, a complete CFD package consisting of pre/post-processing tools and flow solvers suitable for post-combustor intra-engine trace chemistry study is assembled.

Description: The X-43C Flight Demonstrator Project is a joint NASA-USAF hypersonic propulsion technology flight demonstration project that will expand the hypersonic flight envelope for air-breathing engines. The Project will demonstrate sustained accelerating flight through three flights of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs will be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA's Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center off the coast of California over water in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration (approximately 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavy-weight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 (approximately 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air-Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides the background for NASA's current hypersonic flight demonstration efforts.

Description: For the past several years, NASA's Marshall Space Flight Center (MSFC) has been working with Plasma Processes, Inc. (PPI) to fabricate combustion chamber liners using the Vacuum Plasma Spray (VPS) process. Multiple liners of a variety of shapes and sizes have been created. Each liner has been fabricated with GRCop-84 (a copper alloy with chromium and niobium) and a functional gradient coating (FGC) on the hot wall. While the VPS process offers versatility and a reduced fabrication schedule, the material system created with VPS allows the liners to operate at higher temperatures, with maximum blanch resistance and improved cycle life. A subscal unit (5K lbf thrust class) is being cycle tested in a LOX/Hydrogen thrust chamber assembly at MSFC. To date, over 75 hot-fire tests have been accumulated on this article. Tests include conditions normally detrimental to conventional materials, yet the VPS GRCop-84 liner has yet to show any signs of degradation. A larger chamber (15K lbf thrust class) has also been fabricated and is being prepared for hot-fire testing at MSFC near the end of 2003. Linear liners have been successfully created to further demonstrate the versatility of the process. Finally, scale up issues for the VPS process are being tackled with efforts to fabricate a full size, engine class liner. Specifically, a liner for the SSME's Main Combustion Chamber (MCC) has recently been attempted. The SSME size was chosen for convenience, since its design was readily available and its size was sufficient to tackle specific issues. Efforts to fabricate these large liners have already provided valuable lessons for using this process for engine programs. The material quality for these large units is being evaluated with destructive analysis and these results will be available by the end of 2003.

Description: A future upgrade to the Space Shuttle Main Engine (SSME) may be the replacement of the current regenerative cooled tube-wall nozzle with a nozzle using a regeneratively-cooled channel-wall design. The current tube-wall design represents the only major piece of SSME hardware that has not been dramatically updated throughout thc long history of the engine. There are a number of advantages to a channel-wall design including the promise of faster and lower cost fabrication and greater reliability in the field. The technical obstacles in the path of making this happen are many, particularly in the realms of metallurgy and manufacturing techniques. However, one technical area that can and should be addressed in the near term as part of the development of detailed component requirements is a systems type model of the fluid flow and heat transfer processes to which the new design will be exposed. This paper presents the results of an effort to develop a mathematical model of the internal flow for a generic channel-wall nozzle functioning as a direct replacement for the current tube-wall nozzle with a minimum of systems-level changes. Comparisons will be made to mathematical modeling results for the current tube-wall design and the results of various geometrical trade studies will be presented. It is the intent of this work to examine the feasibility of the concept of a direct replacement component with minimum systems-!eve impacts and to highlight potential areas of concern requiring further work in the future.

Description: Results from a series of experiments to investigate techniques for extending the stable flow range of a centrifugal compressor are reported. The research was conducted in a high-speed centrifugal compressor at the NASA Glenn Research Center. The stabilizing effect of steadily flowing air-streams injected into the vaneless region of a vane-island diffuser through the shroud surface is described. Parametric variations of injection angle, injection flow rate, number of injectors, injector spacing, and injection versus bleed were investigated for a range of impeller speeds and tip clearances. Both the compressor discharge and an external source were used for the injection air supply. The stabilizing effect of flow obstructions created by tubes that were inserted into the diffuser vaneless space through the shroud was also investigated. Tube immersion into the vaneless space was varied in the flow obstruction experiments. Results from testing done at impeller design speed and tip clearance are presented. Surge margin improved by 1.7 points using injection air that was supplied from within the compressor. Externally supplied injection air was used to return the compressor to stable operation after being throttled into surge. The tubes, which were capped to prevent mass flux, provided 9.3 points of additional surge margin over the baseline surge margin of 11.7 points.

Description: An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Description: The effect of a nonuniform mean flow on the normal modes; the inflow/outflow nonreflecting boundary conditions; and the sound power are studied. The normal modes in an annular duct are computed using a spectral method in combination with a shooting method. The swirl causes force imbalance which couples the acoustic and vortical modes. The acoustic modes are distinguished from the vortical modes by their large pressure and small vorticity content. The mean swirl also produces a Doppler shift in frequency. This results in more counter-spinning modes cut-on at a given frequency than modes spinning with the swirl. Nonreflecting boundary conditions are formulated using the normal mode solutions. The inflow/outflow boundary conditions are implemented in a linearized Euler scheme and validated by computing the propagation of acoustic and vortical waves in a duct for a variety of swirling mean flows. Numerical results show that the evolution of the vortical disturbances is sensitive to the inflow conditions and the details of the wake excitations. All three components of the wake velocity must be considered to correctly compute the wake evolution and the blade upwash. For high frequencies, the acoustic-vortical mode coupling is weak and a conservation equation for the acoustic energy can be derived. Sound power calculations show significant mean flow swirl effects, but mode interference effects are small.

Description: An acoustic monitoring device has at least two acoustic sensors with a triggering mechanism and a multiplexing circuit. After the occurrence of a triggering event at a sensor, the multiplexing circuit allows a recording component to record acoustic emissions at adjacent sensors. The acoustic monitoring device is attached to a solid medium to detect the occurrence of damage.

Description: Under the SET Program Task 4 - Regional Turboprop/Turbofan Engine Advanced Combustor Study, a total of ten low-emissions combustion system concepts were evaluated analytically for three different gas turbine engine geometries and three different levels of oxides of nitrogen (NOx) reduction technology, using an existing AlliedSignal three-dimensional (3-D) Computational Fluid Dynamics (CFD) code to predict Landing and Takeoff (LTO) engine cycle emission values. A list of potential Barrier Technologies to the successful implementation of these low-NOx combustor designs was created and assessed. A trade study was performed that ranked each of the ten study configurations on the basis of a number of manufacturing and durability factors, in addition to emissions levels. The results of the trade study identified three basic NOx-emissions reduction concepts that could be incorporated in proposed follow-on combustor technology development programs aimed at demonstrating low-NOx combustor hardware. These concepts are: high-flow swirlers and primary orifices, fuel-preparation cans, and double-dome swirlers.

Description: Development of good predictive models for jet noise has always been plagued by the difficulty in obtaining good quality data over a wide range of conditions in different facilities.We consider such issues very carefully in selecting data to be used in developing our model. Flight effects are of critical importance, and none of the means of determining them are without significant problems. Free-jet flight simulation facilities are very useful, and can provide meaningful data so long as they can be analytically transformed to the flight frame of reference. In this report we show that different methodologies used by NASA and industry to perform this transformation produce very different results, especially in the rear quadrant; this compels us to rely largely on static data to develop our model, but we show reasonable agreement with simulated flight data when these transformation issues are considered. A persistent problem in obtaining good quality data is noise generated in the experimental facility upstream of the test nozzle: valves, elbows, obstructions, and especially the combustor can contribute significant noise, and much of this noise is of a broadband nature, easily confused with jet noise. Muffling of these sources is costly in terms of size as well as expense, and it is particularly difficult in flight simulation facilities, where compactness of hardware is very important, as discussed by Viswanathan (Ref. 13). We feel that the effects of jet density on jet mixing noise may have been somewhat obscured by these problems, leading to the variable density exponent used in most jet noise prediction procedures including our own. We investigate this issue, applying Occam s razor, (e.g., Ref. 14), in a search for the simplest physically meaningful model that adequately describes the observed phenomena. In a similar vein, we see no reason to reject the Lighthill approach; it provides a very solid basis upon which to build a predictive procedure, as we believe we demonstrate in this report. Another feature of our approach is that the analyses are all conducted with lossless spectra, rather than Standard Day spectra, as is often done in industry. We feel that it is important to isolate the effects of as many physical processes as practical. Atmospheric attenuation can then be included using the relations developed for NASA by Shields and Bass (Ref. 15), which are available in both FOOTPR and ANOPP. The current approach to coannular jet noise prediction used in FOOTPR is reported in Reference 16, which updates the earlier conventional-velocity-profile (CVP, Ref. 17) and inverted-velocity-profile (IVP, Ref. 18) models.

Description: A simple model of thrust augmentation from a pulsed source is described. In the model it is assumed that the flow into the ejector is quasi-steady, and can be calculated using potential flow techniques. The velocity of the flow is related to the speed of the starting vortex ring formed by the jet. The vortex ring properties are obtained from the slug model, knowing the jet diameter, speed and slug length. The model, when combined with experimental results, predicts an optimum ejector radius for thrust augmentation. Data on pulsed ejector performance for comparison with the model was obtained using a shrouded Hartmann-Sprenger tube as the pulsed jet source. A statistical experiment, in which ejector length, diameter, and nose radius were independent parameters, was performed at four different frequencies. These frequencies corresponded to four different slug length to diameter ratios, two below cut-off, and two above. Comparison of the model with the experimental data showed reasonable agreement. Maximum pulsed thrust augmentation is shown to occur for a pulsed source with slug length to diameter ratio equal to the cut-off value.

Description: This report summarizes progress on NASA Glenn Research Center Grant NAG3-2718 to the University of Missouri at Rolla This grant was awarded on February 22, 2002 and this report covers the performance period to September 30, 2002. There is considerable overlap in research effort with previous NASA Glenn Grant NAG3-2340, as the current effort represents a continuation and extension of this previous grant, which with a no cost supplement terminated on January 31, 2002. This report outlines progress on each task in the original proposal. In addition to progress on several of the specifically proposed tasks, considerable progress has been made in FEM algorithm development with the intent of introducing computational efficiencies required to model high frequency propagation and radiation and to open the possibility of expanding the scope of the modeling capability to three dimensional duct and nacelle geometries. Appended to this document is a paper presented at the 8th AIAA/CEAS Aeroacoustics Conference in June 2002. This paper overlaps the present grant and the previous grant identified above, and it is noted that this paper has also been appended to the final report for NAG3-2304.

Description: Interaction metrics developed for the process control industry are used to evaluate decentralized control of sound radiation from bays on an aircraft fuselage. The metrics are applied to experimentally measured frequency response data from a model of an aircraft fuselage. The purpose is to understand how coupling between multiple bays of the fuselage can destabilize or limit the performance of a decentralized active noise control system. The metrics quantitatively verify observations from a previous experiment, in which decentralized controllers performed worse than centralized controllers. The metrics do not appear to be useful for explaining control spillover which was observed in a previous experiment.

Description: This report provides an extensive analysis of potential wake vortex noise sources that might be utilized to aid in their tracking. Several possible mechanisms of aircraft vortex sound generation are examined on the basis of discrete vortex dynamic models and characteristic acoustic signatures calculated by application of vortex sound theory. It is shown that the most robust mechanisms result in very low frequency infrasound. An instability of the vortex core structure is discussed and shown to be a possible mechanism for generating higher frequency sound bordering the audible frequency range. However, the frequencies produced are still low and cannot explain the reasonably high-pitched sound that has occasionally been observed experimentally. Since the robust mechanisms appear to generate only very low frequency sound, infrasonic tracking of the vortices may be warranted.

Description: A forward swept fan, designated the Quiet High Speed Fan (QHSF), was tested in the NASA Glenn 9-by 15-foot Low Speed Wind Tunnel to investigate its noise reduction relative to a baseline fan of the same aerodynamic performance. The objective of the Quiet High Speed Fan was a 6 decibel reduction in the Effective Perceived Noise relative to the baseline fan at the takeoff condition. The intent of the Quiet High Speed Fan design was to provide both a multiple pure tone noise reduction from the forward sweep of the fan rotor and a rotor-stator interaction blade passing tone noise reduction from a leaned stator. The tunnel noise data indicted that the Quiet High Speed Fan was quieter than the baseline fan for a significant portion of the operating line and was 6 dB quieter near the takeoff condition. Although reductions in the multiple pure tones were observed, the vast majority of the EPNdB reduction was a result of the reduction in the blade passing tone and its harmonics. The baseline fan's blade passing tone was dominated by the rotor-strut interaction mechanism. The observed blade passing tone reduction could be the result of either the redesign of the Quiet High Speed Fan Rotor or the redesigned stator. The exact cause of this rotor-strut noise reduction, whether from the rotor or stator redesign, was not discernable from this experiment.

Description: A recently developed analytic result in acoustics, "Formulation 1B," is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Willliams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experimental results. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, by using both analytical and experimental data on the airfoil surface. The acoustic predictions are compared with analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

Description: The purpose of this paper is to show that computations with an aeroacoustic feedback loop, the jet screech noise, can be obtained using truly unstructured grid technology. Numerical results are presented for a nozzle with two different lip thicknesses which will be referred to in this paper as a thin and a thick lip nozzle respectively. The space-time conservation element and solution element (CE/SE) method is used to solve the conservation laws of the compressible axisymmetric Navier-Stokes equations. The equations are time marched to predict the unsteady flow and the near-field screech tone noise issuing from an underexpanded circular jet. The CE/SE method uses an unstructured grid based data structure. The unstructured grids for these calculations are generated based on the method of Delaunay triangulation. Comparisons of numerical results with available experimental data are shown for flows corresponding to several different jet Mach numbers. Generally good agreement is obtained in terms of flow physics, screech tone frequency, and sound pressure level.

Description: A general characteristic of photovoltaics is that they increase in efficiency as their operating temperature decreases. Based on this principal, the ability to increase a solar aircraft's performance by cooling the solar cells was examined. The solar cells were cooled by channeling some air underneath the cells and providing a convective cooling path to the back side of the array. A full energy balance and flow analysis of the air within the cooling passage was performed. The analysis was first performed on a preliminary level to estimate the benefits of the cooling passage. This analysis established a clear benefit to the cooling passage. Based on these results a more detailed analysis was performed. From this cell temperatures were calculated and array output power throughout a day period were determined with and without the cooling passage. The results showed that if the flow through the cooling passage remained laminar then the benefit in increased output power more than offset the drag induced by the cooling passage.

Description: The present NASA GRC-funded three-year research project is focused on studying PDE driven ejectors applicable to a hybrid Pulse Detonation/Turbofan Engine. The objective of the study is to characterize the PDE-ejector thrust augmentation. A PDE-ejector system has been designed to provide critical experimental data for assessing the performance enhancements possible with this technology. Completed tasks include demonstration of a thrust stand for measuring average thrust for detonation tube multi-cycle operation, and design of a 72-in.-long, 2.25-in.-diameter (ID) detonation tube and modular ejector assembly. This assembly will allow testing of both straight and contoured ejector geometries. Initial ejectors that have been fabricated are 72-in.-long-constant-diameter tubes (4-, 5-, and 6-in.-diameter) instrumented with high-frequency pressure transducers. The assembly has been designed such that the detonation tube exit can be positioned at various locations within the ejector tube. PDE-ejector system experiments with gaseous ethylene/ nitrogen/oxygen propellants will commence in the very near future. The program benefits from collaborations with Prof. Merkle of University of Tennessee whose PDE-ejector analysis helps guide the experiments. The present research effort will increase the TRL of PDE-ejectors from its current level of 2 to a level of 3.

Description: This Technical Memorandum (TM) discusses the harmonic and random plane acoustic waves propagating from inside a duct to its surroundings. Various duct surfaces are considered, such as rigid, flexible, and impedance. In addition, the effects of a mean flow are studied when the duct alone is considered. Results show a significant reduction in overall sound pressure levels downstream of the impedance wall for both mean flow and no mean flow cases and for a narrow duct. When a wider duct is used, the overall sound pressure level (OSPL) reduction downstream of the impedance wall is much smaller. In the far field, the directivity is such that the overall sound pressure level is reduced by about 5 decibels (dB) on the side of the impedance wall. When a flexible surface is used, the far field directivity becomes asymmetric with an increase in the OSPL on the side of the flexible surface of about 7 dB.