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In recent years a significant number of commercially available micro turbines have become available. At the same time unmanned aerial vehicles and smart munitions have decreased in size while their endurance needs have increased. With these new platform requirements comes the need for a propulsion system with reliability, good endurance and low acoustic signature. There has been much research accomplished in the area of steady cold flow primary sources, but little experimental work has been done using a gas turbine as a steady flow hot source. This investigation concerns the performance of an ejector driven by a small gas turbine. Aircraft applicability was a deciding factor in test geometry. Varying both engine throttle and the ejector?s downstream distance resulted in peak augmentation values of nearly 1.4., The original document contains color images.

ELECTRONIC FILE CHARACTERISTICS: 12 files; Adobe Acrobat (.PDF) and DVD video files. PHYSICAL DESCRIPTION: 1 DVD-ROM and 1 CD-ROM; 4 3/4 in.; 264 MB. SYSTEMS DETAIL NOTE: Run time of DVD is 4 min. 50 sec. ABSTRACT: The IPD Project is focused on rocket engine technologies that increase robustness, improve reliability, reduce hardware costs, reduce operations cost and increase performance. The Powerhead is the part of the rocket where propellant is pumped and conditioned prior to burning. The IPD uses the Full Flow Stage Combustion (FFSC) cycle, which provides a catalyst for engine life and reliability improvements 10 times greater than the Space Shuttle Main Engine. IPD uses hydrogen and oxygen as propellants to produce a sea-level thrust of 250,000 pounds with throttle capability. The technology in the engine components can be utilized for long-life reusable rocket engine systems, or for highly reliable, low-cost, expendable rocket engine systems. IPD tested four powerhead components in an unprecedented turbopump and preburner test campaign. Component data and technology will be applied to booster and upper stage engine development for years to come. IPD developed and demonstrated the FFSC cycle during integrated ground engine testing. The first new cycle developed in over 30 years - a historic achievement!, Project no. 50260651

A single-stage transonic research compressor and test rig are to be used to obtain data on the effect of inlet flow distortion on compressor (and therefore engine) stall. Auxiliary injection was examined as a technique for generating distortion in inlet stagnation pressure, or temperature, or to simulate the more complex effects of engine, steam ingestion from a catapult launch. Engineering analyses were developed and programmed in EES to relate inlet conditions to the compressor characteristics, for both pressure and temperature distortion. An injection duct area of 8% was selected to limit the required heater power. A CFD analysis was carried out to predict the compressor inlet flow field and hence position the injection duct exit. It was found that a broad range of distortion parameters could be generated by simply ducting (and heating) atmospheric air (or steam) through and auxiliary inlet throttle valve., The original document contains color images.

Currently, a TCP sender considers all losses as congestion signals and reacts to them by throttling its sending rate. With Internet becoming more heterogeneous with more and more wireless error-prone links, a TCP connection may unduly throttle its sending rate and experience poor performance over paths experiencing random losses unrelated to congestion. The problem of distinguishing congestion losses from random losses is particularly hard when congestion is light: congestion losses themselves appear to be random. The key idea is to "de-randomize" congestion losses. This paper proposes a simple biased queue management scheme that "de-randomizes" congestion losses and enables a TCP receiver to diagnose accurately the cause of a loss and inform the TCP sender to react appropriately. Bounds on the accuracy of distinguishing wireless losses and congestion losses are analytically established and validated through simulations. Congestion losses are identified with an accuracy higher than 95% while wireless losses are identified with an accuracy higher than 75%. A closed form is derived for the achievable improvement by TCP endowed with a discriminator with a given accuracy. Simulations confirm this closed form. TCP-Casablanca, a TCP-Newreno endowed with the proposed discriminator at the receiver, yields through simulations an improvement of more than 100% on paths with low levels of congestion and about 1% random wireless packet loss rates. TCP-Ifrane, a sender-based TCP-Casablanca yields encouraging performance improvement., Prepared in cooperation with the Dept. of Electrical and Computer Engineering, and Coordinated Science Lab, University of Illinois at Urbana-Champaign

The Center for Intelligent Machines and Robotics (CIMAR) at the University of Florida has been a leader in the field of autonomous vehicles since 1991. In order to continue these research activities at CIMAR, a new Kawasaki Mule All-Terrain Vehicle was chosen to be automated as a test-bed for the purpose of developing and testing autonomous vehicle technologies. This Mule will be referred to as the Navigational Test Vehicle II (NTV2), because it is the second generation of the Navigational Test Vehicle (NTV) that was first automated at CIMAR in 1991. This thesis describes in detail the design and implementation of the software and hardware systems necessary for converting the stock Kawasaki Mule ATV into an unmanned vehicle capable of remote control or fully autonomous operation. As part of the automation project, servo actuators were added to the throttle, brake, and steering systems in order to control the mobility functions of the Mule, allowing it to operate as an autonomous vehicle. The operational theory of servo control systems is discussed, and the throttle, brake, and steering actuator systems are described in detail. In addition, transfer function models have been developed to predict the transient response of each actuator control system. A description is given of the Joint Architecture for Unmanned Systems (JAUS), which was used as the vehicle control architecture. Because the Mule control software has been designed according to the modular ideology of the JAUS Reference Architecture, new software and hardware components can be readily implemented as part of the autonomous vehicle system, allowing it to be modernized as new technologies develop. Finally, the goals of this project were accomplished, and the Mule was converted into a reliable development vehicle for autonomous systems technology., The original document contains color images.

The ability to perform an inclination change maximizes the maneuverability of an orbiting space vehicle. Most maneuvers utilize a combined plane change and orbital transfer to the new orbit. This costs more in terms of energy and fuel than an in-plane change of orbits. The amount of DeltaV and fuel required for such an energy-intensive inclination change exceeds the benefit of performing the maneuver. However, this paper demonstrates that a winged re-entry vehicle, based on the currently proposed X-3 7, has the necessary thrust to change planes and then perform an in-plane transfer to achieve a new orbit. Using SIMULINKTM and LABVIEW simulation tools, this research found that the use of the aerodynamic lift of a winged re-entry vehicle produced more than 120 of inclination change with the minimal DeltaV achievable. Through small orbital maneuvers and atmospheric re-entry, the aerodynamics of the lift vector demonstrated that the spacecraft retained sufficient energy to prevent perigee collapse using an orbital regulation code to control throttle selling., The original document contains color images.

A two state, engine speed and wheel speed, powertrain model is developed with bondgraph theory and coded in the C programming language for real-time simulation and comparison to the same model implemented in the MATLAB Simulink/Stateflow environment. Both throttle and braking are inputs with adjustable engine, transmission, and vehicle parameters. The structure of the model implementation provides for decoupled transmission and driveline modules for modularity. Gear shifting is accomplished with a shift delay of 0.8 seconds. A fixed step Runge-Kutta integration scheme is used with a time step of 0.01 seconds. Two throttle input cases are compared for each model with the results being essentially the same for each case - peak differences are attributed to signal shift on transition edges. Simulation time for each case was 200 seconds and was achieved within 5 seconds real-time. This is a 40:1 simulation time to real-time ratio., X Mar 2002 to x Apr 2002, "The original document contains color images"

The concept of a control bifurcation grew out of an AFOSR PRET project to study the robust stabilization of axial flow compressors. Several versions of the Moore Greitzer compressor model were used in this study. McCaughan had studied the classical bifurcations that were present in the three dimensional Moore Greitzer equations(MG3) as the throttle and other parameters were varied. Liaw and Abed considered the throttle and other parameters as a control and developed feedback laws to change the criticality of the primary Hopf bifurcation in MG4.

This dissertation concerns itself with direct injection engine systems and, more specifically, with the examination of technical flow throttles, which are set into an injector in order to control the switch procedure. An optimization with regard to the switch process and efficiency is targeted, so that a particular profile of the throttles might be attained. As far as the throttle profile is concerned, modifications included relate to the diameter and length, a possible conical form of the throttle canal as well as a touching, to some degree, a rounding of the inlet and outlet edges of the throttle. The parameters employed subsequently succeed in simulations with various models, such as with oil, and then with water., Original contains color plates: All DTIC reproductions will be in black and white. Text in German; abstract in English.

Propulsion systems based on the pulsed detonation cycle offer the potential to provide increased performance while simultaneously reducing engine weight, cost, and complexity, relative to conventional propulsion systems currently in service. The increased performance potential is due to the high thermodynamic efficiency of detonation combustion. The near constant volume heat addition process of the detonation cycle, along with the lack of a compression cycle lend to the high efficiency and specific impulse, simplicity, and low-cost of pulse detonation propulsion systems. Pulse detonation Engines (PDEs) have the potential for operation at speeds ranging from static to high-supersonic (M<5), with competitive efficiencies enabling supersonic operation beyond conventional gas turbine engine technology. Currently, no single engine cycle exists that has such a broad range of operability. Pulse Detonation Rocket Engines (PDREs) have the potential to drastically reduce the cost of upper stage and orbit-transfer propulsion systems, and are also attractive for lunar and planetary exploration, and landers and excursion vehicles that require throttling for soft landing. This report discusses the thermodynamic basis of pulse detonation propulsion technology and identifies the major technology initiatives currently underway in the United States.

The present paper reports a study on the dynamic behaviour of a compression system based on a multi-stage centrifugal blower and provided with a device for the active suppression of surge instability. The control device includes a sensor of total pressure at the compressor inlet a proportional- derivative controller and an actuation valve at the plenum exit. A non-linear lumped- parameter model of the controlled compression system is presented. which considers also the time-lags in the compressor and actuator unsteady responses. The numerical simulation shows that: i) the predictions of control effectiveness provided by the non-linear model are rather different from the estimates of a linear approach mainly due to the abrupt stall which affects the compressor characteristics: ii) the present control device is capable of suppressing surge within almost the whole unstable operating range of the compressor with values of the proportional gain small enough to avoid actuator saturation (stroke- end occurrence); iii) the derivative component of the control law exerts a poor influence on the system stabilization and can be thus removed: iv) the proposed control strategy is effective over a wide range of stability parameter B which implies the possibility of suppressing surge also at the highest compressor speeds and when using large plenum volumes. On the basis of the numerical results, a practical control device has been designed. which consists of a transducer of differential pressure between plenum and compressor exit (equivalent to a sensor of compressor face total pressure, a butterfly throttle/actuation valve driven by a Stepper motor and a computer, provided with proper interfaces, for signal acquisition, data processing and motor control., Pres: Symposium RTO Applied Vehicle Tech Panel (AVT), Braunschweig, Germany,p15-1/15-12, 8-11 May 2000. This article is from ADA395700 Active Control Technology for Enhanced Performance Operational Capabilities of Military Aircraft, Land Vehicles and Sea Vehicles (Technologies des systemes a commandes actives pour l'amelioration des performances operationnelles des aeronefs militaires, des vehicules terrestres et des vehicules maritimes)

Emphasis in the period since April has been on observations and analysis of thermal structures associated with microturbulence and air/sea thermal processes at the sea surface. our method of observation is based on an imaging infrared camera that we have adapted to use from a small boat. The camera responds to radiation in the three to five micrometer wavelength range. This is a useful range because the intensity of black body radiation from the water surface in this range is highly sensitive to the temperature of the source. Typically temperature variations of a few hundredths of a degree are detectable. The camera is mounted on a boom extending forward from the bow of the boat about 1 meter above the water. Our first observations at sea were not satisfactory because sky radiation reflected from the wavy water surface obscured the water temperature signal. This difficulty has been removed by surrounding the camera with an aluminum plate shield that acts like a parasol to make an infrared shadow on the sea surface. With this apparatus in operation we have observed the microturbulent structures associated with evaporative cooling of the sea surface. The surface structures show warm pools typically a few centimeters in size, surrounded by narrow lines of cool water. Our interpretation is that the warm pools are broad areas of ascending water and the cool lines are descending. The overall shapes of these structures vary from drawn out figures resembling micro Langmuir cells to more nearly isotropic shapes. The change from isotropic to anisotropic shapes occurs rapidly and so far as we can observe, without relation to driving forces such as waves.

Two directions of this progrma have progressed: (1) Analysis of sea surface structure on the millimeter scale. Past analyses have focussed on the spectrum of waves. This approach does not provide full information because of the non-linear relations between short gravity waves and capillary waves. We are focusing on typical dispersed capillary trains and attempting to relate them to the sharp crested short gravity waves on which they ride. Some of the statistical features that need quantification are the rates of production and decay of the trains and the energy flux implied by the decay rates. (2) We are designing a system to estimate the heat flux from to ocean to the atmosphere in relation to the shortest waves. It is known that the existence of capillaries greatly increases various fluxes across the air-sea boundary, but it is not at all clear how this occurs. Several processes may be at work -- convergence/divergence of orbital motions in waves disturbs the boundary layer at the top of the water where heat flux is throttled by laminar flow. This will lead to variable temperatures at the various phases of capillary waves. Another process derives from the decay of capillary trains, This delivers horizontal momentum to the water in patches corresponding to the locations of the trains. Such a patchy driving force will encourage turbulent motions of a size corresponding to the length of the train. A third process is brought about by wind stress on short gravity waves and the capillaries.

Solid and liquid propellant rocket propulsion systems are weigh- and performance - driven machines whose components experience extreme environments. However, the structural integrity of both types of propulsion systems is currently monitored through type testing, labor-intensive x-ray analysis, and other NDE techniques. Smart structures technologies have the potential not only to alleviate some of the maintenance burden of current rocket propulsion systems but also improve the capability of future systems. This paper discusses the rocket design parameters and operational environments driving the application of passive structural health monitoring systems to both liquid propellant engines, with complex rotating machinery and severe thermal environments, and solid rocket motors, with long required silo lives of highly filled, chemically active, viscoelastic propellant. The potential for application of active smart structures technology to rocket systems will also be discussed including the control of solid propellant to permit throttling of solid motors and control of thermal shrinkage, vibration suppression, and improved creep resistance of rotating components and subassemblies for liquid rocket engines. While the rocket manufacturers are not currently developing smart structures technology for propulsion applications, the potential improvements in increased systems reliability, longer operational lifetimes, reduced weight, increased capability drive the need for research in this area.

A rudder assembly is attached to the nozzle of a jet-powered watercraft by a spring-loaded pivot so as to turn with the nozzle when the craft is steered. The rudder assembly has two flat plate rudders with a baffle plate attached perpendicularly between the rudders. The baffle plate partially covers the nozzle when the watercraft is at rest and is held in position by the spring-loaded pivot. At slow speeds the rudders are parallel to the water jet flow from the nozzle to aid in steering the watercraft. At higher speeds, the force of the water jet against the baffle plate overcomes the spring force to pivot the rudder assembly up and away from the nozzle such that steering is provided by directing the nozzle. In an imminent high-speed collision situation, the panic reaction is to shut off the throttle, which abruptly ends the jet flow from the nozzle. The spring-loaded pivot forces the baffle plate back over the nozzle and the rudders are again positioned parallel with the nozzle to provide steering without any flow from the nozzle.

Automated aircraft control has traditionally been divided into distinct functions that are implemented separately (e.g., autopilot, autothrottle, flight management); each function has its own fault-tolerant computer system, and dependencies among different functions are generally limited to the exchange of sensor and control data. A by-product of this federated architecture is that faults are strongly contained within the computer system of the function where they occur and cannot readily propagate to affect the operation of other functions. More modern avionics architectures contemplate supporting multiple functions on a single, shared, fault-tolerant computer system where natural fault containment boundaries are less sharply defined. Partitioning uses appropriate hardware and software mechanisms to restore strong fault containment to such integrated architectures. This report examines the requirements for partitioning, mechanisms for their realization, and issues in providing assurance for partitioning. Because partitioning shares some concerns with computer security, security models are reviewed and compared with the concerns of partitioning.

This study investigates the ability of the proposed Liquid Fly Back Booster, a replacement for the Space Shuttle's Solid Rocket Boosters (SRBs), to eliminate the need for the Return to Launch Site (RTLS) abort mode. A Fortran model of a nominal launch trajectory was perturbed to simulate a single Space Shuttle Main Engine-out (SSME-out) abort scenario at different times in the launch with the goal being to abort to the southern most landing site possible. The model accounts for vehicle lift, drag, dynamic pressure, throttling times and includes atmospheric effects to enhance fidelity. Results show the RTLS abort mode can be eliminated and successful landings made as far south as Savannah, Georgia. This success is attributed to the throttling capability of the Liquid Fly Back Booster engines.

An Electromagnetic valve actuation (EVA) system was developed and applied to a Kohler Command Series engine. The EVA system was developed for the Kohler engine by Aura Systems of El Segundo CA. Southwest Research Institute (SwRI) was responsible for evaluating the performance of the EVA equipped engine, running on natural gas, in a laboratory test environment. As part of this effort, SwRI applied a personal computer-based engine control system which managed the fueling, ignition, throttling, and intake/exhaust valve control functions. Advantages from utilizing the EVA system on the engine proved to be increased engine power and tall load efficiency at low speed operation, and increased part load efficiency at all engine speeds.

Completed research is reported for an dynamic numerical investigation of unsteady flow in supersonic and hypersonic aircraft inlets. An explicit dynamic solution adaptive mesh computational code was further developed and used to obtain dynamic solutions for an axisymmetric mixed compression inlet and a generic dual mode scramjet inlet isolator diffuser combination. To improve robustness, an existing implicit code was modified for time accuracy and the solution adaptive mesh algorithm was installed. The inlet unstart phenomenon was simulated through perturbation of freestream and downstream conditions (axisymmetric inlet) and through downstream throttling for the dual mode 3-D configuration. Axisymmetric unstart could be induced by a 10% freestream temperature increase or a 5% backpressure increase. Comparison of the results with experiment, where available, indicate that stability margins assessed through inviscid design or quasi-steady experiment may need revision when dynamics are considered. Conclusions are drawn concerning specifics of the flow phenomena and directions for future research are suggested. Lack of highly resolved dynamic experimental data is a pacing item and will prevent full verification of future work.

The Integrated Powerhead Demonstration (IPD) is a 1,112,000 N (250,000 Ib(sub f)) thrust (at sea level) LOX/LH2 demonstration of a full flow cycle in an integrated system configuration. Aerojet and Rocketdyne are on contract to the Air Force Research Laboratory to design, develop, and deliver the required components, and to provide test support to accomplish the demonstration. Rocketdyne is on contract to provide a fuel and oxygen turbopump, a gas-gas injector, and system engineering and integration. Aerojet is on contract to provide a fuel and oxygen preburner, a main combustion chamber, and a nozzle. The IPD components are being designed with Military Spaceplane (MSP) performance and operability requirements in mind. These requirements include: lifetime >/- 200 missions, mean time between overhauls >/- 100 cycles, and a capability to throttle from 20% to 100% of full power. These requirements bring new challenges both in designing and testing the components. This paper will provide some insight into these issues. Lessons learned from operating and supporting the space shuttle main engine (SSME) have been reviewed and incorporated where applicable. The IPD program will demonstrate phase I goals of the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program while demonstrating key propulsion technologies that will be available for MSP concepts. The demonstration will take place on Test Stand 2A at the Air Force Research Laboratory at Edwards AFB. The component tests will begin in 1999 and the integrated system tests will be completed in 2002., Prepared in cooperation with Aerojet, Sacramento, CA.

Thrust augmenting ejectors were analyzed by the author in an attempt to see if they could be used to provide a significant thrust increase over the baseline thrust of a primary core. This was done for the purpose of determining if leaving the ejectors open in the cruise configuration of the High Speed Civil Transport would lead to a thrust increase that would at least offset the weight of the ejectors themselves. If this was found to be true then the fuel economy of the HSCT might be improved by leaving them open and not closing them during cruise. In analyzing the ejectors, no assumptions were made regarding inlet nor outlet configurations, so an attempt was made to find the point of optimal thrust augmentation by varying secondary stream bypass Mach number and the amount of flow entrainment. Two solutions were found to each mixing scenario, one subsonic and the other supersonic. These two solutions were each analyzed and ones not satisfying the Second Law of Thermodynamics were eliminated. Analytic diffuser and bleed losses were also explored in the analysis of the ejector flow. Within the limitations of the assumptions discussed in this paper, appreciable thrust augmentations have been discovered over a large range of bypass Mach numbers and entrained mass flows. This lead the author to the conclusion that ejectors warrant further research beyond a first order analysis, and serious thought should be put into leaving them open in cruise.

Classical bifurcation analysis for nonlinear dynamics is used to derive a nonlinear feedback control law that eliminates the hysteresis loop associated with rotating stall and extends the stable operating range in axial compressors. The proposed control system employs pressure rise as output measurement and throttle position as actuating signal for which both sensor and actuator exist in the current configuration of axial compressors. Thus our results provide a practical solution for rotating stall control in axial compressors., Prepared in collaboration with Louisiana State Univ., Baton Rouge.

Flight simulation has historically been an expensive proposition, particularly if out-the-window views were desired. Advances in computer technology have allowed a modular, off-the-shelf flight simulation (based on 80486 processors or Pentiums) to be assembled that has been adapted, with minimal modification, for conducting general aviation research. This simulation includes variable flight instrumentation, forward, 45 and 90 degree left external world views, and a map display. Control inputs are provided by high-fidelity analog controls (e.g., damped and self-centering yoke, high-performance throttle quadrant, gear, flap, and trim controls; and navigation radio frequency select). The simulation is based upon two commercially available flight simulation software packages, one originally designed as an instrument flight tralner and the other as a 'game'-type flight simulation. The provisions of these packages are discussed highlighting their particular research capabilities. as well as their limitations. The comparatively low cost and ease of assembly/integration allow multiple 'standardized' systems to be distributed for cooperative inter-laboratory studies. The approach appears to have utility for both research and training.

An overview of space electric propulsion (SEP) is presented. Methods of throttling the power levels of electrostatic and electromagnetic thrusters are discussed. Particular attention is given to the concept of thermally-throttling propellant flow using the temperature-viscosity characteristics of xenon gas. The thermoproperties of xenon gas as a function of temperature are determined, and the flow regimes of the propellant at the mass flow rates of interest are studied. The propellant flow is presented separately as Fanno flow and as Rayleigh flow, and then those combined effects are considered. A method for predicting the performance of thermally-throttled systems is presented. Uncertainties in modeling real-world thermal throttling systems are discussed. The possible use of thermal throttling characteristics as a means of propellant pressure regulation is also examined.

An aircraft throttle simulator comprises throttle lever which extend upwardly from spheres fractionally clamped between plates. The levers are constrained for movement in a pattern allowing for forward movement from an idle position to a full throttle position, a lateral shift from the full throttle position, and forward movement to an afterburner position. A linkage is provided to connect each lever to a linear transducer to provide a position signal., Supersedes PAT-APPL-130 950-93.

This research evaluates the use of selective non-catalytic reduction (SNCR) for the removal of oxides of nitrogen (NOx) in a jet engine test cell (JETC). The plume effluent in the exhaust stream of a TF30-P111+ and a TF33-P9 engine in a jet engine test cell was first characterized. A test stand and five-point sample probe were designed and built to make measurements directly behind the engine nozzle. Data were collected in conjunction with actual engine tests. Temperature, NOx, carbon monoxide (CO) concentration, and velocity were among the characteristics measured radially and axially in the plume. These parameters were used to determine baseline conditions for the chemical kinetic modeling of SNCR using ammonia, cyanuric acid, urea, and hydrazine with additives of hydrogen, methane, and hydrogen peroxide. The measurements revealed that temperatures are marginal for the application of SNCR in JETCs except at high throttle and afterburner settings on the P111+ engine. The chemicals themselves have limited range of applicability. Additives help to extend their usefulness, mostly by shifting the temperature window down. Urea and hydrogen peroxide yielded the best range of the chemicals tested. However, this combination may reduce NOx in the entire exhaust stream by up to only 25% at military power on the high-performance P111+ engine. The findings of this work have attractive implications for non-JETC applications operating near the same conditions.

Construction and modification of the novel tubular flow reactor for the synthesis of nanocrystalline materials by thermal evaporation in a forced flux of gas has been an ongoing process over the past year. The first version of the reactor was used to synthesize nanocrystalline silicon (Si). The incoming gas flow rate is set with a mass flow controller. Helium (He) enters at one end of the reactor and flows over the resistively heated crucible where Si is evaporating. The silicon supply is continually replenished with a magnetically coupled loader. An IR pyrometer is used for temperature monitoring and control of the evaporation crucible. A pumping system consisting of a roots blower (booster pump) backed by a single-stage mechanical pump continuously removes the Si particles in the He gas stream from the growth zone over the evaporation source. The pressure in the reactor is controlled by a throttle valve with a capacitance manometer input. The particles entrained in the gas stream flow down the glass tube and are deposited on a liquid nitrogen cooled disc mounted parallel to the gas flow in the center of the collection chamber. jg p.1

The redesign, data acquisition system installation, and software development for the cold-flow Turbine Test Rig is reported. The turbine tested was the Space Shuttle Main Engine High Pressure Fuel Turbopump first stage 'Alternate Turbopump Development' model designed and manufactured by Pratt & Whitney. The purpose of - this research was to establish a continuously operating, cdd-flow, turbine test facility which could be used for future laser-Doppler-velocimetry measurements in the tip leakage region of the turbine rotor blades. Also, to determine both the single and two stage turbine performance parameters, and to provide a test model for comparison with numerical simulations using a three-dimensional viscous flow code. A throttle valve was designed and installed at the turbine exit to control the pressure ratio in order to map the turbine performance. Data acquisition, reduction and display programs were developed using LabVIEW Visual Instrument software in order to collect, and process the data. The first-stage turbine exit velocity profile was measured using a cobra probe. Also, the first-stage turbine performance characteristics were determined. Flow-field characteristics of the first stage turbine were predicted using a three-dimensional viscous flow code., Original contains color plates: All DTIC reproductions will be in black and white.

A throttle or fluid control valve uses at least one compressible multi-wave washer as a fluid control element. The apertures formed by the multi-wave washer are short and of low aspect ratio when the valve is fully open to limit pressure drop across the valve, thus substantially preventing cavitation and flow related noise to high flow velocities. As the valve is closed, the aspect ratio of the apertures increases but surfaces of the apertures remain angled to each other to provide noise cancellation and damping. Several multi-wave washers spaced by sliders are preferably used to further limit pressure drop in any stage and to form a serpentine path in the general direction of flow to provide further noise cancellation and damping. The sliders also enhance the self purging action of the valve structure. A closure structure is also provided for assuring full closure of the valve., Supersedes PAT-APPL-266 062-94.

A method is provided for guiding an in-flight vehicle having course divert capability toward an intercept with a target. Using a sensing device mounted on the vehicle, target data is acquired to establish a relative position vector between the vehicle and the target. A trajectory course for the vehicle is selected such that it is offset from a collision course with the target. The offset is defined within in a target impact plane perpendicular to the relative position vector. The total thrust of the vehicle associated with its course divert capability, is directionally controlled through first and second orthogonal components. To reduce offset the second component is directed to cause the trajectory of the vehicle to spiral around the relative position vector., Supersedes PAT-APPL-77 798-93.

A cascade orificial resistive device for throttling fluid flow which minimizes acoustic noise and internal vibrations is described herein. The device has a hollow body defining a fluid passageway, a plurality of perforated plates mounted within the passageway, a fixed end ring adjacent one end of the perforated plates, and a threadable end ring adjacent an opposite end of the perforated plates to place the plates in compression. Each of the perforated plates is a single piece molded plate having an integral outer ring and an integrally formed center keying mechanism as well as a plurality of orifices. The keying mechanism formed on each plate is designed so that adjacent ones of the plates have their orifices misaligned. In this manner, a pressure drop across each plate is created and the fluid flow through the device is throttled. The device of the present invention has utility in a number of onboard marine vessel systems wherein reduced acoustic noise and internal vibrations are particularly desirable., Supersedes PAT-APPL-899 520.

A rotary servo or throttle valve comprises a housing defining a chamber and a discharge port in fluid communication with the chamber, a rotary valve spool for delivering fluid into the chamber, and a system for delivering pressurized fluid passing from the outlet opening of the valve in the chamber to the discharge port in the housing in which non-cavitating technology can be readily incorporated. The outlet opening of the valve spool is a wedge shaped groove formed in its periphery, the groove having opposed, accurately diverging sidewalls. The diffusing system comprises an elongated member which may be rotated in the chamber to align a plurality of flow paths with the discharge port and the outlet opening. The elongated member and valve spool may be rotated independently in opposite directions to obtain faster flow rate changes., Supersedes PAT-APPL-953 389-92.

In Task D.6 of the Advanced Engine Study, three primary subtasks were accomplished: 1) Design and Parametric Data, 2) Engine Requirement Variation Studies, and 3) Vehicle Study/Engine Study Coordination. Parametric data were generated for vacuum thrusts ranging from 7500 lbf to 50000 lbf, nozzle expansion ratios from 600 to 1200, and engine mixture ratios from 5:1 to 7:1. the Failure Modes and Effects Analysis (FMEA) driven baseline design generated in Tasks D.4 and D.5 was used as a departure point for these parametric analyses. These data are intended to assist in vehicle definition and trade studies. In the Engine Requirements Variation Studies, the individual effects of increasing the throttling ratio from 10:1 to 20:1 and requiring the engine to operate at a maximum mixture ratio of 12:1 were determined. Off-design engine balances were generated at these extreme conditions and individual component operating requirements analyzed in detail. Potential problems were identified and possible solutions generated. In the Vehicle Study/Engine Study coordination subtask, vehicle contractor support was provided as needed, addressing a variety of issues uncovered during vehicle trade studies. This support was primarily provided during Technical Interchange Meetings (TIM) in which Space Exploration Initiative (SEI) studies were addressed.

The Vehicle Simulation software is the genesis of the distributed simulation concept. The first devices to interact over the SIMNET network were two MI vehicle simulators. Gradually the concept diversified to add a greater variety of vehicles, as well as network analysis tools, command and control simulators, and Semi-Automated forces. A vehicle simulator is composed of one or more crewstations. Each crewstation provides the subset of the controls and indicators that would be available to that specific crew member on the actual vehicle. For example, on the M1 vehicle simulator the driver is provided with the steering bar, throttle, and brake control necessary to drive the tank. A critical element of each crewstation is the out-the-window view, or vision block, associated with the station. These vision blocks provide the crew member with a three dimensional graphical representation of the terrain and objects in the environment external to the vehicle. The purpose of the vehicle simulation software is to maintain the vehicle specific model, respond to crew inputs, update crewstation outputs, and communication with the network., See also Volume 2, AD-A244 934.

The 'Programmable Cockpit' is a low-cost facility utilising personal computers linked together to represent the fundamental displays of a fixed-wing aircraft. The cockpit instruments can be displayed either in the conventional manner or in a 'glass-cockpit' type format. Aircraft controls include a sidestick and throttle. It was designed so that the instrument layouts and display formats could be reconfigured rapidly and tested in a reasonable aircraft representation, with the pilot under representative workload conditions. The Programmable Cockpit is to be used to study and develop the pilot vehicle interface for future aircraft systems. Keywords: Head up displays, Cockpit simulators, Australia.

Flight tests were conducted which evaluated the use of ethanol as a blending agent for turbine fuels. As part of this study, a Beech King Air 200 was modified to incorporate a cabin tank which contained ethanol, and the fuel system was modified to deliver the ethanol to the starboard engine. Ten percent ethanol was mixed with JP-4 in the project fuel line, just before entering the engine. The test was performed over a broad range of operating conditions, and key performance parameters were measured. The use of ethanol as a blending agent reduced the power available and resulted in an increase in fuel consumption. The increase in fuel consumption was greater than the decrease in energy density. When operating on the ethanol/JP-4 mixtures, the test pilots noted the test engine required more throttle (throttle stagger). Ethanol, Turbine engines, Alternate fuels, Flight tests.

High-response pressure measurements of a high-speed, 10-stage, axial- flow compressor operating in rotating stall are analyzed. Procedures used to digitize analog voltages and calibrate pressure transducers are presented. From total and static pressures measured at the exit of the test compressor, stall cell Mach number distributions are calculated and used to study the effects of discharge throttle levels and variable vane changes on the 10th-stage rotating stall cells. Results indicate that significant transition zones exist between the reverse flow and peak Mach number of the stall cell cycle. Since the axial Mach numbers of the stall cell cycle are constantly changing, the amount of leading and trailing edge transition zones and fully unstalled flow zones are not easily defined. A method is devised to approximate the different flow zone ranges and correlate them to in-stall pressure characteristic behavior of the 10th stage of the test compressor. Changes in the time-averaged pressure characteristics are found to correlate with changes in the rotating stall flow zones. A lower pressure coefficient appears to correspond to an increase in ratio of trailing to leading edge transition zone size and the average transition zone size. Results also suggest that recovery hysteresis in the test compressor is characterized by reverse flow in the rotating stall cell.

Bifurcation analysis was used to investigate the nonlinear behavior of a simple powered lifting hypersonic vehicle in circular orbit about a spherical nonrotating Earth with gradients in atmospheric density and pressure and an inverse square law for gravity. Vehicle motion is constrained to a vertical plane so only longitudinal dynamics were modeled. Bifurcation analysis was conducted using the AUTO software package. A simple five-state model with three different thrust laws was derived to describe an unaugmented vehicle whose geometric and aerodynamic characteristics follow those of the literature. A parameter representing a body flap deflection was used to conduct one set of bifurcation sweeps for each thrust law. A second set of bifurcation sweeps for each thrust law was obtained using a parameter representing a throttle which scaled the thrust. Secondary parameters representing simple feedback gains were subsequently added. Results were surprising for a simple system with basically linear aerodynamics . Periodic branches arising from the loss of pitch stability or associated with a resonance altitude are routinely found with significant amplitude, and periods on the order of an elliptical orbit's period for a given geocentric radius. Rotational states generally had sub-oscillations of greater frequency.

A pneumatic launcher for instrumented objects includes a barrel, a breech member having a compressed gas reservoir, a pneumatic piston operated valve for discharging gas from the reservoir into the barrel and is characterized by a fail-safe venting feature and a throttling valve for automatically limiting acceleration of the launched object. Patents. (jhd), Supersedes PAT-APPL-633 361-84.