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Thin-Film Thermocouples Hot Tested on Ceramic-Matrix Composite Hoop
Dual, Active, Surface Heat Flux Gage Probe Passes Test
Ratioing Multiwavelength Pyrometer Developed
Resistance Strain Gage Comparison-Tested Against High-Temperature Extensometry
Nonintrusive System Measures ASTOVL Hot Gas Ingestion in Wind Tunnel
Liquid-Crystal, Point-Diffraction Interferometer Invented
Performance-Limiting Micropipe Defects Identified in SiC Wafers
Fiber Optic Control System Integration Program Concluded
Silicon Carbide Junction Field-Effect Transistors Developed
Site-Competition Epitaxy Controls Doping of Silicon Carbide
Parallel Implementation of Real-Time Expert Diagnostic System Achieved
Closed-Loop Control Demonstrated for Rocket Engine Faults
Piloted Evaluation Performed for Integrated Propulsion and Airframe Control Design Methodology
Dynamic Wave Rotor Performance Modeled
Thin-film thermoples are used on the superalloy materials in jet aircraft engines. However, advanced propulsion systems presently being developed must withstand higher operating temperatures, spurring the development of thin-film thermocouples for the ceramic components of these engines.
NASA Lewis has used thin-film thermocouples to measure surface temperatures on a ceramic-matrix composite hoop developed by Allison Engine Co. under an Air Force contract to design material systems that meet propulsion capability goals set by the Integrated High-Performance Turbine Engine Technology (IHPTET) Program. The hoop material contains hydroply filiazane (HPZ) fibers in a silicon carbide matrix.
We fabricated three thin-film thermocouples on the hoop at NASA Lewis. The sensors had several layers: an electrically insulating layer, the thermocouple legs, and a protective overcoat. The electrical insulator between the substrate and the thermocouple was a thermally grown oxide layer enhanced with a 2- to 3-µm-thick layer of sputter-deposited aluminum oxide. The thermocouple was platinum-13% rhodium versus platinum sputtered deposited onto the aluminum oxide to an approximate thickness of 5 µm. An additional sputter-deposited layer of aluminum oxide protected the sensor elements against the harsh testing environment. Platinum-13% rhodium and platinum lead wires 75 µm in diameter were parallel gap welded to the films. The wires were spliced to larger diameter wires routed to the facility instrumentation panel.
The hoop was tested for 50 hr in the Lewis Ceramic Matrix Composite Test Facility, a burner rig that operates on jet fuel at 0.7 to 2 MPa (100 to 300 psig). For this test the facility was designed to run at heat flux levels simulating near-IHPTET conditions. The average gas stream temperature was approximately 1500 °ree;C (2800 °ree;F), and the average material temperature was 1200 °ree;C (2200 °ree;F). The thin-film sensors obtained suface temperature data for over 25 hr of testing-much longer than in previous attempts with wire thermocouples on this material in these conditions. As a result we have obtained a larger data base of surface temperature for this ceramic-matrix composite under harsh operating conditions. Thin-film thermocouples proved to be capable of operating on advanced materials under test conditions in which wire thermocouples were unable to survive.
Lewis contact: Lisa C. Martin, (216) 433-6468
Headquarters program office: OA
Dual, Active, Surface Heat Flux Gage Probe Passes Test
A unique miniature, dual, active, surface plug type of heat flux gage probe was successfully tested in NASA Ames Research Center's 5- by 23-Centimeter Turbulent Flow Air Duct Facility. Heat flux was provided by an arc jet that produces high free-stream temperatures in an extremely hostile environment. The probe was mounted in a fixture bolted to the duct wall and fabricated from 2.54-cm-thick thermal insulation material in an aluminum holder. The probe (designed and fabricated at NASA Lewis) was inserted through the metal and ceramic materials with the front of the gage flush with the ceramic surface. The back surface of the gage was actively and passively air cooled.
Heat fluxes measured by the miniature gage probe were compared with heat fluxes obtained by water-cooled reference calorimeters mounted in a water-cooled duct wall located opposite the miniature gage probe. Both transient and steady-state heat fluxes generally corresponded within a satisfactory ±20%, even though the front surface temperature of the gage was 850 K and the surrounding ceramic surface temperature was 1750 K. (The latter temperature exceeds the melting point of the gage but does not exceed the melting point of the ceramic.) Surprisingly, this large temperature disparity did not greatly affect the accuracy of the heat flux measurement.
These tests suggest that this sensor can be used to measure heat flux under simultaneous conditions of high free-stream temperatures and large temperature gradients along a wall. A possible application is for transient and steady-state heat flux measurements on the walls of advanced gas turbine and rocket engine apparatus.
Lewis contact: Curt H. Liebert, (216) 433-6483
Headquarters program office: OA
An object at absolute zero temperature does not emit thermal radiation, but at temperatures above absolute zero it will emit varying amounts, depending on the temperature, according to Planck's law of radiation. With suitable precalibrated detectors the emitted radiation is measured and used to determine temperature. This is pyrometry. It works most conveniently for blackbodies, but real objects are not black. They are called gray or nongray bodies because they radiate less efficiently than an ideal blackbody. Gray objects have constant emissivity at all wavelengths. Nongray objects have different emissivities at different wavelengths. (Wavelength is one of the characteristics of radiation.) Without knowing the emissivity of an object, pyrometry determines only an erroneous radiant temperature, not the actual temperature.
In application environments, in addition to radiation emitted from the measured surface, radiation originating from other sources and reflected from the measured surface is also detected. The temperature determined from the total radiation would thus be in error. Traditional pyrometers operating at one or two wavelength regions (one-color or two-color pyrometers) cannot correct for errors due to emissivity uncertainty and reflected radiation.
At NASA Lewis an in-house program has developed a new ratioing multiwavelength pyrometer (RMP) that measures temperature by detecting the complete spectra at many wavelengths. The recording times of two spectra are separated by short intervals during which a temperature excursion has occurred such that the emissivity can be considered to remain unchanged. A ratio spectrum is formed by a wavelength-by-wavelength division of these two spectra, whose temperatures are not known. The ratio spectrum does not depend on emissivity because it has been divided out. The equation representing the ratio spectrum is a function of wavelength containing just the two temperatures as unknown parameters. Analysis of this ratio spectrum yields the two temperatures.
In the development of ceramics these nongray emissivity materials are often heated by exposing them to a constant quartz-lamp radiation source. The working principle of RMP allows the ceramic temperatures to be determined by recording the radiation spectra from the ceramics sequentially while the sample temperature is increasing. The ratio of sums and differences of the recorded spectra eliminates the unknown reflected radiation and the unknown emissivity to form a set of new data. Again, analysis of these data as a function of wavelength containing two unknown parameters yields the two temperatures. Other spectra are similarly analyzed to determine their temperatures. Because only the ratios are analyzed, the new pyrometer has the added advantage of not requiring precalibration.
NASA Lewis has demonstrated RMP under simulated conditions. Blackbody radiation transmitted through an optical medium simulated emitted radiation, while intense quartz radiation specularly reflected from the optical medium simulated the reflected radiation. The blackbody radiation source was heated in steps to allow the slow-working spectrometer enough time to completely record a spectrum. A fast spectrometer is being designed to make a portable version of this instrument for use in projects related to the Advanced High-Temperature Engine Materials Technology Program and the Enabling Materials Program.
Lewis contact: Daniel L.P. Ng, (216) 433-3638
Headquarters program office: OA
Each specimen was instrumented with two PdCr wire strain gages applied by flame spraying, one gage on each side. A commercially available air-cooled, resistance-based extensometer with a 12.7-mm gage section was mounted on the thick edge of the specimen; the strain gages were mounted on the face of the specimen. A radiant quartz lamp heater was used for specimen heating, and type K thermocouples were used to monitor and control temperature.
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In general, the PdCr wire strain gage was found to be a useful strain-monitoring device in a thermomechanical loading system equipped with quartz lamp heating. Its apparent strain response was repeatable on both the IN100 and SCS-6/Ti-15-3 specimens. The thermal expansion strain data obtained from the two measurement systems showed excellent agreement to 800 °ree;C. Quasi-static mechanical strain responses of the two systems on both IN100 (to 800 °ree;C) and TMC (to 600 °ree;C) were also in good agreement (within 10%) to 2000 µe. At room temperature strain measurements were accurate to approximately 4500 µe.
Bibliography
Castelli, M.G; and Lei, J.-F.: A Comparison Between High Temperature Extensometry and PdCr Based Resistance Strain Gages With Multiple Application Techniques. HITEMP Review 1994: Advanced High-Temperature Engine Materials Technology Program, NASA CP-10146, Vol. II, 1994, p. 36-1.
Lewis contacts: Dr. Jih-Fen Lei, (216) 433-3922; Michael G. Castelli, (216) 433-8464;
Dr. W. Dan Williams, (216) 433-3725; Rod Ellis, (216) 433-3340
Headquarters program office: OA
Designing high-performance, vertical-lift aircraft requires knowledge of the temperature and velocity of the lift-generated flow field in all operating modes, especially near the engine inlets. During hover or vertical landing, the lift jets impinge on the ground plane and recirculate near the aircraft. Under some conditions the hot exhaust gases can be ingested by the engine inlets, causing loss of efficiency or catastrophic compressor stall. NASA Lewis and McDonnell Douglas conducted an extensive program in Lewis' 9- by 15-Foot Low-Speed Wind Tunnel to study the problem of hot gas ingestion for new ASTOVL (advanced short-takeoff, vertical-landing) aircraft designs.
A nonintrusive optical system based on Rayleigh scattering from ambient air molecules was used to measure the gas temperature and the wingspan component of the velocity in the tunnel. A laser beam was focused in the region under the model in the lift-generated flow field. Light scattered from gas molecules was collected and analyzed by using a Fabry-Perot interferometer, which measured the frequency spectrum of the scattered light. It is this spectrum that contains the information about gas temperature and velocity.
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In a significant breakthrough this technique was demonstrated in the harsh environment of a wind tunnel facility, with "dirty" gases, high vibration levels, and high temperatures. Because Rayleigh scattering is from molecules, no seed particles have to be added to the flow to obtain a scattering signal. In fact, noise caused by scattering from unwanted particles in the flow has been an obstacle to applying Rayleigh techniques outside the laboratory environment. This test demonstrated that the spectrally resolved technique used can extract sufficient Rayleigh signal (i.e., temperature and velocity) in the presence of particle noise to make valid measurements in the flow field. Also, laser beam delivery and signal capture was achieved by using fiber optic cables to isolate the laser and interferometer systems from the noisy wind tunnel environment.
The system took data at over 900 measurement locations in a volume under the model while the tunnel and model simulated ground-effect flow (i.e., hover or landing). Temperature measurements were obtained with a 5% mean uncertainty, and single-component velocity measurements (along the wingspan direction) with 10-m/sec uncertainty.
Lewis contacts: Helen E. Kourous, (216) 433-6569; Dr. Richard G. Seasholtz, (216) 433-3754
Headquarters program office: OA
Liquid-Crystal, Point-Diffraction Interferometer Invented
A new instrument, the liquid-crystal, point-diffraction interferometer (LCPDI), was invented at NASA Lewis to measure optical wavefronts in transparent media, such as glasses, gases, and liquids. Typical applications are in verifying the optical performance of lenses and optical
elements and in measuring temperature, density, or concentration distributions in air or water during wind tunnel or microgravity experiments. The LCPDI represents an improvement over other techniques because it has both a common-path design and the ability to make digital measurements, providing automated data reduction in a robust package.
The LCPDI consists of a thin liquid-crystal layer sandwiched between two glass plates. Transparent electrodes are deposited on the inner surface of each plate. The refractive index of the liquid-crystal layer can be modified by altering the strength of an electric field applied across the electrodes. A transparent microsphere is embedded in the liquid-crystal layer near the center of a clear aperture. The refractive index of the microsphere is unaffected by the electric field.
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A collimated laser beam is passed through the medium under study and then focused onto the microsphere in the LCPDI. Because the focused spot is larger than the microsphere, some of the light passes through the liquid crystals. This portion of the transmitted beam is called the object beam because it contains information about the transparent medium through which it passed. The rest of the beam passes through the microsphere. This light is diffracted into a spherical reference beam. The object and reference beams travel together until they illuminate a viewing screen, where they combine to form an interferogram. A solid-state camera records the interferogram. Phase-stepping interferometry is used to extract the two-dimensional phase of the object beam from the interferogram. A relative phase shift is introduced between the object and reference beams by changing the strength of the electric field applied across the liquid crystals. A sequence of interferograms is recorded with the phase stepped by nominally 90°ree; between frames. The recorded images are then digitally recombined to extract the phase front. This phase front is analyzed to determine the shape of the lens or the properties of the flow under study.
This work was performed in-house. Two papers have so far been generated from this work, a patent application has been filed, and a joint proposal with a microgravity group has been submitted. The program is ongoing.
Bibliography
Mercer, C.R.; and Creath, K.: Liquid Crystal Point Diffraction Interferometer. Optics Letters, vol. 19, no. 12, 1994, pp. 916-918.
Lewis contact: Carolyn R. Mercer, (216) 433-3411
Headquarters program office: OA
Performance-Limiting Micropipe Defects Identified in SiC Wafers
The in-house High-Temperature Integrated Electronics and Sensors (HTIES) Program at NASA Lewis is currently developing a family of silicon carbide (SiC) semiconductor devices for use in high-temperature, high-power, and/or high-radiation conditions. Conventional semiconductors, such as silicon and gallium arsenide (GaAs), cannot adequately perform under these conditions.
Integrated electronics and sensors capable of operating in a hostile environment would have numerous important aerospace-related applications, such as propulsion control, power control, radar and communications, and radiation-hardened circuits. They also would find numerous spinoff
applications in the Earth-bound commercial power and automobile industries.
Theoretical appraisals have indicated that SiC power devices would operate over higher voltage and temperature ranges, have superior switching characteristics, and yet have die sizes nearly 20 times smaller than correspondingly rated silicon-based devices. However, these tremendous theoretical advantages have yet to be realized in experimental SiC devices. Prototype SiC devices constructed to date have failed to operate at currents of more than 2 A. Investigations by the NASA Lewis HTIES research group recently identified the crystal defects restricting SiC power devices to these relatively small electrical current levels.
All useful SiC devices are fabricated starting from commercially available silicon carbide wafers. These wafers contain crystal defects called micropipes, small tubular voids that run through the wafers in a direction normal to the polished wafer surface. The NASA Lewis HTIES team showed that micropipe defects originating in SiC wafers are responsible for premature electrical failures in most SiC devices larger than 1 mm&super;2 in area. In particular, the micropipes were experimentally linked to localized reverse failures at reverse-bias voltages well below the known breakdown voltage inherent to homogeneous SiC. This was found to be the case even when the device is fabricated entirely within high-quality SiC epitaxial layers that have been grown on top of the original SiC wafer, due to the observed fact that the defects propagate from the wafer into the SiC epitaxial layers as they are grown. The accompanying figure shows the visible microplasma that is observed within a micropipe when an SiC diode junction fails prematurely.
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The establishment of micropipes in SiC wafers as a defect limiting the performance of high-voltage SiC devices has serious implications for the near-term realization of high-current SiC power devices. Until micropipe density is significantly reduced, SiC power device ratings will be restricted because the defects prevent scaleup to the large device areas (>1 mm2) needed to carry high current (>10 A). Higher than existing voltage and current ratings will not be attained until improvements in SiC crystal growth enable larger defect-free areas.
The announcement of these findings by the NASA Lewis HTIES team has triggered intensified research worldwide toward eradicating micropipe defects from SiC wafers. A twofold reduction in micropipes was recently announced by Cree Research, Inc.
Lewis contacts: Dr. Philip G. Neudeck, (216) 433-8902; J. Anthony Powell, (216) 433-3652
Headquarters program office: OA
Fiber Optic Control System Integration Program Concluded
The Fiber Optic Control System Integration (FOCSI) Program was a NASA/NAVY program designed to demonstrate that passive optical sensor systems can operate in the severe environment of advanced aircraft. As part of the program an attempt was made to design a standard electro-optic interface that would accommodate analog and digital sensors using various optical modulation techniques.
This program is a foundation for future fly-by-light (FBL) programs. An FBL system utilizes passive optical sensors and fiber optic data links to monitor and control aircraft systems. This technology has many benefits for commercial and military aircraft. It improves reliability and reduces certification costs because FBL aircraft are more immune to electromagnetic effects than fly-by-wire (FBW) aircraft. It reduces waveguide harness and run-length weight and volume. It eliminates short circuits and sparking due to insulation degradation. Other anticipated benefits are lower maintenance costs, absence of signal ground loops, and wider flexibility in control architecture and data bus design.
In the FOCSI program a number of passive optical sensors for the primary flight and propulsion control system were designed to fit alongside electrical sensors in an F-18 aircraft and F404 engine. Flight tests conducted at NASA Dryden indicated that most of the optical sensors tracked the F-18's electrical sensors very well. The flight tests concluded a successful program. The sensors will continue to operate during other missions to obtain some long-term reliability information. In addition to the flight tests a number of useful lessons learned in installation, maintenance, and installed troubleshooting are reported in the final reports of the FOCSI program. The FOCSI program will provide useful information to the military and commercial FBL programs and will help to transition the technology into production aircraft and engine control systems.
Lewis contact: Robert J. Baumbick, (216) 433-3735
Headquarters program office: OA
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Along with small-scale detector and amplification circuits, discrete medium-power diodes and field-effect transistors (FET's) are expected to be among the first SiC electronic components incorporated into these systems. Using its recently developed site-competition epitaxy technique, the NASA Lewis HTIES group has fabricated and characterized several batches of buried-gate junction field-effect transistors (JFET's) that exhibit very promising characteristics. Despite the severe device area limitation (~1 mm&super;2) imposed by present-day SiC wafer defect densities, SiC JFET's with peak drain currents above 1 A at room temperature (top graph) and standoff voltages of 100 V (only 40 V of which are shown) have been realized. The 1.2-A peak drain current exhibited in the graph represents a record for any SiC FET rated at 100 V or more with a less-than-20-V switching gate voltage spread. Further substantial increases in device performance can be expected as SiC wafer defect densities are decreased.
JFET's optimized for high-temperature operation were also fabricated as part of this work. These JFET's operated for a record 30 hr in the 600 °ree;C air environment with only a small degradation in device characteristics due to oxidative reactions of the metals used to make electrical contact with the SiC. High-temperature device lifetime should increase when process refinements designed to prevent contact metal oxidation are implemented.
Lewis contacts: Dr. Philip G. Neudeck, (216) 433-8902; Dr. David J. Larkin, (216) 433-8718
Headquarters program office: OA
Site-Competition Epitaxy Controls Doping of Silicon Carbide
Silicon carbide (SiC) is emerging as the material of choice for high-power and/or microwave-frequency semiconductor devices suitable for high-temperature, high-radiation, and corrosive environments. Currently, the initial step in making SiC semiconductor devices is a process called chemical vapor deposition (CVD), which allows single-crystal layers (epilayers) of varying electrical character to be grown. The SiC epilayers are produced in the CVD process by thermally decomposing commercially available
silicon and carbon source gases onto boule-derived SiC substrates. The electrical character is tailored by adding dopants, impurity elements that affect the epilayer electrical properties, during the CVD SiC epilayer growth by flowing either nitrogen (for n type) or trimethylaluminum (for p type). However, for the inherently superior high-temperature semiconductor properties of SiC to be realized in advanced electronic devices, control over the electronic properties of the CVD SiC epilayers must be improved.
Control over dopant incorporation for CVD SiC epilayers had been very limited. Reproducible doping was typically confined to a relatively narrow doping range, restricting SiC device performance to below theoretically predicted values. This narrow doping range was recently greatly expanded by the NASA Lewis discovery that the ratio of silicon source flow to carbon source flow during the CVD epilayer growth could be used to control dopant incorporation and therefore the electrical properties of the growing SiC epilayers. This process, named site-competition epitaxy, produces much lower doped SiC epilayers than was previously possible. In addition, site-competition epitaxy can produce more highly doped layers when more electrically conductive SiC epilayers are desired. Expanding the reproducible doping range to include lower concentrations has enabled the fabrication of multikilovolt SiC power devices, whereas the availability of higher doping concentrations has resulted in devices with improved performance because of lower parasitic resistances.
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The site-competition epitaxy working model is based on the competition between the SiC and dopant source gases for the available substitutional lattice sites on the growing SiC crystal surface. Dopant incorporation is controlled by appropriately adjusting the Si/C ratio within the growth reactor to affect the amount of dopant atoms incorporated into these sites, either carbon-lattice sites (C sites) or silicon-lattice sites (Si sites), located on the active growth surface of the SiC crystal. Specifically, our model for site-competition epitaxy is based on the principle of competition between nitrogen and carbon for the C sites and between aluminum and silicon for the Si sites of the growing SiC epilayer. The concentration of n-type (nitrogen) dopant atoms, which can occupy only C sites, incorporated into a growing SiC epilayer can be decreased by increasing the carbon source concentration so that C "out competes" N for the available C sites.
In summary, the nitrogen donor concentration in the grown epilayer is proportional to the Si/C ratio during epilayer growth, whereas the aluminum acceptor concentration is inversely proportional to the Si/C ratio. Site-competition epitaxy was used (as depicted) to produce the very low-doped epilayers required for the world's first 6H-SiC 2000-V and 3C-SiC 300-V diodes. More recently, this novel growth technique has enabled the fabrication of SiC junction field-effect transistors (JFET's) that can operate at 600 °ree;C in air for 30 hr. The site-competition epitaxy technique not only greatly expands the doping range for SiC but also allows for more reproducible dopant control within the previously attainable doping range. Additionally, site-competition epitaxy is the only known route for growing degenerately doped epilayers that result in ohmic as-deposited (i.e., unannealed) contacts for a variety of metals on both p-type and n-type SiC epilayers.
This work was performed in-house as part of the High-Temperature Integrated Electronics and Sensors (HTIES) Program at NASA Lewis.
Lewis contacts: Dr. David J. Larkin, (216) 433-8718; Dr. Philip G. Neudeck, (216) 433-8902; J. Anthony Powell, (216) 433-3652; Dr. Lawrence G. Matus, (216) 433-3650
Headquarters program office: OA
Parallel Implementation of Real-Time Expert Diagnostic System Achieved
Intelligent rocket engine control systems are desired that significantly reduce operating costs, increase flight rates, improve performance, and extend engine life. Achieving these goals requires engine controllers that have integrated capabilities to diagnose anomalous operation, that implement accommodating closed-loop control to extend engine life, and that maximize available system performance (ref. 1). A reusable rocket engine diagnostic system (ReREDS) has been developed for use in intelligent control systems of the space shuttle main engines (SSME's) and other reusable rocket engines (refs. 2 and 3). However, a practical implementation of this expert system was needed that would allow failure diagnosis to be carried out in real time. NASA Lewis has developed and demonstrated a methodology for the parallel, real-time implementation of expert systems using the ReREDS expert diagnostic system.
A sequential, or serial, implementation of ReREDS using the CLIPS expert system development tool on an off-the-shelf microcomputer was found to be too slow to be practical. However, as implemented serially, ReREDS already exhibited a certain degree of inherent parallelism. Ten sensor measurements are used to diagnose the presence of five different failures that may manifest themselves in the working SSME. Each module of code that diagnoses one failure is called a failure detector. Although some sensor measurements are shared between failure detectors, the computations within these detectors are completely independent of one other. Thus, the failure detection problem was partitioned so that each failure detector was assigned to its own processor. Because the failure detectors can be processed simultaneously, a speedup in the execution is expected.
This partitioned architecture for the expert diagnostic system and the CLIPS expert system development tool were ported to the NASA Lewis iPSC/860 hypercube parallel computer. Each failure detector resides on a node of the hypercube. The correct operation of each individual failure detector was verified by using simulated SSME sensor data. An ANSI C program was written initializing the copy of CLIPS on each node, so that CLIPS was used as an embedded application on each failure detector node to evaluate the ReREDS failure detection rules. All other program requirements, including opening and closing of sensor measurement data files, preliminary data analysis, and program flow control, are handled in C language. In addition, a server node, programmed entirely in C language, transfers incoming sensor data into an indexed memory array, or blackboard, from which data are distributed on request to the failure detector nodes. A manager node coordinates timing between failure detector nodes and the server node. Using a C language executive on each node and embedded CLIPS language reasoning on each failure detector node provides an efficient structure for realizing context focusing, process interruption, and overall expert system communications and synchronization.
Profiling studies found that the parallel ReREDS implementation could process the sensor measurements and report confidence levels for all five failure modes in 18 msec, nearly three times faster than the serial implementation on the same computer, and much faster than the 100 msec normally required. Additional speedup could be achieved by parallelizing the data source because the failure nodes tended to spend considerable time waiting while contending for data from the server node. However, in general the method used provides a fast, predictable, continuously operating failure detection system. In addition, this method should be useful for implementing other expert systems requiring real-time operation.
This work was carried out through a cooperative agreement between NASA Lewis and Cleveland State University.
References
Lewis contact: John C. DeLaat, (216) 433-3744
Headquarters program office: OA
Closed-Loop Control Demonstrated for Rocket Engine Faults
The goal of the Intelligent Control System Program is to extend the
useful life of a reusable rocket propulsion system while minimizing flight maintenance and maximizing performance through improved control and monitoring algorithms. During the past years the program has successfully demonstrated real-time fault accommodation of two classes of faults: sensors and actuators. The closed-loop simulation of the engine control during the fault and after the accommodation provides important information for the actual implementation of the algorithms. For sensor faults an autoassociative neural network is used to detect an abnormal sensor reading and also to estimate the actual value. For actuator faults a model-based
fault detection system with a reconfigurable control is used in the closed-loop environment.
In sensor failure accommodation a group of analytically redundant measurements, including all the variables used in the closed-loop control for engine performance, is first identified. The data generated from the engine model as well as from the test-bed engine firing are used to train the autoassociative neural network, which operates on the principle of dimensionality reduction. The network is trained to produce an output vector equal to its original input vector. The redundant sensor information is compressed, mixed, and regenerated to provide an estimate of the true measurement. Sensor failure is accommodated by replacing the failed sensor with the estimate for the controller. Simulation results show that the proposed sensor validation scheme can adequately identify the failed sensors and provide reasonable estimates for control purposes.
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Model-based fault detection is used to detect actuator failures. A nominal engine model detects any abnormal operating behavior that can be classified into a specified structure. A fault model with a vector of fault parameters associated with all actuator faults is first developed. Different fault parameters represent different types of faults of a specific actuator. A recursive, least-squares error algorithm performs the real-time estimation of the fault parameters from the most current input/output data and the system matrices, which have been determined a priori. An actuator fault is accommodated on two levels. On the vehicle level the information about the actuator fault is used to determine the thrust command to the engine in order to maintain the mission goal. On the engine level a reconfigurable control is used to execute the engine command with a reduced capability in actuation.
References
Piloted Evaluation Performed for Integrated Propulsion and Airframe Control Design Methodology
The trend is toward military fighter and tactical aircraft with new or enhanced maneuver capabilities, such as short takeoff and vertical landing (STOVL) and high-angle-of-attack performance. An integrated flight and propulsion control system is required to obtain these enhanced capabilities while maintaining reasonable pilot workload. At NASA Lewis an integrated control design methodology called IMPAC, Integrated Methodology for Propulsion and Airframe Control, has been developed and then applied to the development of a STOVL aircraft control design (ref. 1). A piloted evaluation of this control design has been performed in-house at NASA Lewis on a fixed-base flight simulator (ref. 2).
The NASA Lewis flight simulation facility consists of a cockpit equipped with five pilot effectors, a control computer where the integrated control algorithm resides, a simulation computer that houses the integrated propulsion and airframe model, a Unix workstation for graphics development, an image generation system, and three overhead projectors that display the out-the-window scenery and heads-up display symbology on a wraparound forward projection screen. For this program the cockpit and displays were configured to simulate the operation of a STOVL aircraft.
A piecewise linear airframe model was created to simulate the dynamic operation of an ejector-configured conceptual STOVL aircraft. This airframe model was then coupled with a multinozzle turbofan engine model. For this evaluation the integrated engine and airframe model was limited to the flight envelope in the transition from cruise to hover. The integrated flight and propulsion control algorithm was developed using IMPAC. Initial pilot comments and suggestions were very helpful in arriving at the simulator configuration used for the final piloted evaluation.
In the final piloted evaluation two NASA test pilots were asked to fly several predefined evaluation tasks to assess controllability, performance, and workload during a series of four flight scenarios. The selected tasks required the control to undergo the longitudinal, lateral, acceleration, and deceleration commands necessary for a STOVL aircraft flying in the transition flight envelope. The pilots successfully completed the evaluation and commented favorably about the performance and workload for all the tasks.
This work has resulted in the development of a flexible, fixed-base flight simulation facility for integrated airframe and propulsion control systems research. It has also resulted in the successful demonstration of IMPAC as a highly useful control design methodology. Plans include using the simulation facility and IMPAC for other integrated flight and propulsion control programs.
References
Dynamic Wave Rotor Performance Modeled
The wave rotor is being investigated for use as a core gas generator in future gas turbine engines in order to achieve high peak cycle temperatures and pressures. The device uses gas-dynamic waves to transfer energy directly to and from the working fluid through which the waves travel. It consists of a series of constant-area passages that rotate around an axis. Through rotation the ends of the passages are periodically exposed to various circumferentially arranged ports that initiate the traveling waves within the passages. Because each passage of the wave rotor is periodically exposed to both hot and cold flow, the mean temperature of the rotor material should remain considerably below the peak cycle temperature. Preliminary
steady-state design calculations were made for a four-port wave rotor core using subsonic engine technology planned for entry into service in the year 2015. The results indicate that the wave rotor can readily meet aerothermo-dynamic requirements and maintain a mean passage wall temperature approximately 440 °ree;C below the peak cycle temperature.
Although the steady-state performance of the wave rotor appears promising, critical questions concerning dynamic performance when the port conditions are transient (i.e., fuel flow or inlet pressure perturbations) remain. To answer these questions, NASA Lewis developed a model that predicts the state of the fluid in all wave rotor passages as they are exposed to the various ports at the ends. The conditions in the ports may be either steady or transient. The passages are assumed to have uniform properties at any cross section (i.e., one-dimensional flow), and the gas is assumed to be calorically perfect. The model can assess losses induced by viscosity, by heat transfer to and from the passage walls, by the finite opening time of the passages as they enter and exit port regions, and by gas leakage between the passage ends and the stationary walls to and from the cavity in the rotor center. The combustor and cavity, having much larger time constants than the rotor passages, are modeled by using lumped-volume techniques. The wave rotor model has been tested in preliminary simulations using step changes from the design point in fuel flow and exhaust port backpressure.
Lewis contact: Daniel E. Paxson, (216) 433-8334
Headquarters program office: OA
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