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  • Aircraft Propulsion and Power
  • Earth model, also for more shallow analyses !
  • 1995-1999  (75)
  • 1990-1994
  • 1950-1954  (26)
  • 1998  (75)
  • 1954  (5)
  • 1951  (21)
Collection
Keywords
Publisher
Years
  • 1995-1999  (75)
  • 1990-1994
  • 1950-1954  (26)
Year
  • 1
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    Crustal structure above the Iceland mantle plume imaged by the ICEMELT refraction profile (1998)
    In:  Geophys. J. Int., Warszawa, Pergamon, vol. 135, no. 3, pp. 1131-1149, pp. B05310, (ISBN: 0-12-018847-3)
    Details
    Publication Date: 1998
    Keywords: Velocity depth profile ; Refraction seismics ; Earth model, also for more shallow analyses ! ; Flovenz ; GJI
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    BIBLIOGRAPHY SEISMOLOGY
  • 2
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    50 Jahre Helgolandsprengung (1998)
    In:  DGG-Mittl., Münster, 3, vol. 75, no. 2, pp. 2-32, pp. L12311, (ISBN 0-471-26610-8)
    Details
    Publication Date: 1998
    Keywords: Deep seismic sounding (espec. cont. crust) ; Refraction seismics ; Reflection seismics ; Acoustics ; Travel time ; Earth model, also for more shallow analyses ! ; Velocity depth profile
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    BIBLIOGRAPHY SEISMOLOGY
  • 3
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    Three-dimensional seismic structure beneath Shillong Plateau and Assam Valley, northeast India (1998)
    In:  Bull. Seism. Soc. of America, Amsterdam, Univ. Tokyo, vol. 88, no. 3, pp. 667-676, pp. L14309, (ISSN: 1340-4202)
    Details
    Publication Date: 1998
    Keywords: Three dimensional ; Earth model, also for more shallow analyses ! ; Velocity depth profile ; BSSA
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    BIBLIOGRAPHY SEISMOLOGY
  • 4
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    Compressional and shear velocity structure of the lithosphere in Northern Iceland (1998)
    In:  Bulletin of the Seismological Society of America, Hannover, Conseil de l'Europe, vol. 88, no. 6, pp. 1561-1571, pp. L24314, (ISSN: 1340-4202)
    Details
    Publication Date: 1998
    Keywords: Velocity depth profile ; Earth model, also for more shallow analyses ! ; Seismology ; BSSA
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    BIBLIOGRAPHY SEISMOLOGY
  • 5
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    CRUST 5.1: A global crustal model at 5° x 5° (1998)
    In:  J. Geophys. Res., Edmonton, Conseil de l'Europe, vol. 103, no. B1, pp. 727-747, pp. B10410, (ISSN: 1340-4202)
    Details
    Publication Date: 1998
    Keywords: Earth model, also for more shallow analyses ! ; Seismics (controlled source seismology) ; JGR
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    BIBLIOGRAPHY SEISMOLOGY
  • 6
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    Three-dimensional seismic structure and moment tensors of non-double-couple earthquakes at the Hengill-Grensdalur volcanic complex, Iceland (1998)
    In:  Geophys. J. Int., Luxembourg, Conseil de l'Europe, vol. 133, no. 2, pp. 309-325, pp. L24308, (ISSN: 1340-4202)
    Details
    Publication Date: 1998
    Keywords: Moment tensor ; Earth model, also for more shallow analyses ! ; Three dimensional ; Volcanology ; Seismology ; Geothermics ; GJI
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    BIBLIOGRAPHY SEISMOLOGY
  • 7
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    On the density distribution within the Earth (1998)
    In:  Geophys. J. Int., New York, Allerton Press, vol. 132, no. 2, pp. 374-382, pp. L01606, (ISSN: 1340-4202)
    Details
    Publication Date: 1998
    Keywords: Earth model, also for more shallow analyses ! ; Gravimetry, Gravitation ; GJI
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    BIBLIOGRAPHY SEISMOLOGY
  • 8
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    New perspectives on mantle dynamics from high-resolution seismic tomographic model P1200 (1998)
    In:  Pure and Applied Geophysics, Potsdam, ZIPE, vol. 151, no. 2-4, pp. 503-525, pp. 2156, (ISBN: 0-12-018847-3)
    Details
    Publication Date: 1998
    Keywords: Tomography ; GeodesyY ; ConvolutionE ; Teleseismic events ; Seismology ; Earth model, also for more shallow analyses ! ; Cadek ; Cizkova ; PAG
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    BIBLIOGRAPHY SEISMOLOGY
  • 9
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    Thermal-mechanical model for crustal thickening in the central Andes driven by ablative subduction (1998)
    In:  Geology, Stockholm, Wissenschaftliche Buchgesellschaft, vol. 26, no. 6, pp. 511-514, pp. L15318, (ISSN: 1340-4202)
    Details
    Publication Date: 1998
    Description: The thick (up to 65 km) crust of the high-elevation regions of the Andean orogenic belt results from structural thickening of the felsic upper crust of South America. The great thickness implies shortening or loss of several hundred kilometers of lithospheric mantle and lower crust. Through numerical simulation of the thermal and mechanical processes of crustal thickening, we test a model of lithospheric-mantle removal by ablative subduction. The model combines kinematic assumptions regarding mantle motion with dynamic calculations of crustal deformation. Predictions include the formation of an outwardly growing orogenic plateau - analogous to the Altiplano - that results from high-temperature, low-viscosity flow of the lower crust due to thickening of the high-heat-production upper crust.
    Keywords: Plate tectonics ; Earth model, also for more shallow analyses ! ; Geothermics ; Subduction zone ; Geol. aspects ; Rheology
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    BIBLIOGRAPHY SEISMOLOGY
  • 10
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    Global teleseismic earthquake relocation with improved travel times and procedures for depth determination (1998)
    In:  Bull. Seism. Soc. of America, Münster, 3, vol. 88, no. 3, pp. 722-743, pp. L12311, (ISBN 0-471-26610-8)
    Details
    Publication Date: 1998
    Keywords: Teleseismic events ; Location ; Travel time ; Hypocentral depth ; Earth model, also for more shallow analyses ! ; BSSA ; Hilst
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    BIBLIOGRAPHY SEISMOLOGY
  • 11
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    EurID: a regionalized 3-D seismological model of Europe (1998)
    In:  Phys. Earth Plan. Int., Luxembourg, Pergamon, vol. 106, no. 1-2, pp. 31-62, pp. 2131, (ISBN 1-86239-117-3)
    Details
    Publication Date: 1998
    Keywords: Three dimensional ; Seismology ; Modelling ; Surface waves ; Velocity depth profile ; Earth model, also for more shallow analyses ! ; PEPI
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    BIBLIOGRAPHY SEISMOLOGY
  • 12
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    The Applicability of Jet-Shear-Layer Mixing and Effervescent Atomization for Low-NO(x) Combustors (1998)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: An investigation has been conducted to develop appropriate technologies for a low-NO(x), liquid-fueled combustor. The combustor incorporates an effervescent atomizer used to inject fuel into a premixing duct. Only a fraction of the combustion air is used in the premixing process. This fuel-rich mixture is introduced into the remaining combustion air by a rapid jet-shear-layer mixing process involving radial fuel-air jets impinging on axial air jets in the primary combustion zone. Computational modeling was used as a tool to facilitate a parametric analysis appropriate to the design of an optimum low-NO(x) combustor. A number of combustor configurations were studied to assess the key combustor technologies and to validate the three-dimensional modeling code. The results from the experimental testing and computational analysis indicate a low-NO(x) potential for the jet-shear-layer combustor. Key features found to affect NOx emissions are the primary combustion zone fuel-air ratio, the number of axial and radial jets, the aspect ratio and radial location of the axial air jets, and the radial jet inlet hole diameter. Each of these key parameters exhibits a low-NO(x) point from which an optimized combustor was developed Also demonstrated was the feasibility of utilizing an effervescent atomizer for combustor application. Further developments in the jet-shear-layer mixing scheme and effervescent atomizer design promise even lower NO(x) with high combustion efficiency.
    Keywords: Aircraft Propulsion and Power
    Type: Journal of Engineering for Gas Turbines and Power; Volume 120; 17-23
    Format: text
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    NASA TECHNICAL REPORTS
  • 13
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    Calculation of Tip Clearance Effects in a Transonic Compressor Rotor (1998)
    In:  CASI
    Details
    Publication Date: 2013-08-29
    Description: The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concern- ing the use of the simple clearance model Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computed results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.
    Keywords: Aircraft Propulsion and Power
    Type: Journal of Turbomachinery; Volume 120; 131-140
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    NASA TECHNICAL REPORTS
  • 14
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    Simulated Bladed MMC Disk LCF Validation (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The goal of this program was to evaluate the low cycle fatigue behavior of an SCS-6/Ti-6Al-4V sub-component under bi-axial loading conditions at 316 C(600 F). A simulated bladed TMC disk was designed having thirty four blades representing the number that would be used in Allied Signal's JTAGG II impeller. The outer diameter of the bladed ring was 254 mm (10.0 inch) and the inner diameter 114.3 mm (4.50 inch). The outer and inner diameter of the composite zone was 177.8 mm (7.00 inch) and 127.O mm(5.00 inch) respectively. Stress analysis showed that the fatigue life of the bladed composite ring would be about 12000 cycles for the test conditions applied. A modal analysis was conducted which showed that the blades would have sufficient life margin from dynamic excitation. The arbor design was the same as that employed in the spin-to burst test of NAS3-27027. A systematic stress analysis of each part making up the arbor was undertaken to assure the design would meet the low cycle fatigue requirements of the program. The Textron Systems grooved foil-fiber process was chosen to make the SCS-6/Ti-6Al-4V core ring based on the success they had in contract NAS3-27027. Fiber buckling, however, was observed at several locations in the first ring made which rendered it unsuitable for spin testing. The fiber buckling was attributed to cracking of the graphite tooling during the consolidation process. On this basis a second ring was made but it too contained fiber buckling defects. Analysis by Textron indicated that the fiber buckling was most likely due to poor placement of the SCS-6 fiber in the etched grooves of the Ti-6Al-4V foil. This was also a contributor to the defects in the first ring. Since there was little indication of control in the process to manufacture a quality ring a third attempt at making a ring was not undertaken.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-207063 , NAS 1.26:207063
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    NASA TECHNICAL REPORTS
  • 15
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    Study of a Wake Recovery Mechanism in a High-Speed Axial Compressor Stage (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This work addresses the significant differences in compressor rotor wake mixing loss which exist in a stage environment relative to a rotor in isolation. The wake decay for a rotor in isolation is due solely to viscous dissipation which is an irreversible process and thus leads to a loss in both total pressure and efficiency. Rotor wake decay in the stage environment is due to both viscous mixing and the inviscid strain imposed on the wake fluid particles by the stator velocity field. This straining process, referred to by Smith (1993) as recovery, is reversible and for a 2D rotor wake leads to an inviscid reduction of the velocity deficit of the wake. A model for the rotor wake decay process is developed and used to quantify the viscous dissipation effects relative to those of inviscid wake stretching. The model is verified using laser anemometer measurements acquired in the wake of a transonic rotor operated in isolation and in a stage configuration at near peak efficiency and near stall operating conditions. Additional insight is provided by a time-accurate 3D Navier-Stokes simulation of the compressor stator flow field at the corresponding stage loading levels. Results from the wake decay model exhibit good agreement with the experimental data. Data from the model, laser anemometer measurements, and numerical simulations indicate that for the rotor/stator spacing used in this work, which is typical of core compressors, rotor wake straining (stretching) is the primary decay process in the stator passage with viscous mixing playing only a minor role. The implications of these results on compressor stage design are discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-206594 , E-11045 , NAS 1.26:206594
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    NASA TECHNICAL REPORTS
  • 16
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    Numerical Investigation of Hot Gas Ingestion by STOVL Aircraft (1998)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This report compiles the various research activities conducted under the auspices of the NASA Grant NAG3-1026, "Numerical Investigation of Hot Gas Ingestion by STOVL Aircraft" during the period of April 1989 to April 1994. The effort involved the development of multigrid based algorithms and computer programs for the calculation of the flow and temperature fields generated by Short Take-off and Vertical Landing (STOVL) aircraft, while hovering in ground proximity. Of particular importance has been the interaction of the exhaust jets with the head wind which gives rise to the hot gas ingestion process. The objective of new STOVL designs to reduce the temperature of the gases ingested into the engine. The present work describes a solution algorithm for the multi-dimensional elliptic partial-differential equations governing fluid flow and heat transfer in general curvilinear coordinates. The solution algorithm is based on the multigrid technique which obtains rapid convergence of the iterative numerical procedure for the discrete equations. Initial efforts were concerned with the solution of the Cartesian form of the equations. This algorithm was applied to a simulated STOVL configuration in rectangular coordinates. In the next phase of the work, a computer code for general curvilinear coordinates was constructed. This was applied to model STOVL geometries on curvilinear grids. The code was also validated in model problems. In all these efforts, the standard k-Epsilon model was used.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-4769 , NAS 1.26:4769 , E-10676
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    NASA TECHNICAL REPORTS
  • 17
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    Ducted-Fan Engine Acoustic Predictions using a Navier-Stokes Code (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: A Navier-Stokes computer code is used to predict one of the ducted-fan engine acoustic modes that results from rotor-wake/stator-blade interaction. A patched sliding-zone interface is employed to pass information between the moving rotor row and the stationary stator row. The code produces averaged aerodynamic results downstream of the rotor that agree well with a widely used average-passage code. The acoustic mode of interest is generated successfully by the code and is propagated well upstream of the rotor; temporal and spatial numerical resolution are fine enough such that attenuation of the signal is small. Two acoustic codes are used to find the far-field noise. Near-field propagation is computed by using Eversman's wave envelope code, which is based on a finite-element model. Propagation to the far field is accomplished by using the Kirchhoff formula for moving surfaces with the results of the wave envelope code as input data. Comparison of measured and computed far-field noise levels show fair agreement in the range of directivity angles where the peak radiation lobes from the inlet are observed. Although only a single acoustic mode is targeted in this study, the main conclusion is a proof-of-concept: Navier-Stokes codes can be used both to generate and propagate rotor/stator acoustic modes forward through an engine, where the results can be coupled to other far-field noise prediction codes.
    Keywords: Aircraft Propulsion and Power
    Type: Journal of Sound and Vibration; Volume 123; No. 4; 643-664
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    NASA TECHNICAL REPORTS
  • 18
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    Effect of Installation of Mixer/Ejector Nozzles on the Core Flow Exhaust of High-Bypass-Ratio Turbofan Engines (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: The aerospace industry is currently investigating the effect of installing mixer/ejector nozzles on the core flow exhaust of high-bypass-ratio turbofan engines. This effort includes both full-scale engine tests at sea level conditions and subscale tests in static test facilities. Subscale model tests are to be conducted prior to full-scale testing. With this approach, model results can be analyzed and compared with analytical predications. Problem areas can then be identified and design changes made and verified in subscale prior to committing to any final design configurations for engine ground tests. One of the subscale model test programs for the integrated mixer/ejector development was a joint test conducted by the NASA Lewis Research Center and Pratt & Whitney Aircraft. This test was conducted to study various mixer/ejector nozzle configurations installed on the core flow exhaust of advanced, high-bypass-ratio turbofan engines for subsonic, commercial applications. The mixer/ejector concept involves the introduction of largescale, low-loss, streamwise vortices that entrain large amounts of secondary air and rapidly mix it with the primary stream. This results in increased ejector pumping relative to conventional ejectors and in more complete mixing within the ejector shroud. The latter improves thrust performance through the efficient energy exchange between the primary and secondary streams. This experimental program was completed in April 1997 in Lewis' CE-22 static test facility. Variables tested included the nozzle area ratio (A9/A8), which ranged from 1.6 to 3.0. This ratio was varied by increasing or decreasing the nozzle throat area, A8. Primary nozzles tested included both lobed mixers and conical primaries. These configurations were tested with and without an outer shroud, and the shroud position was varied by inserting spacers in it. In addition, data were acquired with and without secondary flow.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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    NASA TECHNICAL REPORTS
  • 19
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    F100 Engine Emissions Tested in NASA Lewis' Propulsion Systems Laboratory (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-02
    Description: Recent advances in atmospheric sciences have shown that the chemical composition of the entire atmosphere of the planet (gases and airborne particles) has been changed due to human activity and that these changes have changed the heat balance of the planet. National Research Council findings indicate that anthropogenic aerosols1 reduce the amount of solar radiation reaching the Earth's surface. Atmospheric global models suggest that sulfate aerosols change the energy balance of the Northern Hemisphere as much as anthropogenic greenhouse gases have. In response to these findings, NASA initiated the Atmospheric Effects of Aviation Project (AEAP) to advance the research needed to define present and future aircraft emissions and their effects on the Earth's atmosphere. Although the importance of aerosols and their precursors is now well recognized, the characterization of current subsonic engines for these emissions is far from complete. Furthermore, since the relationship of engine operating parameters to aerosol emissions is not known, extrapolation to untested and unbuilt engines necessarily remains highly uncertain. Tests in 1997-an engine test at the NASA Lewis Research Center and the corresponding flight measurement test at the NASA Langley Research Center-attempted to address both issues by measuring emissions when fuels containing different levels of sulfur were burned. Measurement systems from four research groups were involved in the Lewis engine test: A Lewis gas analyzer suite to measure the concentration of gaseous species 1. including NO, NOx, CO, CO2, O2, THC, and SO2 as well as the smoke number; 2. A University of Missouri-Rolla Mobile Aerosol Sampling System to measure aerosol and particulate properties including the total concentration, size distribution, volatility, and hydration property; 3. An Air Force Research Laboratory Chemical Ionization Mass Spectrometer to measure the concentration of SO2 and SO3/H2SO4; and 4. An Aerodyne Research Inc. Tunable Diode Laser System to measure the concentrations of SO2, SO3, NO, NO2, CO2, and H2O. By September 1997, an F100 engine operating at several power levels at sea level and up to six simulated altitudes had been tested with commercial jet fuels with three levels of sulfur content and one military jet fuel. The data are being vigorously analyzed. A complete report is anticipated for the 1998 Atmospheric Effects of Aviation Project Annual Conference.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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  • 20
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    Solar Electric Propulsion for Mars Exploration (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Highly propellant-efficient electric propulsion is being combined with advanced solar power technology to provide a non-nuclear transportation option for the human exploration of Mars. By virtue of its high specific impulse, electric propulsion offers a greater change in spacecraft velocity for each pound of propellant than do conventional chemical rockets. As a result, a mission to Mars based on solar electric propulsion (SEP) would require fewer heavy-lift launches than a traditional all-chemical space propulsion scenario would. Performance, as measured by mass to orbit and trip time, would be comparable to the NASA design reference mission for human Mars exploration, which utilizes nuclear thermal propulsion; but it would avoid the issues surrounding the use of nuclear reactors in space.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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  • 21
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    High-Temperature Magnetic Bearings Being Developed for Gas Turbine Engines (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Magnetic bearings are the subject of a new NASA Lewis Research Center and U.S. Army thrust with significant industry participation, and cooperation with other Government agencies. The NASA/Army emphasis is on high-temperature applications for future gas turbine engines. Magnetic bearings could increase the reliability and reduce the weight of these engines by eliminating the lubrication system. They could also increase the DN (diameter of bearing times the rpm) limit on engine speed and allow active vibration cancellation systems to be used, resulting in a more efficient, "more electric" engine. Finally, the Integrated High Performance Turbine Engine Technology (IHPTET) program, a joint Department of Defense/industry program, identified a need for a high-temperature (1200 F) magnetic bearing that could be demonstrated in their Phase III engine. This magnetic bearing is similar to an electric motor. It has a laminated rotor and stator made of cobalt steel. Wound around the stator's circumference are a series of electrical wire coils which form a series of electric magnets that exert a force on the rotor. A probe senses the position of the rotor, and a feedback controller keeps it centered in the cavity. The engine rotor, bearings, and casing form a flexible structure with many modes. The bearing feedback controller, which could cause some of these modes to become unstable, could be adapted to varying flight conditions to minimize seal clearances and monitor the health of the system.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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    NASA TECHNICAL REPORTS
  • 22
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    Jet Injection Used to Control Rotating Stall in a High-Speed Compressor (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: In a joint effort between the Massachusetts Institute of Technology (MIT) and the NASA Lewis Research Center, a new technology was demonstrated to identify and control rotating stall and surge in a single-stage, high-speed compressor. Through the use of highvelocity, high-frequency jet injectors, the instabilities of surge and stall were controlled in a high-speed compressor rig. Through the use of active stall control, modal instabilities that normally occur in the pressure measurements prior to stall were normalized and the range of the compressor was extended. Normally the events of rotating stall and surge instabilities limit the operation of the aeroengine compressor to a region below the surge line. To enhance the performance of the compressor, the Lewis/MIT team used active stall control methods to extend the normal operation of the compressor beyond the original stall point.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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    NASA TECHNICAL REPORTS
  • 23
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    Wave Rotor Research and Technology Development (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Wave rotor technology offers the potential to increase the performance of gas turbine engines significantly, within the constraints imposed by current material temperature limits. The wave rotor research at the NASA Lewis Research Center is a three-element effort: 1) Development of design and analysis tools to accurately predict the performance of wave rotor components; 2) Experiments to characterize component performance; 3) System integration studies to evaluate the effect of wave rotor topping on the gas turbine engine system.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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    NASA TECHNICAL REPORTS
  • 24
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    Lift Fan Nozzle for Joint Strike Fighter Tested in NASA Lewis' Powered Lift Rig (1998)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Under a nonreimbursable space act agreement between the NASA Lewis Research Center and the Allison Advanced Development Company, Allison tested a lift fan nozzle in Lewis' Powered Lift Rig. This test was in support of the Joint Strike Fighter program (formerly the Joint Advanced Strike Technology) sponsored by the Department of Defense, which will develop and field an affordable, multirole, next-generation, strike fighter aircraft for the Navy, Air Force, Marine Corps, and foreign allies. Allison, along with Pratt & Whitney Company, is part of the Lockheed Martin Corporation team that is scheduled to build a concept demonstrator aircraft by fiscal year 2001. The test was initiated in April and successfully completed in mid-July of 1997. Allison supplied a one-third-scale model of the lift fan nozzle, and Lewis provided the facility and the necessary support team. Various configurations, including pitching vectored angles ranging from 15deg forward to 60deg backward, were tested over a range of nozzle pressure ratios. Nozzle flow rates, thrust, and static pressures were measured for each of the configurations. Results from the test met the design requirements for the Joint Strike Fighter program and were in agreement with Allison's internal computational fluid dynamics (CFD) analyses. Data obtained from this test will also be used in the full-scale design of the lift fan system.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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  • 25
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    An Analysis of Thrust Normalization of Ground Flowfield Pressures, Temperatures, and Velocities, for an AV-8B -408 Harrier During Hover (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: An external environment test for an AV-8B Harrier during hover and vertical operations was conducted at NAWCAD at Patuxent River, Maryland in July 1997. Four boundary layer rakes were instrumented with static and total pressures, and thermocouples for measuring temperatures. These rakes were installed at 30, 50, 75, and 100 foot from the hover center. The 50 ft and 100 ft rakes were offset 20 deg from the other two to minimize interference effects. In order to measure a complete flowfield footprint, it was necessary to have the Harrier change its heading relative to the rakes from 0 to 180 deg. A 20 deg increment in azimuth was used. This permitted the four rakes to measure the flowfield at 72 locations relative to the aircraft. However, as the Harrier burns fuel, the hover thrust must be reduced by the pilot in order to maintain a constant height above ground. The typical test procedure employed was: (1) vertical takeoff at an initial heading; (2) 20 second hover dwell at that heading; (3) pedal turn to a second heading, followed by a 20 second dwell hover; (4) pedal turn to a third heading, followed by a 20 second dwell hover; and (5) vertical landing at the third heading. Additional information is contained in the original extended abstract.
    Keywords: Aircraft Propulsion and Power
    Type: International Powered Lift Conference; Sep 02, 1998 - Sep 04, 1998; London; United Kingdom
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    Integrated Fiber-Optic Light Probe: Measurement of Static Deflections in Rotating Turbomachinery (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: At the NASA Lewis Research Center, in cooperation with Integrated Fiber Optic Systems, Inc., an integrated fiber-optic light probe system was designed, fabricated, and tested for monitoring blade tip deflections, vibrations, and to some extent, changes in the blade tip clearances of a turbomachinery fan or a compressor rotor. The system comprises a set of integrated fiber-optic light probes that are positioned to detect the passing blade tip at the leading and trailing edges. In this configuration, measurements of both nonsynchronous blade vibrations and steady-state blade deflections can be made from the timing information provided by each light probe-consisting of an integrated fiber-optic transmitting channel and numerical aperture receiving fibers, all mounted in the same cylindrical housing. With integrated fiber-optic technology, a spatial resolution of 50 mm is possible while the outer diameter is kept below 2.5 mm. To evaluate these probes, we took measurements in a single-stage compressor facility and an advanced fan rig in Lewis' 9- by 15-Foot Low-Speed Wind Tunnel.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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    Experimental Study of a Nozzle Using Fluidic Counterflow for Thrust Vectoring (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A static experimental investigation of a counterflow thrust vectoring nozzle concept was performed. The study was conducted in the NASA Langley Research Center Jet Exit Test Facility. Internal performance characteristics were defined over a nozzle pressure ratio (jet total to ambient) range of 3.5 to 10.0. The effects of suction collar geometry and suction slot height on nozzle performance were examined. In the counterflow concept, thrust vectoring is achieved by applying a vacuum to a slot adjacent to a primary jet that is shrouded by a suction collar. Two flow phenomena work to vector the primary jet depending upon the test conditions and configuration. In one case, the vacuum source creates a secondary reverse flowing stream near the primary jet. The shear layers between the two counterflowing streams mix and entrain mass from the surrounding fluid. The presence of the collar inhibits mass entrainment and the flow near the collar accelerates, causing a drop in pressure on the collar. The second case works similarly except that the vacuum is not powerful enough to create a counterflowing stream and instead a coflowing stream is present. The primary jet is vectored if suction is applied asymmetrically on the top or bottom of the jet.
    Keywords: Aircraft Propulsion and Power
    Type: AIAA Paper 98-3255 , 34th AIAA/ASME SAE/ASEE Joint Propulsion Conference and Exhibit; Jul 13, 1998 - Jul 15, 1998; Cleveland, OH; United States
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    Data Mining and Optimization Tools for Developing Engine Parameters Tools (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This project was awarded for understanding the problem and developing a plan for Data Mining tools for use in designing and implementing an Engine Condition Monitoring System. Tricia Erhardt and I studied the problem domain for developing an Engine Condition Monitoring system using the sparse and non-standardized datasets to be available through a consortium at NASA Lewis Research Center. We visited NASA three times to discuss additional issues related to dataset which was not made available to us. We discussed and developed a general framework of data mining and optimization tools to extract useful information from sparse and non-standard datasets. These discussions lead to the training of Tricia Erhardt to develop Genetic Algorithm based search programs which were written in C++ and used to demonstrate the capability of GA algorithm in searching an optimal solution in noisy, datasets. From the study and discussion with NASA LeRC personnel, we then prepared a proposal, which is being submitted to NASA for future work for the development of data mining algorithms for engine conditional monitoring. The proposed set of algorithm uses wavelet processing for creating multi-resolution pyramid of tile data for GA based multi-resolution optimal search.
    Keywords: Aircraft Propulsion and Power
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  • 29
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    Three-Component Force Measurements on a Scramjet in a Reflected-Shock Tunnel (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A three-component stress-wave force-balance for a large scramjet has been designed, calibrated and tested in the HYPULSE reflected shock tunnel at GASL Inc., New York. The scramjet model is over 3-foot long and weighs in excess of 90 Ibm. The stress-wave force-balance is comprised of three stress bars which are attached to the model. Calibration results indicate that the force balance responds well within about 1 ms and that the sensitivity of the balance to the distribution of load is not large. Results with and without fuel injection were obtained in the tunnel operated for Mach 7 and Mach 10 flight simulation. These tests showed the force-balance can resolve axial force increments due to combustion of about 40 lb in the presence of model lift forces of 500-700 lb.
    Keywords: Aircraft Propulsion and Power
    Type: JANNAF Airbreathing Propulsion Subcommittee and 35th Combustion Subcommittee Meeting; 1; 35-52; CPIA-Publ-682-Vol-1
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  • 30
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    Efficient Aeroelastic Analysis of Propulsion Turbomachinery (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The purpose of this report is to provide a description of the overall thrust, past relevant accomplishments and future work for research that was performed on the titled project. The purpose of this work was to perform applied research at the NASA Lewis Research Center in the area of the aeroelastic behavior of turbomachinery used in air-breathing propulsion. This activity was undertaken in close cooperation with the Structural Dynamics Branch of the Structures Division at NASA Lewis.
    Keywords: Aircraft Propulsion and Power
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  • 31
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    Neural Net-Based Redesign of Transonic Turbines for Improved Unsteady Aerodynamic Performance (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A recently developed neural net-based aerodynamic design procedure is used in the redesign of a transonic turbine stage to improve its unsteady aerodynamic performance. The redesign procedure used incorporates the advantages of both traditional response surface methodology (RSM) and neural networks by employing a strategy called parameter-based partitioning of the design space. Starting from the reference design, a sequence of response surfaces based on both neural networks and polynomial fits are constructed to traverse the design space in search of an optimal solution that exhibits improved unsteady performance. The procedure combines the power of neural networks and the economy of low-order polynomials (in terms of number of simulations required and network training requirements). A time-accurate, two-dimensional, Navier-Stokes solver is used to evaluate the various intermediate designs and provide inputs to the optimization procedure. The optimization procedure yields a modified design that improves the aerodynamic performance through small changes to the reference design geometry. The computed results demonstrate the capabilities of the neural net-based design procedure, and also show the tremendous advantages that can be gained by including high-fidelity unsteady simulations that capture the relevant flow physics in the design optimization process.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208754 , A-9900398 , NAS 1.15:208754 , Propulsion; Jun 20, 1999 - Jun 24, 1999; Los Angeles, CA; United States
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  • 32
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    Challenges to Laser-Based Imaging Techniques in Gas Turbine Combustor Systems for Aerospace Applications (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Increasingly severe constraints on emissions, noise and fuel efficiency must be met by the next generation of commercial aircraft powerplants. At NASA Lewis Research Center (LeRC) a cooperative research effort with industry is underway to design and test combustors that will meet these requirements. To accomplish these tasks, it is necessary to gain both a detailed understanding of the combustion processes and a precise knowledge of combustor and combustor sub-component performance at close to actual conditions. To that end, researchers at LeRC are engaged in a comprehensive diagnostic investigation of high pressure reacting flowfields that duplicate conditions expected within the actual engine combustors. Unique, optically accessible flame-tubes and sector rig combustors, designed especially for these tests. afford the opportunity to probe these flowfields with the most advanced, laser-based optical diagnostic techniques. However, these same techniques, tested and proven on comparatively simple bench-top gaseous flame burners, encounter numerous restrictions and challenges when applied in these facilities. These include high pressures and temperatures, large flow rates, liquid fuels, remote testing, and carbon or other material deposits on combustor windows. Results are shown that document the success and versatility of these nonintrusive optical diagnostics despite the challenges to their implementation in realistic systems.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208649 , NAS 1.15:208649 , AIAA Paper 98-2778 , E-11368 , Advanced Measurement and Ground Testing Technology Conference; Jun 15, 1998 - Jun 18, 1998; Albuquerque, NM; United States
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    Initial Flight Test Evaluation of the F-15 ACTIVE Axisymmetric Vectoring Nozzle Performance (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A full envelope database of a thrust-vectoring axisymmetric nozzle performance for the Pratt & Whitney Pitch/Yaw Balance Beam Nozzle (P/YBBN) is being developed using the F-15 Advanced Control Technology for Integrated Vehicles (ACTIVE) aircraft. At this time, flight research has been completed for steady-state pitch vector angles up to 20' at an altitude of 30,000 ft from low power settings to maximum afterburner power. The nozzle performance database includes vector forces, internal nozzle pressures, and temperatures all of which can be used for regression analysis modeling. The database was used to substantiate a set of nozzle performance data from wind tunnel testing and computational fluid dynamic analyses. Findings from initial flight research at Mach 0.9 and 1.2 are presented in this paper. The results show that vector efficiency is strongly influenced by power setting. A significant discrepancy in nozzle performance has been discovered between predicted and measured results during vectoring.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206558 , H-2267 , NAS 1.15:206558 , AIAA Paper 98-3871 , Propulsion; Jul 13, 1998 - Jul 15, 1998; Cleveland, OH; United States
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    Effects of Tip Clearance and Casing Recess on Heat Transfer and Stage Efficiency in Axial Turbines (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Calculations were performed to assess the effect of the tip leakage flow on the rate of heat transfer to blade, blade tip and casing. The effect on exit angle and efficiency was also examined. Passage geometries with and without casing recess were considered. The geometry and the flow conditions of the GE-E 3 first stage turbine, which represents a modem gas turbine blade were used for the analysis. Clearance heights of 0%, 1%, 1.5% and 3% of the passage height were considered. For the two largest clearance heights considered, different recess depths were studied. There was an increase in the thermal load on all the heat transfer surfaces considered due to enlargement of the clearance gap. Introduction of recessed casing resulted in a drop in the rate of heat transfer on the pressure side but the picture on the suction side was found to be more complex for the smaller tip clearance height considered. For the larger tip clearance height the effect of casing recess was an orderly reduction in the suction side heat transfer as the casing recess height was increased. There was a marked reduction of heat load and peak values on the blade tip upon introduction of casing recess, however only a small reduction was observed on the casing itself. It was reconfirmed that there is a linear relationship between the efficiency and the tip gap height. It was also observed that the recess casing has a small effect on the efficiency but can have a moderating effect on the flow underturning at smaller tip clearances.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-208514 , E-11287 , NAS 1.26:208514 , ICOMP-98-04 , Turbo; Jun 02, 1998 - Jun 05, 1998; Stockholm; Sweden
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    Energy Efficient Engine Low Pressure Subsystem Aerodynamic Analysis (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The objective of this study was to demonstrate the capability to analyze the aerodynamic performance of the complete low pressure subsystem (LPS) of the Energy Efficient Engine (EEE). Detailed analyses were performed using three- dimensional Navier-Stokes numerical models employing advanced clustered processor computing platforms. The analysis evaluates the impact of steady aerodynamic interaction effects between the components of the LPS at design and off- design operating conditions. Mechanical coupling is provided by adjusting the rotational speed of common shaft-mounted components until a power balance is achieved. The Navier-Stokes modeling of the complete low pressure subsystem provides critical knowledge of component acro/mechanical interactions that previously were unknown to the designer until after hardware testing.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208402 , E-11234 , NAS 1.15:208402 , AIAA Paper 98-3119 , Propulsion; Jul 13, 1998 - Jul 15, 1998; Cleveland, OH; United States
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    Calculation of Multistage Turbomachinery Using Steady Characteristic Boundary Conditions (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A multiblock Navier-Stokes analysis code for turbomachinery has been modified to allow analysis of multistage turbomachines. A steady averaging-plane approach was used to pass information between blade rows. Characteristic boundary conditions written in terms of perturbations about the mean flow from the neighboring blade row were used to allow close spacing between the blade rows without forcing the flow to be axisymmetric. In this report the multiblock code is described briefly and the characteristic boundary conditions and the averaging-plane implementation are described in detail. Two approaches for averaging the flow properties are also described. A two-dimensional turbine stator case was used to compare the characteristic boundary conditions with standard axisymmetric boundary conditions. Differences were apparent but small in this low-speed case. The two-stage fuel turbine used on the space shuttle main engines was then analyzed using a three-dimensional averaging-plane approach. Computed surface pressure distributions on the stator blades and endwalls and computed distributions of blade surface heat transfer coefficient on three blades showed very good agreement with experimental data from two tests.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206613 , E-11033 , NAS 1.15:206613 , AIAA Paper 98-0968 , Aerospace Sciences Meeting and Exhibit; Jan 12, 1998 - Jan 15, 1998; Reno, NV; United States
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    Flashback Arrestor for LPP, Low NOx Combustors (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-27
    Description: Lean premixed, prevaporized (LPP) high temperature combustor designs as explored for the Advanced Subsonic Transport (AST) and High Speed Civil Transport (HSCT) combustors can achieve low NO(x), emission levels. An enabling device is needed to arrest flashback and inhibit preignition at high power conditions and during transients (surge and rapid spool down). A novel flashback arrestor design has demonstrated the ability to arrest flashback and inhibit preignition in a 4.6 cm diameter tubular reactor at full power inlet temperatures (725 C) using Jet-A fuel at 0.4 less than or equal To phi less than or equal to 3.5. Several low pressure loss (0.2 to 0.4% at 30 m/s) flashback arrestor designs were developed which arrested flashback at all of the test conditions. Flame holding was also inhibited off the flash arrestor face or within the downstream tube even velocities (less than or equal to 3 to 6 m/s), thus protecting the flashback arrestor and combustor components. Upstream flow conditions influence the specific configuration based on using either a 45% or 76% upstream geometric blockage. Stationary, lean premixed dry low NO(x) gas turbine combustors would also benefit from this low pressure drop flashback arrestor design which can be easily integrated into new and existing designs.
    Keywords: Aircraft Propulsion and Power
    Type: Turbo Expo ''98 Congress and Exhibition; 2-5 Ju. 1998; Stockholm; Sweden
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    Experiments on Cruise Propulsion with a Hydrogen Scramjet (1998)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Measurement of drag have been made in a shock tunnel on a simple integrated vehicle engine combination for hypersonic cruise with hydrogen scramjet propulsion. The test flow Mach number was 6.4, and the velocity was 2.45 kms(exp -1). Zero Drag, which is the necessary condition for cruise, was achieved as the equivalence ratio approached one. It was found that an analysis using established aerodynamic concept was adequate for predicting drag in the case of no combustion. When combustion occurred results of direct connect experiments provided was qualitative guide to the measured levels of drag, and indicated that thrust nozzle combustion was taking place. An heuristic analysis is used to point to the important effect this may have on propulsive lift.
    Keywords: Aircraft Propulsion and Power
    Type: Paper 2250 , The Aeronautical Journal; 102; 1011; 37-43
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    JANNAF Airbreathing Propulsion Subcommittee and 35th Combustion Subcommittee Meeting (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-08-13
    Description: This document, CPIA Publication 682, Volume 1, is a compilation of 5 unclassified/unlimited technical papers (approved for public release) which were presented at the 1 998 meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Airbreathing Propulsion Subcommittee (APS) and Combustion Subcommittee (CS) held jointly with the Propulsion Systems Hazards Subcommittee (PSHS). The meeting was held on 7-11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include HyTech technology development, hydrocarbon fuel development for hypersonic applications, pulse detonation propulsion system development and arc heaters for direct-connect scramjet testing.
    Keywords: Aircraft Propulsion and Power
    Type: CPIA-Publ-682-Vol-1 , Airbreathing Propulsion Subcommittee; Dec 07, 1998 - Dec 11, 1998; Tucson, AZ; United States|Combustion Subcommittee; Dec 07, 1998 - Dec 11, 1998; Tucson, AZ; United States
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    The Control System for the X-33 Linear Aerospike Engine (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The linear aerospike engine is being developed for single-stage -to-orbit (SSTO) applications. The primary advantages of a linear aerospike engine over a conventional bell nozzle engine include altitude compensation, which provides enhanced performance, and lower vehicle weight resulting from the integration of the engine into the vehicle structure. A feature of this integration is the ability to provide thrust vector control (TVC) by differential throttling of the engine combustion elements, rather than the more conventional approach of gimballing the entire engine. An analysis of the X-33 flight trajectories has shown that it is necessary to provide +/- 15% roll, pitch and yaw TVC authority with an optional capability of +/- 30% pitch at select times during the mission. The TVC performance requirements for X-33 engine became a major driver in the design of the engine control system. The thrust level of the X-33 engine as well as the amount of TVC are managed by a control system which consists of electronic, instrumentation, propellant valves, electro-mechanical actuators, spark igniters, and harnesses. The engine control system is responsible for the thrust control, mixture ratio control, thrust vector control, engine health monitoring, and communication to the vehicle during all operational modes of the engine (checkout, pre-start, start, main-stage, shutdown and post shutdown). The methodology for thrust vector control, the health monitoring approach which includes failure detection, isolation, and response, and the basic control system design are the topic of this paper. As an additional point of interest a brief description of the X-33 engine system will be included in this paper.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-207923 , NAS 1.26:207923
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    Small Engine Technology (Set) Task 8 Aeroelastic Prediction Methods (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: AlliedSignal Engines, in cooperation with NASA LeRC, completed an evaluation of recently developed aeroelastic computer codes using test cases from the AlliedSignal Engines fan blisk database. Test data for this task includes strain gage, light probe, performance, and steady-state pressure information obtained for conditions where synchronous or flutter vibratory conditions were found to occur. Aeroelastic codes evaluated include the quasi 3-D UNSFLO (developed at MIT and modified to include blade motion by AlliedSignal), the 2-D FREPS (developed by NASA LeRC), and the 3-D TURBO-AE (under development at NASA LeRC). Six test cases each where flutter and synchronous vibrations were found to occur were used for evaluation of UNSFLO and FREPS. In addition, one of the flutter cases was evaluated using TURBO-AE. The UNSFLO flutter evaluations were completed for 75 percent radial span and provided good agreement with the experimental test data. Synchronous evaluations were completed for UNSFLO but further enhancement needs to be added to the code before the unsteady pressures can be used to predict forced response vibratory stresses. The FREPS evaluations were hindered as the steady flow solver (SFLOW) was unable to converge to a solution for the transonic flow conditions in the fan blisk. This situation resulted in all FREPS test cases being attempted but no results were obtained during the present program. Currently, AlliedSignal is evaluating integrating FREPS with our existing steady flow solvers to bypass the SFLOW difficulties. ne TURBO-AE steady flow solution provided an excellent match with the AlliedSignal Engines calibrated DAWES 3-D viscous solver. Finally, the TURBO-AE unsteady analyses also matched experimental observations by predicting flutter for the single test case evaluated.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-202328 , NAS 1.26:202328 , E-10674 , Rept-21-9157
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    Carbon-carbon cylinder block (1998)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: A lightweight cylinder block composed of carbon-carbon is disclosed. The use of carbon-carbon over conventional materials, such as cast iron or aluminum, reduces the weight of the cylinder block and improves thermal efficiency of the internal combustion reciprocating engine. Due to the negligible coefficient of thermal expansion and unique strength at elevated temperatures of carbon-carbon, the piston-to-cylinder wall clearance can be small, especially when the carbon-carbon cylinder block is used in conjunction with a carbon-carbon piston. Use of the carbon-carbon cylinder block has the effect of reducing the weight of other reciprocating engine components allowing the piston to run at higher speeds and improving specific engine performance.
    Keywords: Aircraft Propulsion and Power
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    A Method for Generating Reduced-Order Linear Models of Multidimensional Supersonic Inlets (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Simulation of high speed propulsion systems may be divided into two categories, nonlinear and linear. The nonlinear simulations are usually based on multidimensional computational fluid dynamics (CFD) methodologies and tend to provide high resolution results that show the fine detail of the flow. Consequently, these simulations are large, numerically intensive, and run much slower than real-time. ne linear simulations are usually based on large lumping techniques that are linearized about a steady-state operating condition. These simplistic models often run at or near real-time but do not always capture the detailed dynamics of the plant. Under a grant sponsored by the NASA Lewis Research Center, Cleveland, Ohio, a new method has been developed that can be used to generate improved linear models for control design from multidimensional steady-state CFD results. This CFD-based linear modeling technique provides a small perturbation model that can be used for control applications and real-time simulations. It is important to note the utility of the modeling procedure; all that is needed to obtain a linear model of the propulsion system is the geometry and steady-state operating conditions from a multidimensional CFD simulation or experiment. This research represents a beginning step in establishing a bridge between the controls discipline and the CFD discipline so that the control engineer is able to effectively use multidimensional CFD results in control system design and analysis.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-207405 , E-11164 , NAS 1.26:207405
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    Study of Low Reynolds Number Effects on the Losses in Low-Pressure Turbine Blade Rows (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Experimental data from jet-engine tests have indicated that unsteady blade row interactions and separation can have a significant impact on the efficiency of low-pressure turbine stages. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Several recent studies have revealed that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study numerical experiments have been performed to study the models available for low Reynolds number flows, and to quantify the Reynolds number dependence of low-pressure turbine cascades and stages. The predicted aerodynamic results exhibit good agreement with design data.
    Keywords: Aircraft Propulsion and Power
    Type: AIAA Paper 98-3575 , Joint Propulsion; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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    Damage Tolerance and Reliability of Turbine Engine Components (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: A formal method is described to quantify structural damage tolerance and reliability in the presence of multitude of uncertainties in turbine engine components. The method is based at the materials behavior level where primitive variables with their respective scatters are used to describe that behavior. Computational simulation is then used to propagate those uncertainties to the structural scale where damage tolerance and reliability are usually specified. Several sample cases are described to illustrate the effectiveness, versatility, and maturity of the method. Typical results from these methods demonstrate that the methods are mature and that they can be used for future strategic projections and planning to assure better, cheaper, faster products for competitive advantages in world markets. These results also indicate that the methods are suitable for predicting remaining life in aging or deteriorating structures.
    Keywords: Aircraft Propulsion and Power
    Type: SC.78: Qualification of Life Extension Schemes for Engine Components; Oct 05, 1998 - Oct 09, 1998; Corfu; Greece
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    Analysis of Inlet-Compressor Acoustic Interactions Using Coupled CFD Codes (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A problem that arises in the numerical simulation of supersonic inlets is the lack of a suitable boundary condition at the engine face. In this paper, a coupled approach, in which the inlet computation is coupled dynamically to a turbomachinery computation, is proposed as a means to overcome this problem. The specific application chosen for validation of this approach is the collapsing bump experiment performed at the University of Cincinnati. The computed results are found to be in reasonable agreement with experimental results. The coupled simulation results could also be used to aid development of a simplified boundary condition.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208839 , NAS 1.15:208839 , AIAA Paper 98-0749 , E-11451 , Aerospace Sciences; Jan 11, 1999 - Jan 14, 1999; Reno, NV; United States
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    Engine Technology Challenges for the High-Speed Civil Transport Plane (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Ongoing NASA-funded and privately funded studies continue to indicate that an opportunity exists for a second generation supersonic commercial airliner, or High-Speed Civil Transport (HSCT), to become a key part of the 21 st century international air transportation system. Long distance air travel is projected to be the fastest growing segment of the air transportation market by the turn of the century with increases at about 5 percent per annum over the next two decades. This projection suggests that by the year 2015, more than 600,000 passengers per day will be traveling long distances, predominantly over water. These routes would provide the greatest potential for an HSCT to become a significant part of the international air transportation system. The potential market for an HSCT is currently projected to be anywhere from 500-1500 aircraft over the 2005-2030 time period. Such an aircraft fleet size would represent a considerable share of the potential long-range aircraft market. However, this projected HSCT fleet can become a reality only if technologies are developed which will allow an HSCT design that is (1) environmentally compatible and (2) economically viable. Simply stated, the HSCT will be a technology driven airplane. Without significant advances in airframe and propulsion technologies over the levels currently available, there will be no second generation supersonic airliner! This paper will briefly describe the propulsion technology challenges which must be met prior to any product launch decision being made by industry and the progress toward meeting these challenges through NASAs High-Speed Research (HSR) Program, a partnership between NASA and Boeing, General Electric and Pratt & Whitney.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208405 , E-11238 , NAS 1.15:208405 , AIAA Paper 98-2505 , Advanced Measurement and Ground Testing Technology; Jun 15, 1998 - Jun 18, 1998; Albuquerque, NM; United States
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    High Stability Engine Control (HISTEC): Flight Demonstration Results (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Future aircraft turbine engines, both commercial and military, must be able to accommodate expected increased levels of steady-state and dynamic engine-face distortion. The current approach of incorporating sufficient design stall margin to tolerate these increased levels of distortion would significantly reduce performance. The High Stability Engine Control (HISTEC) program has developed technologies for an advanced, integrated engine control system that uses measurement- based estimates of distortion to enhance engine stability. The resulting distortion tolerant control reduces the required design stall margin, with a corresponding increase in performance and/or decrease in fuel burn. The HISTEC concept was successfully flight demonstrated on the F-15 ACTIVE aircraft during the summer of 1997. The flight demonstration was planned and carried out in two parts, the first to show distortion estimation, and the second to show distortion accommodation. Post-flight analysis shows that the HISTEC technologies are able to successfully estimate and accommodate distortion, transiently setting the stall margin requirement on-line and in real-time. Flight demonstration of the HISTEC technologies has significantly reduced the risk of transitioning the technology to tactical and commercial engines.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208655 , NAS 1.15:208655 , SAE-985556 , E-11375 , 1998 World Aviation Congress and Exposition; Sep 28, 1998 - Sep 30, 1998; Anaheim, CA; United States
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    Development and Testing of a High Stability Engine Control (HISTEC) System (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Flight tests were recently completed to demonstrate an inlet-distortion-tolerant engine control system. These flight tests were part of NASA's High Stability Engine Control (HISTEC) program. The objective of the HISTEC program was to design, develop, and flight demonstrate an advanced integrated engine control system that uses measurement-based, real-time estimates of inlet airflow distortion to enhance engine stability. With improved stability and tolerance of inlet airflow distortion, future engine designs may benefit from a reduction in design stall-margin requirements and enhanced reliability, with a corresponding increase in performance and decrease in fuel consumption. This paper describes the HISTEC methodology, presents an aircraft test bed description (including HISTEC-specific modifications) and verification and validation ground tests. Additionally, flight test safety considerations, test plan and technique design and approach, and flight operations are addressed. Some illustrative results are presented to demonstrate the type of analysis and results produced from the flight test program.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206562 , H-2269 , NAS 1.15:206562 , AIAA Paper 98-3715 , Propulsion; Jul 13, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 50
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    High Stability Engine Control (HISTEC) Flight Test Results (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The High Stability Engine Control (HISTEC) Program, managed and funded by the NASA Lewis Research Center, is a cooperative effort between NASA and Pratt & Whitney (P&W). The program objective is to develop and flight demonstrate an advanced high stability integrated engine control system that uses real-time, measurement-based estimation of inlet pressure distortion to enhance engine stability. Flight testing was performed using the NASA Advanced Controls Technologies for Integrated Vehicles (ACTIVE) F-15 aircraft at the NASA Dryden Flight Research Center. The flight test configuration, details of the research objectives, and the flight test matrix to achieve those objectives are presented. Flight test results are discussed that show the design approach can accurately estimate distortion and perform real-time control actions for engine accommodation.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208481 , E-11255 , NAS 1.15:208481 , AIAA Paper 98-3757 , Propulsion; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 51
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    Model Engine Performance Measurement From Force Balance Instrumentation (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A large scale model representative of a low-noise, high bypass ratio turbofan engine was tested for acoustics and performance in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel. This test was part of NASA's continuing Advanced Subsonic Technology Noise Reduction Program. The low tip speed fan, nacelle, and an un-powered core passage (with core inlet guide vanes) were simulated. The fan blades and hub are mounted on a rotating thrust and torque balance. The nacelle, bypass duct stators, and core passage are attached to a six component force balance. The two balance forces, when corrected for internal pressure tares, measure the total thrust-minus-drag of the engine simulator. Corrected for scaling and other effects, it is basically the same force that the engine supports would feel, operating at similar conditions. A control volume is shown and discussed, identifying the various force components of the engine simulator thrust and definitions of net thrust. Several wind tunnel runs with nearly the same hardware installed are compared, to identify the repeatability of the measured thrust-minus-drag. Other wind tunnel runs, with hardware changes that affected fan performance, are compared to the baseline configuration, and the thrust and torque effects are shown. Finally, a thrust comparison between the force balance and nozzle gross thrust methods is shown, and both yield very similar results.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208486 , E-11263 , NAS 1.15:208486 , AIAA Paper 98-3112 , Propulsion; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 52
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    General Aviation Light Aircraft Propulsion: From the 1940's to the Next Century (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Current general aviation light aircraft are powered by engines that were originally designed in the 1940's. This paper gives a brief history of light aircraft engine development, explaining why the air-cooled, horizontally opposed piston engine became the dominant engine for this class of aircraft. Current engines are fairly efficient, and their designs have been updated through the years, but their basic design and operational characteristics are archaic in comparison to modem engine designs, such as those used in the automotive industry. There have been some innovative engine developments, but in general they have not been commercially successful. This paper gives some insight into the reasons for this lack of success. There is now renewed interest in developing modem propulsion systems for light aircraft, in the fore-front of which is NASA's General Aviation Propulsion (GAP) program. This paper gives an overview of the engines being developed in the GAP program, what they will mean to the general aviation community, and why NASA and its industry partners believe that these new engine developments will bring about a new era in general aviation light aircraft.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208411 , NAS 1.15:208411 , E-11246 , AIAA Paper 98-3116 , Joint Propulsion Conference; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 53
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    The High Stability Engine Control (HISTEC) Program: Flight Demonstration Phase (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Future aircraft turbine engines, both commercial and military, must be able to accommodate expected increased levels of steady-state and dynamic engine-face distortion. The current approach of incorporating sufficient design stall margin to tolerate these increased levels of distortion would significantly reduce performance. The objective of the High Stability Engine Control (HISTEC) program is to design, develop, and flight-demonstrate an advanced, integrated engine control system that uses measurement-based estimates of distortion to enhance engine stability. The resulting distortion tolerant control reduces the required design stall margin, with a corresponding increase in performance and decrease in fuel burn. The HISTEC concept has been developed and was successfully flight demonstrated on the F-15 ACTIVE aircraft during the summer of 1997. The flight demonstration was planned and carried out in two phases, the first to show distortion estimation, and the second to show distortion accommodation. Post-flight analysis shows that the HISTEC technologies are able to successfully estimate and accommodate distortion, transiently setting the stall margin requirement on-line and in real-time. This allows the design stall margin requirement to be reduced, which in turn can be traded for significantly increased performance and/or decreased weight. Flight demonstration of the HISTEC technologies has significantly reduced the risk of transitioning the technology to tactical and commercial engines.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208482 , E-11257 , NAS 1.15:208482 , AIAA Paper 98-3756 , Joint Propulsion Conference; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 54
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    A Modified Through-Flow Wave Rotor Cycle with Combustor Bypass Ducts (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A wave rotor cycle is described which avoids the inherent problem of combustor exhaust gas recirculation (EGR) found in four-port, through-flow wave rotor cycles currently under consideration for topping gas turbine engines. The recirculated hot gas is eliminated by the judicious placement of a bypass duct which transfers gas from one end of the rotor to the other. The resulting cycle, when analyzed numerically, yields an absolute mean rotor temperature 18% below the already impressive value of the conventional four-port cycle (approximately the turbine inlet temperature). The absolute temperature of the gas leading to the combustor is also reduced from the conventional four-port design by 22%. The overall design point pressure ratio of this new bypass cycle is approximately the same as the conventional four-port cycle. This paper will describe the EGR problem and the bypass cycle solution including relevant wave diagrams. Performance estimates of design and off-design operation of a specific wave rotor will be presented. The results were obtained using a one-dimensional numerical simulation and design code.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206971 , NAS 1.15:206971 , E-10987-1 , AIAA Paper 97-3140 , Joint Propulsion Conference and Exhibit; Jul 06, 1997 - Jul 09, 1997; Seattle, WA; United States
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  • 55
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    Application of Pressure Sensitive Paint to Confined Flow at Mach Number 2.5 (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Pressure sensitive paint (PSP) is a novel technology that is being used frequently in external aerodynamics. For internal flows in narrow channels, and applications at elevated nonuniform temperatures, however, there are still unresolved problems that complicate the procedures for calibrating PSP signals. To address some of these problems, investigations were carried out in a narrow channel with supersonic flows of Mach 2.5. The first set of tests focused on the distribution of the wall pressure in the diverging section of the test channel downstream of the nozzle throat. The second set dealt with the distribution of wall static pressure due to the shock/wall interaction caused by a 25 deg. wedge in the constant Mach number part of the test section. In addition, the total temperature of the flow was varied to assess the effects of temperature on the PSP signal. Finally, contamination of the pressure field data, caused by internal reflection of the PSP signal in a narrow channel, was demonstrated. The local wall pressures were measured with static taps, and the wall pressure distributions were acquired by using PSP. The PSP results gave excellent qualitative impressions of the pressure field investigated. However, the quantitative results, specifically the accuracy of the PSP data in narrow channels, show that improvements need to be made in the calibration procedures, particularly for heated flows. In the cases investigated, the experimental error had a standard deviation of +/- 8.0% for the unheated flow, and +/- 16.0% for the heated flow, at an average pressure of 11 kpa.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-107527 , E-10842 , NAS 1.15:107527 , AIAA Paper 97-3214 , Joint Propulsion Conference; Jul 06, 1997 - Jul 09, 1997; Seattle, WA; United States
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  • 56
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    PSP Measurement of Stator Vane Surface Pressures in a High Speed Fan (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents measurements of static pressures on the stator vane suction side of a high-speed single stage fan using the technique of pressure sensitive paint (PSP). The paper illustrates development in application of the relatively new experimental technique to the complex environment of internal flows in turbomachines. First, there is a short explanation of the physics of the PSP technique and a discussion of calibration methods for pressure sensitive paint in the turbomachinery environment. A description of the image conversion process follows. The recorded image of the stator vane pressure field is skewed due to the limited optical access and must be converted to the meridional plane projection for comparison with analytical predictions. The experimental results for seven operating conditions along an off-design rotational speed line are shown in a concise form, including performance map points, mindspan static tap pressure distributions, and vane suction side pressure fields. Then, a comparison between static tap and pressure sensitive paint data is discussed. Finally, the paper lists shortcomings of the pressure sensitive paint technology and lessons learned in this high-speed fan application.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-207403 , E-11150 , NAS 1.26:207403 , Gas Turbine and Aeroengine; Jun 02, 1998 - Jun 05, 1998; Stockholm; Sweden
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  • 57
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    Propulsion System for Very High Altitude Subsonic Unmanned Aircraft (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper explains why a spark ignited gasoline engine, intake pressurized with three cascaded stages of turbocharging, was selected to power NASA's contemplated next generation of high altitude atmospheric science aircraft. Beginning with the most urgent science needs (the atmospheric sampling mission) and tracing through the mission requirements which dictate the unique flight regime in which this aircraft has to operate (subsonic flight at greater then 80 kft) we briefly explore the physical problems and constraints, the available technology options and the cost drivers associated with developing a viable propulsion system for this highly specialized aircraft. The paper presents the two available options (the turbojet and the turbocharged spark ignited engine) which are discussed and compared in the context of the flight regime. We then show how the unique nature of the sampling mission, coupled with the economic considerations pursuant to aero engine development, point to the spark ignited engine as the only cost effective solution available. Surprisingly, this solution compares favorably with the turbojet in the flight regime of interest. Finally, some remarks are made about NASA's present state of development, and future plans to flight demonstrate the three stage turbocharged powerplant.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206636 , NAS 1.15:206636 , E-11101 , Aerospace Power Systems 1998; Apr 21, 1998 - Apr 23, 1998; Williamsburg, VA; United States
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  • 58
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    Selected Performance Measurements of the F-15 Active Axisymmetric Thrust-vectoring Nozzle (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Flight tests recently completed at the NASA Dryden Flight Research Center evaluated performance of a hydromechanically vectored axisymmetric nozzle onboard the F-15 ACTIVE. A flight-test technique whereby strain gages installed onto engine mounts provided for the direct measurement of thrust and vector forces has proven to be extremely valuable. Flow turning and thrust efficiency, as well as nozzle static pressure distributions were measured and analyzed. This report presents results from testing at an altitude of 30,000 ft and a speed of Mach 0.9. Flow turning and thrust efficiency were found to be significantly different than predicted, and moreover, varied substantially with power setting and pitch vector angle. Results of an in-flight comparison of the direct thrust measurement technique and an engine simulation fell within the expected uncertainty bands. Overall nozzle performance at this flight condition demonstrated the F100-PW-229 thrust-vectoring nozzles to be highly capable and efficient.
    Keywords: Aircraft Propulsion and Power
    Type: International Symposium on Airbreathing Engines (ISABE); Sep 05, 1999 - Sep 10, 1999; Florence; Italy
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  • 59
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    Model-Based Fault Diagnosis for Turboshaft Engines (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Tests are described which, when used to augment the existing periodic maintenance and pre-flight checks of T700 engines, can greatly improve the chances of uncovering a problem compared to the current practice. These test signals can be used to expose and differentiate between faults in various components by comparing the responses of particular engine variables to the expected. The responses can be processed on-line in a variety of ways which have been shown to reveal and identify faults. The combination of specific test signals and on-line processing methods provides an ad hoc approach to the isolation of faults which might not otherwise be detected during pre-flight checkout.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208825 , NAS 1.15:208825 , ARL-TR-1447 , E-11433 , Fault Detection, Supervision and Safety for Technical Processes - SAFEPROCESS ''97; Aug 26, 1997 - Aug 28, 1997; Kingston Upon Hull; United Kingdom
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  • 60
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    A Simplified Model for the Investigation of Acoustically Driven Combustion Instabilities (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A simplified one-dimensional model of reactive flow is presented which captures features of aeropropulsion systems, including acoustically driven combustion instabilities. Although the resulting partial differential equations are one dimensional, they qualitatively describe observed phenomena, including, resonant frequencies and the admission of both steady and unsteady behavior. A number of simulations are shown which exhibit both steady and unsteady behavior, including flame migration and thermo acoustic instabilities. Finally, we present examples of unsteady flow resulting from fuel modulation.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208488 , NAS 1.15:208488 , AIAA Paper 98-3764 , E-11267 , Joint Propulsion; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 61
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    MoSi2-Base Hybrid Composite Passed Engine Test (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The intermetallics compound molybdenum disilicide (MoSi2) is an attractive high-temperature structural material for advanced engine applications. It has excellent oxidation resistance, a high melting point, relatively low density, and high thermal conductivity, and it is easily machined. Past research'at the NASA Lewis Research Center has resulted in the development of a hybrid composite consisting of a MoSi2 matrix reinforced with silicon nitride (Si3N4) Particulate and silicon carbide (SiC) fibers. This composite has demonstrated attractive strength, toughness, thermal fatigue, and oxidation resistance, including resistance to "pest" oxidation. These properties attracted the interest of the Office of Naval Research and Pratt & Whitney, and a joint NASA/Navy/Pratt & Whitney effort was developed to continue to mature the MoSi2 Composite technology. A turbine blade outer air seal, which was part of the Integrated High Performance Turbine Engine Technology (IHPTET) program, was chosen as a first component on which to focus. The first tasks of the materials development effort were to develop improved processing methods to reduce costs and to use fine-diameter fibers that enable the manufacturing of complex shapes. Tape-casting methods were developed to fully infiltrate the fine SiC fibers with matrix powders. The resulting composites were hot pressed to 100-percent density. Composites with cross-plied fiber architectures with 30 vol. % hi-nicalon SiC fibers and 30 vol. % nitride particles are now made routinely and demonstrate a good balance of properties. The next task entailed the measurement of a wide variety of mechanical properties to confirm the suitability of this composite in engines. In particular, participants in this effort demonstrated that composites made with Hi-Nicalon fibers had strength and toughness properties equal to or better than those of the composites made with the large-diameter fibers that had been used previously. Another critically important property measured was impact resistance. Aircraft engine components require sufficient toughness to resist manufacturing defects, assembly damage, stress concentrations at notches, and foreign object damage. Engine company designers indicated that impact resistance would have to be measured before they would seriously consider these types of composites. The Charpy V-notch test was chosen to assess impact resistance, and both monolithic and composite versions Of MOSi2 were tested from -300 to 1400 C. The results (see the following graphs) show that nitride-particulate-reinforced MoSi2 exhibited impact resistance higher than that of many monolithic ceramics and intermetallics, and that the fiber-reinforced composites had even higher values, approaching that of cast superalloys.
    Keywords: Aircraft Propulsion and Power
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  • 62
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    Energy Efficient Engine Low Pressure Subsystem Flow Analysis (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The objective of this project is to provide the capability to analyze the aerodynamic performance of the complete low pressure subsystem (LPS) of the Energy Efficient Engine (EEE). The analyses were performed using three-dimensional Navier-Stokes numerical models employing advanced clustered processor computing platforms. The analysis evaluates the impact of steady aerodynamic interaction effects between the components of the LPS at design and off-design operating conditions. Mechanical coupling is provided by adjusting the rotational speed of common shaft-mounted components until a power balance is achieved. The Navier-Stokes modeling of the complete low pressure subsystem provides critical knowledge of component aero/mechanical interactions that previously were unknown to the designer until after hardware testing.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-206597 , NAS 1.26:206597 , E-11067
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  • 63
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    Flight Testing the Linear Aerospike SR-71 Experiment (LASRE) (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The design of the next generation of space access vehicles has led to a unique flight test that blends the space and flight research worlds. The new space vehicle designs, such as the X-33 vehicle and Reusable Launch Vehicle (RLV), are powered by linear aerospike rocket engines. Conceived of in the 1960's, these aerospike engines have yet to be flown, and many questions remain regarding aerospike engine performance and efficiency in flight. To provide some of these data before flying on the X-33 vehicle and the RLV, a spacecraft rocket engine has been flight-tested atop the NASA SR-71 aircraft as the Linear Aerospike SR-71 Experiment (LASRE). A 20 percent-scale, semispan model of the X-33 vehicle, the aerospike engine, and all the required fuel and oxidizer tanks and propellant feed systems have been mounted atop the SR-71 airplane for this experiment. A major technical objective of the LASRE flight test is to obtain installed-engine performance flight data for comparison to wind-tunnel results and for the development of computational fluid dynamics-based design methodologies. The ultimate goal of firing the aerospike rocket engine in flight is still forthcoming. An extensive design and development phase of the experiment hardware has been completed, including approximately 40 ground tests. Five flights of the LASRE and firing the rocket engine using inert liquid nitrogen and helium in place of liquid oxygen and hydrogen have been successfully completed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206567 , H-2280 , NAS 1.15:206567 , Sep 15, 1998 - Sep 17, 1998; Reno, NV; United States
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  • 64
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    Wave Augmented Diffusers for Centrifugal Compressors (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A conceptual device is introduced which would utilize unsteady wave motion to slow and turn flows in the diffuser section of a centrifugal compressor. The envisioned device would substantially reduce the size of conventional centrifugal diffusers by eliminating the relatively large ninety degree bend needed to turn the flow from the radial/tangential to the axial direction. The bend would be replaced by a wall and the flow would instead exit through a series of rotating ports located on a disk, adjacent to the diffuser hub, and fixed to the impeller shaft. The ports would generate both expansion and compression waves which would rapidly transition from the hub/shroud (axial) direction to the radial/tangential direction. The waves would in turn induce radial/tangential and axial flow. This paper presents a detailed description of the device. Simplified cycle analysis and performance results are presented which were obtained using a time accurate, quasi-one-dimensional CFD code with models for turning, port flow conditions, and losses due to wall shear stress. The results indicate that a periodic wave system can be established which yields diffuser performance comparable to a conventional diffuser. Discussion concerning feasibility, accuracy, and integration follow.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208480 , E-11253 , NAS 1.15:208480 , AIAA Paper 98-3401 , Propulsion; Jul 12, 1998 - Jul 15, 1998; Cleveland, OH; United States
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  • 65
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    Fuel Injector Patternation Evaluation in Advanced Liquid-Fueled, High Pressure, Gas Turbine Combustors, Using Nonintrusive Optical Diagnostic Techniques (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Planar laser-induced fluorescence (PLIF) imaging and planar Mie scattering are used to examine the fuel distribution pattern (patternation) for advanced fuel injector concepts in kerosene burning, high pressure gas turbine combustors. Three diverse fuel injector concepts for aerospace applications were investigated under a broad range of operating conditions. Fuel PLIF patternation results are contrasted with those obtained by planar Mie scattering. Further comparison is also made for one injector with data obtained through phase Doppler measurements. Differences in spray patterns for diverse conditions and fuel injector configurations are readily discernible. An examination of the data has shown that a direct determination of the fuel spray angle at realistic conditions is also possible. The results obtained in this study demonstrate the applicability and usefulness of these nonintrusive optical techniques for investigating fuel spray patternation under actual combustor conditions.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-206292 , NAS 1.15:206292 , E-10977 , JANNAF CS/PSHS/APS Joint Meeting; Oct 27, 1997 - Oct 31, 1997; West Palm Beach, FL; United States
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  • 66
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    Study of Boundary Layer Development in a Two-Stage Low-Pressure Turbine (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Experimental data from jet-engine tests have indicated that unsteady blade row (wake) interactions and separation can have a significant impact on the efficiency of turbine stages. The effects of these interactions can be intensified in low-pressure turbine stages because of the low Reynolds number operating environment. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Thus, during the last decade a significant amount of effort has been put into determining the effects of transition and turbulence on the performance of low pressure turbine stages. Experimental investigations have been performed, for example, by Hodson et al. and Halstead et al. These investigations have helped identify/clarify the roles that factors such as the Reynolds number, free stream turbulence intensity, pressure gradient and curvature have in the generation of losses. In parallel to the experimental investigations, there have been significant analytical efforts to improve the modeling of transition. Examples of such efforts include the works of Mayle and Gostelow et al. These newer models show promise of providing accurate transition predictions over a wide range of flow conditions, although they have yet to be implemented into the numerical flow analyses used by the turbine design community. Some recent computational investigations of interest include the works of Chernobrovkin and Lakshminarayana and Eulitz and Engel. The focus of the current effort has been to -use a viscous, unsteady quasi-three-dimensional Navier-Stokes analysis to study boundary layer development in a two-stage low-pressure turbine. A two-layer algebraic turbulence model, along with a natural transition model and a bubble transition model, have been used, The geometry used in the simulations has been the subject of extensive experiments. The predicted results have been compared with experimental data, including airfoil loadings and time-averaged/unsteady integral boundary layer quantities.
    Keywords: Aircraft Propulsion and Power
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  • 67
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    Optical Fuel Injector Patternation Measurements in Advanced Liquid-Fueled, High Pressure, Gas Turbine Combustors (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Planar laser-induced fluorescence (PLIF) imaging and planar Mie scattering are used to examine the fuel distribution pattern (patternation) for advanced fuel injector concepts in kerosene burning, high pressure gas turbine combustors. Three fuel injector concepts for aerospace applications were investigated under a broad range of operating conditions. Fuel PLIF patternation results are contrasted with those obtained by planar Mie scattering. For one injector, further comparison is also made with data obtained through phase Doppler measurements. Differences in spray patterns for diverse conditions and fuel injector configurations are readily discernible. An examination of the data has shown that a direct determination of the fuel spray angle at realistic conditions is also possible. The results obtained in this study demonstrate the applicability and usefulness of these nonintrusive optical techniques for investigating fuel spray patternation under actual combustor conditions.
    Keywords: Aircraft Propulsion and Power
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    GaAs/Ge Solar Powered Aircraft (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Unmanned Aerial Vehicles (UAV) are being proposed for many applications for many applications including surveillance, mapping and atmospheric studies. These applications require a lightweight, low speed, medium to long duration aircraft. Due to the weight, speed, and altitude constraints imposed on such an aircraft, solar array generated electric power can be a viable alternative to air-breathing engines for certain missions. Development of such an aircraft is currently being funded under the Environmental Research Aircraft and Sensor Technology (ERAST) program. NASA Lewis Research Center (LeRC) has built a Solar Electric Airplane to demonstrate UAV technology. This aircraft utilizes high efficiency Applied Solar Energy Corporation (ASEC) GaAs/Ge space solar cells. The cells have been provided by the Air Force through the ManTech Office.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208652 , NAS 1.15:208652 , E-11371
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  • 69
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    The Numerical Investigation of a Dual-Mode Scramjet Combustor (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A numerical investigation of a multiple-jet array dual-mode scramjet combustor has been performed utilizing a three-dimensional Navier-Stokes code with finite-rate chemistry. Results indicate substantial upstream interaction in the form of an oblique shock/expansion train upstream of the combustor, culminating in completely subsonic flow in the vicinity of fuel injectors. The flow returns to supersonic velocities in the downstream (diverging) portion of the combustor. Mixing and combustion are rapid in this flow and predicted combustion efficiency closely matches experimental data. However, comparisons of wall pressure between the simulation and the experiment show i) substantial underprediction of the upstream interaction distance and ii) moderate overprediction of peak pressure in the vicinity of the entrance of the combustor. This can be at least partially explained by examination of available experimental data; this data shows a very significant movement of the entering vitiated airflow to the sides of the combustor (around the injector array and the upstream interaction front as a whole). This important effect is currently being examined by an extension of the modeling to include the entire half-duct of the same combustor geometry.
    Keywords: Aircraft Propulsion and Power
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  • 70
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    Improved Modeling of Finite-Rate Turbulent Combustion Processes in Research Combustors (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The objective of this thesis is to further develop and test a stochastic model of turbulent combustion in recirculating flows. There is a requirement to increase the accuracy of multi-dimensional combustion predictions. As turbulence affects reaction rates, this interaction must be more accurately evaluated. In this work a more physically correct way of handling the interaction of turbulence on combustion is further developed and tested. As turbulence involves randomness, stochastic modeling is used. Averaged values such as temperature and species concentration are found by integrating the probability density function (pdf) over the range of the scalar. The model in this work does not assume the pdf type, but solves for the evolution of the pdf using the Monte Carlo solution technique. The model is further developed by including a more robust reaction solver, by using accurate thermodynamics and by more accurate transport elements. The stochastic method is used with Semi-Implicit Method for Pressure-Linked Equations. The SIMPLE method is used to solve for velocity, pressure, turbulent kinetic energy and dissipation. The pdf solver solves for temperature and species concentration. Thus, the method is partially familiar to combustor engineers. The method is compared to benchmark experimental data and baseline calculations. The baseline method was tested on isothermal flows, evaporating sprays and combusting sprays. Pdf and baseline predictions were performed for three diffusion flames and one premixed flame. The pdf method predicted lower combustion rates than the baseline method in agreement with the data, except for the premixed flame. The baseline and stochastic predictions bounded the experimental data for the premixed flame. The use of a continuous mixing model or relax to mean mixing model had little effect on the prediction of average temperature. Two grids were used in a hydrogen diffusion flame simulation. Grid density did not effect the predictions except for peak temperature and tangential velocity. The hybrid pdf method did take longer and required more memory, but has a theoretical basis to extend to many reaction steps which cannot be said of current turbulent combustion models.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/TM-1998-208823 , E-11430 , NAS 1.15:208823
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  • 71
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    Parallel 3D Multi-Stage Simulation of a Turbofan Engine (1998)
    In:  Other Sources
    Details
    Publication Date: 2019-07-10
    Description: A 3D multistage simulation of each component of a modern GE Turbofan engine has been made. An axisymmetric view of this engine is presented in the document. This includes a fan, booster rig, high pressure compressor rig, high pressure turbine rig and a low pressure turbine rig. In the near future, all components will be run in a single calculation for a solution of 49 blade rows. The simulation exploits the use of parallel computations by using two levels of parallelism. Each blade row is run in parallel and each blade row grid is decomposed into several domains and run in parallel. 20 processors are used for the 4 blade row analysis. The average passage approach developed by John Adamczyk at NASA Lewis Research Center has been further developed and parallelized. This is APNASA Version A. It is a Navier-Stokes solver using a 4-stage explicit Runge-Kutta time marching scheme with variable time steps and residual smoothing for convergence acceleration. It has an implicit K-E turbulence model which uses an ADI solver to factor the matrix. Between 50 and 100 explicit time steps are solved before a blade row body force is calculated and exchanged with the other blade rows. This outer iteration has been coined a "flip." Efforts have been made to make the solver linearly scaleable with the number of blade rows. Enough flips are run (between 50 and 200) so the solution in the entire machine is not changing. The K-E equations are generally solved every other explicit time step. One of the key requirements in the development of the parallel code was to make the parallel solution exactly (bit for bit) match the serial solution. This has helped isolate many small parallel bugs and guarantee the parallelization was done correctly. The domain decomposition is done only in the axial direction since the number of points axially is much larger than the other two directions. This code uses MPI for message passing. The parallel speed up of the solver portion (no 1/0 or body force calculation) for a grid which has 227 points axially.
    Keywords: Aircraft Propulsion and Power
    Type: High Performance Computing and Communication (HPCC)/Computational AeroSsience (CAS) Workshop; United States
    Format: text
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  • 72
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    Advanced Low Emissions Subsonic Combustor Study (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: Recent advances in commercial and military aircraft gas turbines have yielded significant improvements in fuel efficiency and thrust-to-weight ratio, due in large part to increased combustor operating pressures and temperatures. However, the higher operating conditions have increased the emission of oxides of nitrogen (NOx), which is a pollutant with adverse impact on the atmosphere and environment. Since commercial and military aircraft are the only important direct source of NOx emissions at high altitudes, there is a growing consensus that considerably more stringent limits on NOx emissions will be required in the future for all aircraft. In fact, the regulatory communities have recently agreed to reduce NOx limits by 20 percent from current requirements effective in 1996. Further reductions at low altitude, together with introduction of limits on NOx at altitude, are virtual certainties. In addition, the U.S. Government recently conducted hearings on the introduction of federal fees on the local emission of pollutants from all sources, including aircraft. While no action was taken regarding aircraft in this instance, the threat of future action clearly remains. In these times of intense and growing international competition, the U.S. le-ad in aerospace can only be maintained through a clear technological dominance that leads to a product line of maximum value to the global airline customer. Development of a very low NOx combustor will be essential to meet the future needs of both the commercial and military transport markets, if additional economic burdens and/or operational restrictions are to be avoided. In this report, Pratt & Whitney (P&W) presents the study results with the following specific objectives: Development of low-emissions combustor technologies for advances engines that will enter into service circa 2005, while producing a goal of 70 percent lower NOx emissions, compared to 1996 regulatory levels. Identification of solution approaches to barriers to the productization and economic viability of the low-emissions technologies. Preparation of these technologies to facilitate an annular rig high-pressure demonstration.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-207931 , E-11199 , GRC-E-DAA-TN10123
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  • 73
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    Optical Closed-Loop Propulsion Control System Development (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: The overall objective of this program was to design and fabricate the components required for optical closed-loop control of a F404-400 turbofan engine, by building on the experience of the NASA Fiber Optic Control System Integration (FOCSI) program. Evaluating the performance of fiber optic technology at the component and system levels will result in helping to validate its use on aircraft engines. This report includes descriptions of three test plans. The EOI Acceptance Test is designed to demonstrate satisfactory functionality of the EOI, primarily fail-safe throughput of the F404 sensor signals in the normal mode, and validation, switching, and output of the five analog sensor signals as generated from validated optical sensor inputs, in the optical mode. The EOI System Test is designed to demonstrate acceptable F404 ECU functionality as interfaced with the EOI, making use of a production ECU test stand. The Optical Control Engine Test Request describes planned hardware installation, optical signal calibrations, data system coordination, test procedures, and data signal comparisons for an engine test demonstration of the optical closed-loop control.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-208416 , E-11272 , NAS 1.26:208416 , R98AEB237
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  • 74
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    Turbulent Radiation Effects in HSCT Combustor Rich Zone (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A joint UTRC-University of Connecticut theoretical program was based on describing coupled soot formation and radiation in turbulent flows using stretched flamelet theory. This effort was involved with using the model jet fuel kinetics mechanism to predict soot growth in flamelets at elevated pressure, to incorporate an efficient model for turbulent thermal radiation into a discrete transfer radiation code, and to couple die soot growth, flowfield, and radiation algorithm. The soot calculations used a recently developed opposed jet code which couples the dynamical equations of size-class dependent particle growth with complex chemistry. Several of the tasks represent technical firsts; among these are the prediction of soot from a detailed jet fuel kinetics mechanism, the inclusion of pressure effects in the soot particle growth equations, and the inclusion of the efficient turbulent radiation algorithm in a combustor code.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-1998-206532 , NAS 1.26:206532 , E-11021
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  • 75
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    Active Pattern Factor Control for Gas Turbine Engines (1998)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Small variations in fuel/air mixture ratios within gas turbine combustors can result in measurable, and potentially detrimental, exit thermal gradients. Thermal gradients can increase emissions, as well as shorten the design life of downstream turbomachinery, particularly stator vanes. Uniform temperature profiles are usually sought through careful design and manufacturing of related combustor components. However, small componentto-component variations as well as numerous aging effects degrade system performance. To compensate for degraded thermal performance, researchers are investigating active, closed-loop control schemes.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1997; NASA/TM-1998-206312
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  • 76
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    Investigation of Noise Field and Velocity Profiles of an Afterburning Engine (1954)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Sound pressure levels, frequency spectrum, and jet velocity profiles are presented for an engine-afterburner combination at various values of afterburner fuel - air ratio. At the high fuel-air ratios, severe low-frequency resonance was encountered which represented more than half the total energy in the sound spectrum. At similar thrust conditions, lower sound pressure levels were obtained from a current fighter air craft with a different afterburner configuration. The lower sound pressure levels are attributed to resonance-free afterburner operation and thereby indicate the importance of acoustic considerations in afterburner design.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E54G07
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  • 77
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    An Investigation in the Ames 40- by 80-Foot Wind Tunnel of a YT-56A Turboprop Engine Incorporating a Decoupler and a Controlled-Feathering Device (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation of a decoupler and a controlled-feathering device incorporated with the YT-56A turboprop engine has been made to determine the effectiveness of these devices in reducing the high negative thrust (drag) which accompanies power failure of this type of engine. Power failures were simulated by fuel cut-off, both without either device free to operate, and with each device free to operate singly. The investigation was made through an airspeed range from 50 to 230 mph. It was found that with neither device free to operate, the drag levels realized after power failures at airspeeds above 170 mph would impose vertical tail loads higher than those allowable for the YC-130, the airplane for which the test power package was designed. These levels were reached in approximately one second. The maximum drag realized after power failure was not appreciably altered by the use of the decoupler although the decoupler did put a limit on the duration of the peak drag. The controlled-feathering device maintained a level of essentially zero drag after power failure. The use of the decoupler in the YT-56A engine complicates windmilling air-starting procedures and makes it necessary to place operating restrictions on the engine to assure safe flight at low-power conditions,
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SA54I09
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  • 78
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    Preliminary Altitude Operational Characteristics of a J57-P1 Turbojet Engine (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The operational characteristics of a J57-P1 turbojet engine have been investigated at altitudes between 15,000 and 66,000 feet in the Lewis altitude wind tunnel. Included in this study is a discussion of fuel nozzle coking, the altitude operating limits with and without the standard engine control, the compressor surge characteristics, and the engine starting and windmilling characteristics. Severe circumferential turbine outlet temperature gradients which occurred at high altitude as a result of fuel nozzle coking were alleviated by the manufacturer's change in the fuel flow divider schedule and in a nozzle gasket material. Compressor air bleed is required to prevent surge of the outboard compressor in the low engine speed region. The maximum altitude at which the engine was operated without the control was about 66,000 feet at 0.8 flight Mach number and at a reduced engine speed to avoid compressor surge; with the engine control in operation, the altitude operating limit is reduced to approximately 59,000 feet. The maximum altitude at which the engine was started was about 40,000 feet.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE54C31
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  • 79
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    Design and Evaluation of a Turbojet Exhaust Simulator, Utilizing a Solid-Propellant Rocket Motor, for use in Free-Flight Aerodynamic Research Models (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A method has been developed for modifying a rocket motor so that its exhaust characteristics simulate those of a turbojet engine. The analysis necessary to the design is presented along with tests from which the designs are evaluated. Simulation was found to be best if the exhaust characteristics to be duplicated were those of a turbojet engine at high altitudes and with the afterburner operative.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-L54I15
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  • 80
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    Experimental Permeability Measurements on a Strut-Supported Transpiration-Cooled Turbine Blade with Stainless-Steel Shell made by the Federal-Mogul Corporation under Bureau of Aeronautics Contract N0as 51613-C (1954)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A turbine blade with a porous stainless-steel shell sintered to a supporting steel strut has been fabricated for tests at the NACA by Federal-Mogul Corporation under contract from the Bureau of Aeronautics, Department of the Navy. The apparent permeability of this blade, on the average, more nearly approaches the values specified by the NAGA than did two strut-supported bronze blades in a previous investigation. Random variations of permeability in the present blade are substantialy greater than those of the bronze blades, but projected improvements in certain phases of the fabrication process are expected to reduce these variations.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE54D29
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  • 81
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    Crustal layers of the continents and oceans (1951)
    In:  Bull. Geol. Soc. Am., Milano, California Institute of Technology Pasadena, vol. 62, no. 6, pp. 427-440, pp. L24306, (ISBN: 0534351875, 2nd edition)
    Details
    Publication Date: 1951
    Keywords: CRUST ; Earth model, also for more shallow analyses !
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    BIBLIOGRAPHY SEISMOLOGY
  • 82
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    Strain characteristics of the earth's interior (1951)
    Dover Publ.
    In:  Professional Paper, Internal constitution of the earth - Physic of the earth, Dover, 439 pp., Dover Publ., vol. 7, no. XV:, pp. 382-407, (ISBN: 3-540-23712-7)
    Details
    Publication Date: 1951
    Keywords: Earth model, also for more shallow analyses ! ; Strain ; TIDES ; (The Earth's free) oscillations
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  • 83
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    Colloquium on plastic flow and deformation within the earth (Hershey, Pennsylvania, September 12-14, 1950) (1951)
    In:  Trans., Am. Geophys. Union, Milano, California Institute of Technology Pasadena, vol. 32, no. 6, pp. 497-543, pp. L24306, (ISBN: 0534351875, 2nd edition)
    Details
    Publication Date: 1951
    Keywords: Elasticity ; Rock mechanics ; Rheology ; Fluids ; Earth model, also for more shallow analyses ! ; EOS ; BIBTEX? ; ENDNOTE?
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  • 84
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    Internal constitution of the earth - Physic of the earth (1951)
    Dover Publ.
    In:  Dover, 439 pp., Dover Publ., vol. 7, no. XVI:, pp. 385-389, (ISBN 0-12-305355-2)
    Details
    Publication Date: 1951
    Keywords: Handbook of geophysics ; Earth model, also for more shallow analyses !
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  • 85
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    Hypotheses on the development of the earth (1951)
    Dover Publ.
    In:  Professional Paper, Internal constitution of the earth - Physic of the earth, Dover, 439 pp., Dover Publ., vol. 7, no. IX:, pp. 178-226, (ISBN: 3-540-23712-7)
    Details
    Publication Date: 1951
    Keywords: Earth model, also for more shallow analyses ! ; evolution
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  • 86
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    Forces in the earth (1951)
    Dover Publ.
    In:  Professional Paper, Internal constitution of the earth - Physic of the earth, Dover, 439 pp., Dover Publ., vol. 7, no. VIII:, pp. 167-177, (ISBN: 3-540-23712-7)
    Details
    Publication Date: 1951
    Keywords: Earth model, also for more shallow analyses ! ; Elasticity ; TIDES ; (The Earth's free) oscillations ; Geothermics
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  • 87
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    Preface (1951)
    Dover Publ.
    In:  Professional Paper, Internal constitution of the earth - Physic of the earth, Dover, 439 pp., Dover Publ., vol. 7, no. VIIa, pp. 438-470, (ISBN: 3-540-23712-7)
    Details
    Publication Date: 1951
    Keywords: Review article ; Earth model, also for more shallow analyses !
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  • 88
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    The cooling of the earth and the temperature in its interior (1951)
    Dover Publ.
    In:  Professional Paper, Internal constitution of the earth - Physic of the earth, Dover, 439 pp., Dover Publ., vol. 7, no. VII:, pp. 150-166, (ISBN: 3-540-23712-7)
    Details
    Publication Date: 1951
    Keywords: Geothermics ; Earth model, also for more shallow analyses !
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  • 89
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    Structure of the crust (1951)
    Dover Publ.
    In:  Professional Paper, Internal constitution of the earth - Physic of the earth, Dover, 439 pp., Dover Publ., vol. 7, no. XII:, pp. 314-339, (ISBN: 3-540-23712-7)
    Details
    Publication Date: 1951
    Keywords: CRUST ; Earth model, also for more shallow analyses !
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  • 90
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    Comparison of High-Speed Operating Characteristics of Size 215 Cylindrical-Roller Bearings as Determined in Turbojet Engine and in Laboratory Test Rig (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A comparison of the operating characteristics of 75-millimeter-bore (size 215) cylindrical-roller one-piece inner-race-riding cage-type bearings was made using a laboratory test rig and a turbojet engine. Cooling correlation parameters were determined by means of dimensional analysis, and the generalized results for both the inner- and outer-race bearing operating temperatures are compared for the laboratory test rig and the turbojet engine. Inner- and outer-race cooling-correlation curves were obtained for the turbojet-engine turbine-roller bearing with the same inner- and outer-race correlation parameters and exponents as those determined for the laboratory test-rig bearing. The inner- and outer-race turbine roller-bearing temperatures may be predicted from a single curve, regardless of variations in speed, load, oil flow, oil inlet temperature, oil inlet viscosity, oil-jet diameter or any combination of these parameters. The turbojet-engine turbine-roller-bearing inner-race temperatures were 30 to 60 F greater than the outer-race-maximum temperatures, the exact values depending on the operating condition and oil viscosity; these results are in contrast to the laboratory test-rig results where the inner-race temperatures were less than the outer-race-maximum temperatures. The turbojet-engine turbine-roller bearing, maximum outer-race circumferential temperature variation was approximately 30 F for each of the oils used. The effect of oil viscosity on inner- and outer-race turbojet-engine turbine-roller-bearing temperatures was found to be significant. With the lower viscosity oil (6x10(exp -7) reyns (4.9 centistokes) at 100 F; viscosity index, 83), the inner-race temperature was approximately 30 to 35 F less than with the higher viscosity oil (53x10(exp -7) reyns (42.8 centistokes) at 100 F; viscosity index, 150); whereas the outer-race-maximum temperatures were 12 to 28 F lower with the lower viscosity oil over the DN range investigated.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E51I05
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  • 91
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    Tests of Two-Blade Propellers in the Langley 8-Foot High-Speed Tunnel to Determine the Effect on Propeller Performance of a Modification of Inboard Pitch Distribution (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests of two propellers having two blades and differing only in the inboard pitch distribution were made in the Langley 8-foot highspeed tunnel to determine the effect of inboard pitch distribution on propeller performance. propeller was designed for operation in the reduced velocity region ahead of an NACA cowling; the inboard pitch distribution of the modified propeller was increased for operation at or near free-stream velocities, such as would be obtained in a pusher installation. conditions covering climb, cruise, and high-speed operation. Wake surveys were taken behind the propellers in order to determine the distribution of thrust along the blades and to aid in the analysis of the results. Test results showed that the modified propeller was about 2.5 percent less efficient for a typical climb condition at all altitudes, 2 percent more efficient for one cruise condition, and 5 percent more efficient for high-speed operation. speed condition, the modified propeller showed a 6-percent loss in efficiency due to compressibility; whereas the original propeller showed an 11-percent efficiency loss due to compressiblity. The lower compressibility loss for the modified propeller resulted from the fact that the inboard sections of this propeller could operate at increased thrust loading after compressibility losses had occurred at the outboard sections.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TN-2268
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  • 92
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    Experimental Investigation of Air-Cooled Turbine Blades in Turbojet Engine. 7: Rotor-Blade Fabrication Procedures (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted to determine the cooling effectiveness of a wide variety of air-cooled turbine-blade configurations. The blades, which were tested in the turbine of a - commercial turbojet engine that was modified for this investigation by replacing two of the original blades with air-cooled blades located diametrically opposite each other, are untwisted, have no aerodynamic taper, and have essentially the same external profile. The cooling-passage configuration is different for each blade, however. The fabrication procedures were varied and often unique. The blades were fabricated using methods most suitable for obtaining a small number of blades for use in the cooling investigations and therefore not all the fabrication procedures would be directly applicable to production processes, although some of the ideas and steps might be useful. Blade shells were obtained by both casting and forming. The cast shells were either welded to the blade base or cast integrally with the base. The formed shells were attached to the base by a brazing and two welding methods. Additional surface area was supplied in the coolant passages by the addition of fins or tubes that were S-brazed. to the shell. A number of blades with special leading- and trailing-edge designs that provided added cooling to these areas were fabricated. The cooling effectiveness and purposes of the various blade configurations are discussed briefly.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E51E23 , REPT-2203
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  • 93
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    Effect of Angle of Incidence of Second-Stage Vane Assembly on Third-Stage Compressor-Blade Vibration and Engine Performance (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Strain-gages were used to measure blade vibrations causing failures in the third stage of a production 11-stage axial-flow compressor. After the serious third-stage vibration was detected, a series of investigations were conducted with second-stage vane assemblies of varying angles of incidence. Curves presented herein show the effect of varying the angle of incidence of second-stage vane assembly on third-stage rotor-blade vibration amplitude and engine performance. A minimum vibration amplitude was obtained without greatly affecting the engine performance with a second-stage vane assembly of 9deg. greater angle of incidence than the assembly normally furnished with the engine.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE51F08
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  • 94
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    Altitude Performance of J35-A-17 Turbojet Engine in an Altitude Chamber (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation of the altitude performance characteristics of an Allison J35-A-17 turbojet engines have been conducted in an altitude chamber at the NACA Lewis laboratory. Engine performance was obtained over a range of altitudes from 20,000 to 60,000 feet at a flight Mach number of 0.62 and a range of flight Mach numbers from 0.42 to 1.22 at an altitude of 30,000 feet. The performance of the engine over the range investigated could be generalized up to an altitude of 30,000 feet. Performance of the engine at any flight Mach number in the range investigated can be predicted for those operating condition a t which critical flow exits in the exhaust nozzle with the exception of the variables corrected net thrust, and net-thrust specific fuel consumption.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E50I15
    Format: application/pdf
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  • 95
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    Flight Performance of a Jet Power Plant (1951)
    In:  CASI
    Details
    Publication Date: 2019-08-13
    Description: The performance of a jet power plant consisting of a compressor and a turbine is determined by the characteristic curves of these component parts and is controllable by the characteristics of the compressor and the turbine i n relation t o each other. The normal. output, overload, and throttled load of the Jet power plant are obtained on the basis of assumed straight-line characteristics.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TM-1258
    Format: application/pdf
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  • 96
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    Graphic Analysis of American and British Axial-Flow Turbojet Engine Performance Trends (Current and Future) (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: This report presents a compilation of static sea-level data on existing or designed American and British axial-flow turbojet engines in terms of basic engine parameters such as thrust and air flow. In the data presented, changes in the over-U engine performance with time sre examined as well as the relation of the various engine parameters to each other.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-51K29
    Format: application/pdf
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  • 97
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    Matching Characteristics of J35-A-23 Compressor and Two-stage Turbine (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Component data on the J35-A-23 compressor and two-stage turbine were used to determine the problems in matching the two units for operatio n in a turbojet engine. Possible operating regions were determined an d an equilibrium operating line was also determined for the assumed c onditions of zero flight speed and a jet nozzle area approximately 5. 5 percent greater than the wide-open nozzle area.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-E51H15
    Format: application/pdf
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  • 98
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    Control Performance of General Electric Fuel and Torque Regulator Operating on T31-3 Turbine-Propeller Engine in Sea-Level Test Stand (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A .General Electric fuel and torque regulator was tested in conjunction with a T31-3 turbine-propeller engine in the sea-level static test stand at the NACA Lewis laboratory. The engine and control were operated over the entire speed range: 11,000 rpm, nominal flight idle, to 13,000 rpm, full power. Steady-state and transient data were recorded and are presented with a description of the four control loops being used in the system. Results of this investigation indicated that single-lever control operation was satisfactory under conditions of test. Transient data presented showed that turbine-outlet temperature did overshoot maximum operating value on acceleration but that the time duration of overshoot did not exceed approximately 1 second. This temperature limiting resulted from a control on fuel flow as a function of engine speed. Speed and torque first reached their desired values 0.4 second from the time of change in power-setting lever position. Maximum speed overshoot was 3 percent.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE1H20
    Format: application/pdf
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  • 99
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    Preliminary Transient Performance Data for Afterburner Operation of Westinghouse Electronic Power Regulator on XJ34-WE-32 Turbojet Engine in Altitude Wind Tunnel (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: At the request of the Bureau of Aeronautics, Department of the Navy, an investigation of the Westinghouse XJ34-WE-32 turbojet engine is being conducted in the NACA Lewis altitude wind tunnel to determine the steady-state and transient operating characteristics of the controlled and uncontrolled engine at various altitudes and ram pressure ratios. As part of this program, transient performance data that illustrate the operation of the engine is obtained in the form of oscillographic traces. Similar data for engine operation i n the afterburning range, covering a range of throttle settings from the minimum value giving rated speed (throttle position, 72 degrees) to full afterburning (throttle position, ll0 degrees), is presented herein. These data thus serve to indicate the transient characteristics of the engine when the throttle is advance into, withdrawn from, and moved within the afterburning range in a stepwise manner, as well as the steady-state stability of the engine during afterburning .
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50L29
    Format: application/pdf
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  • 100
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    Factors Affecting the Starting Characteristics of Gas-Turbine Engines (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: This report summarizes the effects of fuel volatility and engine design variables on the problem of starting gas-turbine engines at sea-level and altitude conditions. The starting operation for engines with tubular combustors is considered as three steps; namely, (1) ignition of a fuel-air mixture in the combustor, (2) propagation of flame through cross-fire tubes to all combustors, and (3) acceleration of the engine from windmilling or starting speed to the operating speed range. Pertinent data from laboratory researches, single-combustor studies, and full-scale engine investigations are presented on each phase of the starting problem.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE51B02
    Format: application/pdf
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