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  • Life and Medical Sciences  (629)
  • Aircraft Propulsion and Power
  • 1995-1999  (719)
  • 1955-1959
  • 1935-1939
  • 1996  (719)
Collection
Keywords
Publisher
Years
  • 1995-1999  (719)
  • 1955-1959
  • 1935-1939
Year
  • 1
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    Air Force Seal Programs (1996)
    In:  CASI
    Details
    Publication Date: 2011-09-13
    Description: Seal technology development is an important part of the Air Force's participation in the Integrated High Performance Turbine Engine Technology (IHPTET) initiative, the joint DOD, NASA, ARPA, and industry endeavor to double turbine engine capabilities by the turn of the century. Significant performance and efficiency improvements can be obtained through reducing internal flow system leakage, but seal environment requirements continue to become more extreme as the engine thermodynamic cycles advance towards these IHPTET goals. Seal technology continues to be pursued by the Air Force to control leakage at the required conditions. This presentation briefly describes current seal research and development programs and gives a summary of seal applications in demonstrator and developmental engines.
    Keywords: Aircraft Propulsion and Power
    Type: Seals Code Development Workshop; 73-80; NASA-CP-10181
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    NASA TECHNICAL REPORTS
  • 2
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    Turbomachine Sealing (1996)
    In:  CASI
    Details
    Publication Date: 2011-09-13
    Description: Designers and customers are demanding higher performance turbomachine systems that have long life between overhauls and satisfy the more restrictive environmental constraints. This overview provides sources of design data, numerical, and experimental results along with selected new seal configurations and static sealing challenges such as in the combustors. The following categories are presented: (1) Seal Rotordynamic Data Base (experimental analytical program at Texas A&M); (2) Secondary Flow Interactions (validation studies at CFDRC, Huntsville AL); (3) Contact Sealing (selected types with finger seal model); and (4) Environmental Constraints (emphasis on combustors).
    Keywords: Aircraft Propulsion and Power
    Type: Seals Code Development Workshop; 5-40; NASA-CP-10181
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  • 3
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    Aspirating Seal Development: Analytical Modeling and Seal Test Rig (1996)
    In:  CASI
    Details
    Publication Date: 2011-09-13
    Description: This effort is to develop large diameter (22 - 36 inch) Aspirating Seals for application in aircraft engines. Stein Seal Co. will be fabricating the 36-inch seal(s) for testing. GE's task is to establish a thorough understanding of the operation of Aspirating Seals through analytical modeling and full-scale testing. The two primary objectives of this project are to develop the analytical models of the aspirating seal system, to upgrade using GE's funds, GE's 50-inch seal test rig for testing the Aspirating Seal (back-to-back with a corresponding brush seal), test the aspirating seal(s) for seal closure, tracking and maneuver transients (tilt) at operating pressures and temperatures, and validate the analytical model. The objective of the analytical model development is to evaluate the transient and steady-state dynamic performance characteristics of the seal designed by Stein. The transient dynamic model uses a multi-body system approach: the Stator, Seal face and the rotor are treated as individual bodies with relative degrees of freedom. Initially, the thirty-six springs are represented as a single one trying to keep open the aspirating face. Stops (Contact elements) are provided between the stator and the seal (to compensate the preload in the fully-open position) and between the rotor face and Seal face (to detect rub). The secondary seal is considered as part of the stator. The film's load, damping and stiffness characteristics as functions of pressure and clearance are evaluated using a separate (NASA) code GFACE. Initially, a laminar flow theory is used. Special two-dimensional interpolation routines are written to establish exact film load and damping values at each integration time step. Additionally, other user-routines are written to read-in actual pressure, rpm, stator-growth and rotor growth data and, later, to transfer these as appropriate loads/motions in the system-dynamic model. The transient dynamic model evaluates the various motions, clearances and forces as the seals are subjected to different aircraft maneuvers: Windmilling restart; start-ground idle; ground idle-takeoff; takeoff-burst chop, etc. Results of this model show that the seal closes appropriately and does not ram into the rotor for all of the conditions analyzed. The rig upgrade design for testing Aspirating Seals has been completed. Long lead-time items (forgings, etc.) have been ordered.
    Keywords: Aircraft Propulsion and Power
    Type: Seals Code Development Workshop; 89-114; NASA-CP-10181
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  • 4
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    Innovative Sealing Technology for Next Generation Subsonic Engines (1996)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: The next generation of subsonic engines can be expected to continue the historical trend towards increased thrust to weight (T/W) and decreased specific fuel consumption (SFC). Development programs currently underway throughout the gas turbine industry such as DOD's Integrated High Performance Turbine Engine Technology (IHPTET), and more recently NASA's Advanced Subsonic Transport (AST) programs, have altered these trends in both pace and magnitude. Advanced seals and sealing technologies have become a prominent part of these efforts due to the large potential performance gains which can be realized. Allison has recently completed a study for NASA the goal of which was to quantize the potential performance benefits which might accrue through the use of advanced seals in future subsonic gas turbine engines. For the study, two engines where analyzed, a small turboshaft and a larger turbofan engine to help assess the effect of engine size on the results. Engines were analyzed stage by stage with the most sensitive areas highlighted. Leakage characteristics for advanced seals were then substituted into secondary airflow models, and the leakage reductions documented. These leakage reductions were then converted to changes in performance, i.e. increased range, decreased takeoff gross weight, etc. and presented. It was found that the development and use of a realtively few advanced seals, less than 5, could for example reduce SFC by 10% or more.
    Keywords: Aircraft Propulsion and Power
    Type: Seals Code Development Workshop; 327-336; NASA-CP-10181
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  • 5
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    Aerospace Seal Requirements (1996)
    In:  CASI
    Details
    Publication Date: 2004-12-03
    Description: Cycle studies have shown the benefits of increasing engine pressure ratios and cycle temperatures to decrease engine weight and improve performance of commercial turbine engines. NASA is working with industry to define technology requirements of advanced engines and engine technology to meet the goals of NASA's Advanced Subsonic Technology Initiative. As engine operating conditions become more severe and customers demand lower operating costs, NASA and engine manufacturers are investigating methods of improving engine efficiency and reducing operating costs. A number of new technologies are being examined that will allow next generations engines to operate at higher pressures and temperatures. Improving seal performance - reducing leakage and increasing service life while operating under more demanding conditions - will play an important role in meeting overall program goals of reducing specific fuel consumption and ultimately reducing direct operating costs. This paper provides an overview of the Advanced Subsonic Technology Program goals discusses the motivation for advanced seal development, and highlights seal technology requirements to meet future engine performance goals.
    Keywords: Aircraft Propulsion and Power
    Type: Seals Code Development Workshop; 41-54; NASA-CP-10181
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  • 6
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    Inlet and Propulsion Integration of Scram Propelled Vehicles (1996)
    In:  CASI
    Details
    Publication Date: 2011-10-14
    Description: The material to be presented in these two lectures begins with cycle considerations of the turbojet engine combined with a ramjet engine to provide thrust over the range of Mach 0 to 5. We will then examine in some detail the aerodynamic behavior that occurs in the inlet operating near the peak speed. Following that, we shall view a numerical simulation through a baseline scramjet engine, starting at the entrance to the inlet, proceeding into the combustor and through the nozzle. In the next segment, we examine a combined rocket and ramjet propulsion system. Analysis and test results will be examined with a view toward evaluation of the concept as a practical device. Two other inlets will then be reviewed: a Mach 12 inlet and a Mach 18 configuration. Finally, we close our lectures with a discussion of the Detonation Wave engine, and inspect the physical and chemical behavior obtained from numerical simulation. A few final remarks will be made regarding the application of CFD for hypersonic propulsion components.
    Keywords: Aircraft Propulsion and Power
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  • 7
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    Integrated Flight and Propulsion Controls for Advanced Aircraft Configurations (1996)
    In:  CASI
    Details
    Publication Date: 2011-10-14
    Description: The research vision of the NASA Lewis Research Center in the area of integrated flight and propulsion controls technologies is described. In particular, the integrated method for propulsion and airframe controls developed at the Lewis Research Center is described including its application to an advanced aircraft configuration. Additionally, future research directions in integrated controls are described.
    Keywords: Aircraft Propulsion and Power
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  • 8
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    Comparisons Between Unsteady Aerodynamic Events in a Gas Turbine Generator and an Identical Compressor Rig (1996)
    In:  CASI
    Details
    Publication Date: 2011-10-14
    Description: Extensive testing done on a T55-L-712 turboshaft engine compressor in a compressor test rig is being followed by engine tests in progress as part of the Army Non-Recoverable Stall Program. Goals include a greater understanding of the gas turbine engine start cycle and compressor/engine operation in the regions 'beyond' the normal compressor stall line (rotating stall/surge). Rig steady state instrumentation consisted of 497 steady state pressure sensors and 153 temperature sensors. Engine instrumentation was placed in similar radial/axial locations and consists of 122 steady state pressure sensors and 65 temperature sensors. High response rig instrumentation consisted of 34 wall static pressure transducers. Rig and engine high response pressure transducers were located in the same axial/radial/circumferential locations in front of the first three stages. Additional engine high response instrumentation was placed in mach probes in front of the engine and on the compressor hub. This instrumentation allows for the generation of detailed stage characteristics, overall compressor mapping, and detailed analysis of dynamic compressor events.
    Keywords: Aircraft Propulsion and Power
    Type: Loss Mechanisms and Unsteady Flows in Turbomachines; AGARD-CP-571
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  • 9
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    A Pre-Mixed Shock-Induced-Combustion Approach to Inlet and Combustor Design for Hypersonic Applications (1996)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The need for efficient access to space has created interest in airbreathing propulsion as a means of achieving that goal. The NASP program explored a single-stage-to-orbit approach which could require scramjet airbreathing propulsion out to Mach 16 to 20. Recent interest in global access could require hypersonic cruise engines operating efficiently in the Mach 10 to 12 speed range. A common requirement of both these types of propulsion systems is that they would have to be fully integrated with the aero configuration so that the forebody becomes a part of the external compression inlet and the nozzle expansion is completed on the vehicle aftbody.
    Keywords: Aircraft Propulsion and Power
    Type: Transportation Beyond 2000: Technologies Needed for Engineering Design; 639-652; NASA-CP-10184-Pt-2
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  • 10
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    A Giant Step in Jetliner Propulsion (1996)
    In:  CASI
    Details
    Publication Date: 2018-06-09
    Description: GE Aircraft's GE90 is a high bypass turbofan jetliner engine capable of well over 84,700 pounds thrust. The turbofan is a propulsion system that compresses some of the air taken in, burns it in a combuster and expells it to generate power for driving the fan and compressor. A greater amount of air bypasses the combustion process. The GE90 pushes the cooler bypass air rearward with a fan to mix it with the hot exhaust gas; the result is a gain in thrust with minimal fuel expenditure. Over a billion dollars and several years went into its development, which included incorporating technologies developed by Lewis Research Center work done in the 1970s and from projects with SNECMA of France. The engine will power the 777 and other subsonic commercial widebodies.
    Keywords: Aircraft Propulsion and Power
    Type: Spinoff 1996; 56-57; NASA/NP-1996-10-222-HQ
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  • 11
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    LDV Measurements in an Annular Combustor Model (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This thesis covers the design and setup of a laser doppler velocimeter (LDV) system used to take velocity measurements in an annular combustor model. The annular combustor model is of contemporary design using 60 degree flat vane swirlers, producing a strong recirculation zone. Detailed measurements are taken of the swirler inlet air flow and of the downstream enclosed swirling flow. The laser system used is a two color, two component system set up in forward scatter. Detailed are some of the special considerations needed for LDV use in the confined turbulent flow of the combustor model. LDV measurements in a single swirler rig indicated that the flow changes radically in the first duct height. After this, a flow profile is set up and remains constant in shape. The magnitude of the velocities gradually decays due to viscous damping.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-182207 , NAS 1.26:182207 , E-9865
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  • 12
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    Reynolds-Averaged Navier-Stokes Studies of Low Reynolds Number Effects on the Losses in a Low Pressure Turbine (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experimental data from jet-engine tests have indicated that unsteady blade-row interaction effects 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. Preliminary studies indicate that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study, numerical experiments have been performed to quantify the Reynolds number dependence of unsteady wake/separation bubble interaction on the performance of a low-pressure turbine.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198534 , NAS 1.26:198534 , E-10457
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  • 13
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    CFD Study of Turbo-Ramjet Interactions in Hypersonic Airbreathing Propulsion System (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Advanced airbreathing propulsion systems used in Mach 4-6 mission scenarios, usually involve turbo-ramjet configurations. As the engines transition from turbojet to ramjet, there is an operational envelope where both engines operate simultaneously. In the first phase of our study, an over/under nozzle configuration was analyzed. The two plumes from the turbojet and ramjet interact at the end of a common 2-D cowl, where they both reach an approximate Mach 3.0 condition and then jointly expand to Mach 3.6 at the common nozzle exit plane. For the problem analyzed, the turbojet engine operates at a higher nozzle pressure ratio than the ramjet, causes the turbojet plume overpowers the ramjet plume, deflecting it approximately 12 degrees downward and in turn the turbojet plume is deflected 6 degrees upward. In the process, shocks were formed at the deflections and a shear layer formed at the confluence of the two jets. This particular case was experimentally tested and the data were used to compare with a computational fluid dynamics (CFD) study using the PARC2D code. The CFD results were in good agreement with both static pressure distributions on the cowl separator and on nozzle walls. The thrust coefficients were also in reasonable agreement. In addition, inviscid relationships were developed around the confluence point, where the two exhaust jets meet, and these results compared favorably with the CFD results. In the second phase of our study, a 3-D CFD solution was generated to compare with the 2-D solution. The major difference between the 2-D and 3-D solutions was the interaction of the shock waves, generated by the plume interactions, on the sidewall. When a shock wave interacts with a sidewall and sidewall boundary layer, it is called a glancing shock sidewall interaction. These interactions entrain boundary layer flow down the shockline into a vortical flow pattern. The 3-D plots show the streamlines being entrained down the shockline. The pressure of the flow also decreases slightly as the sidewall is approached. Other difference between the 2-D and 3-D solutions were a lowering of the nozzle thrust coefficient value from 0.9850 (2-D) to 0.9807 (3-D), where the experimental value was 0.9790. In the third phase of our study, a different turbo-ramjet configuration was analyzed. The confluence of a supersonic turbojet and a subsonic ramjet in the turbine based combined-cycle (TBCC) propulsion system was studied by a 2-D CFD code. In the analysis, Mach 1.4 primary turbojet was mixed with the subsonic ramjet secondary flow in an ejector mode operation. Reasonable agreements were obtained with the supplied I-D TBCC solutions. For low downstream backpressure, the Fabri choke condition (Break-Point condition) was observed in the secondary flow within mixing zone. For sufficient high downstream backpressure, the Fabri choke no longer exist, the ramjet flow was reduced and the ejector flow became backpressure dependent. Highly non-uniform flow at ejector exit were observed, indicated that for smooth downstream combustion, the mixing of the two streams probably required some physical devices.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-202418 , NAS 1.26:202418
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  • 14
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    Improved NASA-ANOPP Noise Prediction Computer Code for Advanced Subsonic Propulsion Systems (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Recent experience using ANOPP to predict turbofan engine flyover noise suggests that it over-predicts overall EPNL by a significant amount. An improvement in this prediction method is desired for system optimization and assessment studies of advanced UHB engines. An assessment of the ANOPP fan inlet, fan exhaust, jet, combustor, and turbine noise prediction methods is made using static engine component noise data from the CF6-8OC2, E(3), and QCSEE turbofan engines. It is shown that the ANOPP prediction results are generally higher than the measured GE data, and that the inlet noise prediction method (Heidmann method) is the most significant source of this overprediction. Fan noise spectral comparisons show that improvements to the fan tone, broadband, and combination tone noise models are required to yield results that more closely simulate the GE data. Suggested changes that yield improved fan noise predictions but preserve the Heidmann model structure are identified and described. These changes are based on the sets of engine data mentioned, as well as some CFM56 engine data that was used to expand the combination tone noise database. It should be noted that the recommended changes are based on an analysis of engines that are limited to single stage fans with design tip relative Mach numbers greater than one.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-195480 , NAS 1.26:195480 , E-9710
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  • 15
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    A Gas Chromatograph/Mass Spectrometer System for UltraLow-Emission Combustor Exhaust Studies (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A gas chromatograph (GC)/mass spectrometer (MS) system that allows the speciation of unburnt hydrocarbons in the combustor exhaust has been developed at the NASA Lewis Research Center. Combustion gas samples are withdrawn through a water-cooled sampling probe which, when not in use, is protected from contamination by a high-pressure nitrogen purge. The sample line and its connecting lines, filters, and valves are all ultraclean and are heated to avoid condensation. The system has resolution to the parts-per-billion (ppb) level.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107253 , NAS 1.15:107253 , ARL-MR-293 , E-10152
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  • 16
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    SEADYN Analysis of a Tow Line for a High Altitude Towed Glider (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The concept of using a system, consisting of a tow aircraft, glider and tow line, which would enable subsonic flight at altitudes above 24 km (78 kft) has previously been investigated. The preliminary results from these studies seem encouraging. Under certain conditions these studies indicate the concept is feasible. However, the previous studies did not accurately take into account the forces acting on the tow line. Therefore in order to investigate the concept further a more detailed analysis was needed. The code that was selected was the SEADYN cable dynamics computer program which was developed at the Naval Facilities Engineering Service Center. The program is a finite element based structural analysis code that was developed over a period of 10 years. The results have been validated by the Navy in both laboratory and at actual sea conditions. This code was used to simulate arbitrarily-configured cable structures subjected to excitations encountered in real-world operations. The Navy's interest was mainly for modeling underwater tow lines, however the code is also usable for tow lines in air when the change in fluid properties is taken into account. For underwater applications the fluid properties are basically constant over the length of the tow line. For the tow aircraft/glider application the change in fluid properties is considerable along the length of the tow line. Therefore the code had to be modified in order to take into account the variation in atmospheric properties that would be encountered in this application. This modification consisted of adding a variable density to the fluid based on the altitude of the node being calculated. This change in the way the code handled the fluid density had no effect on the method of calculation or any other factor related to the codes validation.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-202308 , NAS 1.26:202308 , E-10589
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  • 17
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    The Effect of Wake Passing on Turbine Blade Film Cooling (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The effect of upstream blade row wake passing on the showerhead film cooling performance of a downstream turbine blade has been investigated through a combination of experimental and computational studies. The experiments were performed in a steady-flow annular turbine cascade facility equipped with an upstream rotating row of cylindrical rods to produce a periodic wake field similar to that found in an actual turbine. Spanwise, chordwise, and temporal resolution of the blade surface temperature were achieved through the use of an array of nickel thin-film surface gauges covering one unit cell of showerhead film hole pattern. Film effectiveness and Nusselt number values were determined for a test matrix of various injectants, injectant blowing ratios, and wake Strouhal numbers. Results indicated a demonstratable reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated well by a chordwise-constant decrement of 0.094-St. Measurable spanwise film effectiveness variations were found near the showerhead region, but meaningful unsteady variations and downstream spanwise variations were not found. Nusselt numbers were less sensitive to wake and injection changes. Computations were performed using a three-dimensional turbulent Navier-Stokes code which was modified to model wake passing and film cooling. Unsteady computations were found to agree well with steady computations provided the proper time-average blowing ratio and pressure/suction surface flow split are matched. The remaining differences were isolated to be due to the enhanced mixing in the unsteady solution caused by the wake sweeping normally on the pressure surface. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107380 , NAS 1.15:107380 , E-10568
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  • 18
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    Atomization and Dispersion of a Liquid Jet Injected Into a Crossflow of Air (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: In recent years, environmental regulations have become more stringent, requiring lower emissions of mainly nitrogen oxides (NOx), as well as carbon monoxide (CO) and unburned hydrocarbons (UHC). These regulations have forced the gas turbine industry to examine non-conventional combustion strategies, such as the lean burn approach. The reasoning behind operating under lean conditions is to maintain the temperature of combustion near and below temperatures required for the formation of thermal nitric oxide (NO). To be successful, however, the lean processes require careful preparation of the fuel/air mixture to preclude formation of either locally rich reaction zones, which may give rise to NO formation, or locally lean reaction zones, which may give rise to inefficient fuel processing. As a result fuel preparation is crucial to the development and success of new aeroengine combustor technologies. A key element of the fuel preparation process is the fuel nozzle. As nozzle technologies have developed, airblast atomization has been adopted for both industrial and aircraft gas turbine applications. However, the majority of the work to date has focused on prefilming nozzles, which despite their complexity and high cost have become an industry standard for conventional combustion strategies. It is likely that the new strategies required to meet future emissions goals will utilize novel fuel injector approaches, such as radial injection. This thesis proposes and demonstrates an experiment to examine, on a mechanistic level (i.e., the physics of the action), the processes associated with the atomization, evaporation, and dispersion of a liquid jet introduced, from a radial, plain-jet airblast injector, into a crossflow of air. This understanding requires the knowledge not only of what factors influence atomization, but also the underlying mechanism associated with liquid breakup and dispersion. The experimental data acquired identify conditions and geometries for improved performance of radial airblast injectors.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198543 , NAS 1.26:198543 , E-10507 , UCI-ARTR-95-4
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  • 19
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    Quick-Mixing Studies Under Reacting Conditions (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The low-NO(x) emitting potential of rich-burn/quick-mix/lean-burn )RQL) combustion makes it an attractive option for engines of future stratospheric aircraft. Because NO(x) formation is exponentially dependent on temperature, the success of the RQL combustor depends on minimizing high temperature stoichiometric pocket formation in the quick-mixing section. An experiment was designed and built, and tests were performed to characterize reaction and mixing properties of jets issuing from round orifices into a hot, fuel-rich crossflow confined in a cylindrical duct. The reactor operates on propane and presents a uniform, non-swirling mixture to the mixing modules. Modules consisting of round orifice configurations of 8, 9, 10, 12, 14, and 18 holes were evaluated at a momentum-flux ratio of 57 and jet-to-mainstream mass-flaw ratio of 2.5. Temperatures and concentrations of O2, CO2, CO, HC, and NO(x) were obtained upstream, down-stream, and within the orifice plane to determine jet penetration as well as reaction processes. Jet penetration was a function of the number of orifices and affected the mixing in the reacting system. Of the six configurations tested, the 14-hole module produced jet penetration close to the module half-radius and yielded the best mixing and most complete combustion at a plane one duct diameter from the orifice leading edge. The results reveal that substantial reaction and heat release occur in the jet mixing zone when the entering effluent is hot and rich, and that the experiment as designed will serve to explore satisfactorily jet mixing behavior under realistic reacting conditions in future studies.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-195375 , NAS 1.26:195375 , E-9074
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    NASA TECHNICAL REPORTS
  • 20
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    Dynamic Inlet Distortion Prediction with a Combined Computational Fluid Dynamics and Distortion Synthesis Approach (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A procedure has been developed for predicting peak dynamic inlet distortion. This procedure combines Computational Fluid Dynamics (CFD) and distortion synthesis analysis to obtain a prediction of peak dynamic distortion intensity and the associated instantaneous total pressure pattern. A prediction of the steady state total pressure pattern at the Aerodynamic Interface Plane is first obtained using an appropriate CFD flow solver. A corresponding inlet turbulence pattern is obtained from the CFD solution via a correlation linking root mean square (RMS) inlet turbulence to a formulation of several CFD parameters representative of flow turbulence intensity. This correlation was derived using flight data obtained from the NASA High Alpha Research Vehicle flight test program and several CFD solutions at conditions matching the flight test data. A distortion synthesis analysis is then performed on the predicted steady state total pressure and RMS turbulence patterns to yield a predicted value of dynamic distortion intensity and the associated instantaneous total pressure pattern.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198053 , NAS 1.26:198053 , H-2129
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    NASA TECHNICAL REPORTS
  • 21
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    Wave rotor demonstrator engine assessment (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The objective of the program was to determine a wave rotor demonstrator engine concept using the Allison 250 series engine. The results of the NASA LERC wave rotor effort were used as a basis for the wave rotor design. A wave rotor topped gas turbine engine was identified which incorporates five basic requirements of a successful demonstrator engine. Predicted performance maps of the wave rotor cycle were used along with maps of existing gas turbine hardware in a design point study. The effects of wave rotor topping on the engine cycle and the subsequent need to rematch compressor and turbine sections in the topped engine were addressed. Comparison of performance of the resulting engine is made on the basis of wave rotor topped engine versus an appropriate baseline engine using common shaft compressor hardware. The topped engine design clearly demonstrates an impressive improvement in shaft horsepower (+11.4%) and SFC (-22%). Off design part power engine performance for the wave rotor topped engine was similarly improved including that at engine idle conditions. Operation of the engine at off design was closely examined with wave rotor operation at less than design burner outlet temperatures and rotor speeds. Challenges identified in the development of a demonstrator engine are discussed. A preliminary design was made of the demonstrator engine including wave rotor to engine transition ducts. Program cost and schedule for a wave rotor demonstrator engine fabrication and test program were developed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198496 , NAS 1.26:198496 , E-10307
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    Dynamic and Transient Performance of Turbofan/Turboshaft Convertible Engine With Variable Inlet Guide Vanes (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A convertible engine called the CEST TF34, using the variable inlet guide vane method of power change, was tested on an outdoor stand at the NASA Lewis Research Center with a waterbrake dynamometer for the shaft load. A new digital electronic system, in conjunction with a modified standard TF34 hydromechanical fuel control, kept engine operation stable and safely within limits. All planned testing was completed successfully. Steady-state performance and acoustic characteristics were reported previously and are referenced. This report presents results of transient and dynamic tests. The transient tests measured engine response to several rapid changes in thrust and torque commands at constant fan (shaft) speed. Limited results from dynamic tests using the pseudorandom binary noise technique are also presented. Performance of the waterbrake dynamometer is discussed in an appendix.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-4696 , E-9637 , NAS 1.15:4696
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  • 23
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    Flight and Static Exhaust Flow Properties of an F110-GE-129 Engine in an F-16XL Airplane During Acoustic Tests (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The exhaust flow properties (mass flow, pressure, temperature, velocity, and Mach number) of the F110-GE-129 engine in an F-16XL airplane were determined from a series of flight tests flown at NASA Dryden Flight Research Center, Edwards, California. These tests were performed in conjunction with NASA Langley Research Center, Hampton, Virginia (LARC) as part of a study to investigate the acoustic characteristics of jet engines operating at high nozzle pressure conditions. The range of interest for both objectives was from Mach 0.3 to Mach 0.9. NASA Dryden flew the airplane and acquired and analyzed the engine data to determine the exhaust characteristics. NASA Langley collected the flyover acoustic measurements and correlated these results with their current predictive codes. This paper describes the airplane, tests, and methods used to determine the exhaust flow properties and presents the exhaust flow properties. No acoustics results are presented.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-104326 , H-2122 , NAS 1.15:104326
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  • 24
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    Multi-Dimensional Measurements of Combustion Species in Flame Tube and Sector Gas Turbine Combustors (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The higher temperature and pressure cycles of future aviation gas turbine combustors challenge designers to produce combustors that minimize their environmental impact while maintaining high operation efficiency. The development of low emissions combustors includes the reduction of unburned hydrocarbons, smoke, and particulates, as well as the reduction of oxides of nitrogen (NO(x)). In order to better understand and control the mechanisms that produce emissions, tools are needed to aid the development of combustor hardware. Current methods of measuring species within gas turbine combustors use extractive sampling of combustion gases to determine major species concentrations and to infer the bulk flame temperature. These methods cannot be used to measure unstable combustion products and have poor spatial and temporal resolution. The intrusive nature of gas sampling may also disturb the flow structure within a combustor. Planar laser-induced fluorescence (PLIF) is an optical technique for the measurement of combustion species. In addition to its non-intrusive nature, PLIF offers these advantages over gas sampling: high spatial resolution, high temporal resolution, the ability to measure unstable species, and the potential to measure combustion temperature. This thesis considers PLIF for in-situ visualization of combustion species as a tool for the design and evaluation of gas turbine combustor subcomponents. This work constitutes the first application of PLIF to the severe environment found in liquid-fueled, aviation gas turbine combustors. Technical and applied challenges are discussed. PLIF of OH was used to observe the flame structure within the post flame zone of a flame tube combustor, and within the flame zone of a sector combustor, for a variety of fuel injector configurations. OH was selected for measurement because it is a major combustion intermediate, playing a key role in the chemistry of combustion, and because its presence within the flame zone can serve as a qualitative marker of flame temperature. All images were taken in the environment of actual engines during flight, using actual jet fuel. The results of the PLIF study led directly to the modification of a fuel injector.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107329 , NAS 1.15:107329 , E-10215
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    Development of an Intelligent Monitoring and Control System for a Heterogeneous Numerical Propulsion System Simulation (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The NASA Numerical Propulsion System Simulation (NPSS) project is exploring the use of computer simulation to facilitate the design of new jet engines. Several key issues raised in this research are being examined in an NPSS-related research project: zooming, monitoring and control, and support for heterogeneity. The design of a simulation executive that addresses each of these issues is described. In this work, the strategy of zooming, which allows codes that model at different levels of fidelity to be integrated within a single simulation, is applied to the fan component of a turbofan propulsion system. A prototype monitoring and control system has been designed for this simulation to support experimentation with expert system techniques for active control of the simulation. An interconnection system provides a transparent means of connecting the heterogeneous systems that comprise the prototype.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-202435 , NAS 1.26:202435
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  • 26
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    Heat Transfer Experiments in the Internal Cooling Passages of a Cooled Radial Turbine Rotor (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental study was conducted (1) to experimentally measure, assess and analyze the heat transfer within the internal cooling configuration of a radial turbine rotor blade and (2) to obtain heat transfer data to evaluate and improve computational fluid dynamics (CFD) procedures and turbulent transport models of internal coolant flows. A 1.15 times scale model of the coolant passages within the NASA LERC High Temperature Radial Turbine was designed, fabricated of Lucite and instrumented for transient beat transfer tests using thin film surface thermocouples and liquid crystals to indicate temperatures. Transient heat transfer tests were conducted for Reynolds numbers of one-fourth, one-half, and equal to the operating Reynolds number for the NASA Turbine. Tests were conducted for stationary and rotating conditions with rotation numbers in the range occurring in the NASA Turbine. Results from the experiments showed the heat transfer characteristics within the coolant passage were affected by rotation. In general, the heat transfer increased and decreased on the sides of the straight radial passages with rotation as previously reported from NASA-HOST-sponsored experiments. The heat transfer in the tri-passage axial flow region adjacent to the blade exit was relatively unaffected by rotation. However, the heat transfer on one surface, in the transitional region between the radial inflow passage and axial, constant radius passages, decreased to approximately 20 percent of the values without rotation. Comparisons with previous 3-D numerical studies indicated regions where the heat transfer characteristics agreed and disagreed with the present experiment.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198492 , E-10155 , NAS 1.26:198492
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    Experimental Reacting Hydrogen Shear Layer Data at High Subsonic Mach Number (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The flow in a planar shear layer of hydrogen reacting with hot air was measured with a two-component laser Doppler velocimeter (LDV) system, a schlieren system, and OH fluorescence imaging. It was compared with a similar air-to-air case without combustion. The high-speed stream's flow speed was about 390 m/s, or Mach 0.71, and the flow speed ratio was 0.34. The results showed that a shear layer with reaction grows faster than one without; both cases are within the range of data scatter presented by the established data base. The coupling between the streamwise and the cross-stream turbulence components inside the shear layers was low, and reaction only increased it slightly. However, the shear layer shifted laterally into the lower speed fuel stream, and a more organized pattern of Reynolds stress was present in the reaction shear layer, likely as a result of the formation of a larger scale structure associated with shear layer corrugation from heat release. Dynamic pressure measurements suggest that coherent flow perturbations existed inside the shear layer and that this flow became more chaotic as the flow advected downstream. Velocity and thermal variable values are listed in this report for a computational fluid dynamics (CFD) benchmark.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TP-3342 , NAS 1.60:3342 , E-7693
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    High-speed engine/component performance assessment using exergy and thrust-based methods (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This investigation summarizes a comparative study of two high-speed engine performance assessment techniques based on energy (available work) and thrust-potential (thrust availability). Simple flow-fields utilizing Rayleigh heat addition and one-dimensional flow with friction are used to demonstrate the fundamental inability of conventional energy techniques to predict engine component performance, aid in component design, or accurately assess flow losses. The use of the thrust-based method on these same examples demonstrates its ability to yield useful information in all these categories. Energy and thrust are related and discussed from the stand-point of their fundamental thermodynamic and fluid dynamic definitions in order to explain the differences in information obtained using the two methods. The conventional definition of energy is shown to include work which is inherently unavailable to an aerospace Brayton engine. An engine-based energy is then developed which accurately accounts for this inherently unavailable work; performance parameters based on this quantity are then shown to yield design and loss information equivalent to the thrust-based method.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198271 , NAS 1.26:198271
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  • 29
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    Active Control of Fan Noise: Feasibility Study (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The objective of this effort is to develop an analytical model for the coupling of active noise control (ANC) piston-type actuators that are mounted flush to the inner and outer walls of an annular duct to the modes in the duct generated by the actuator motion. The analysis will be used to couple the ANC actuators to the modal analysis propagation computer program for the annular duct, to predict the effects of active suppression of fan-generated engine noise sources. This combined program will then be available to assist in the design or evaluation of ANC systems in fan engine annular exhaust ducts. An analysis has been developed to predict the modes generated in an annular duct due to the coupling of flush-mounted ring actuators on the inner and outer walls of the duct. The analysis has been combined with a previous analysis for the coupling of modes to a cylindrical duct in a FORTRAN computer program to perform the computations. The method includes the effects of uniform mean flow in the duct. The program can be used for design or evaluation purposes for active noise control hardware for turbofan engines. Predictions for some sample cases modeled after the geometry of the NASA Lewis ANC Fan indicate very efficient coupling in both the inlet and exhaust ducts for the m = 6 spinning mode at frequencies where only a single radial mode is cut-on. Radial mode content in higher order cut-off modes at the source plane and the required actuator displacement amplitude to achieve 110 dB SPL levels in the desired mode were predicted. Equivalent cases with and without flow were examined for the cylindrical and annular geometry, and little difference was found for a duct flow Mach number of 0.1. The actuator ring coupling program will be adapted as a subroutine to the cylindrical duct modal analysis and the exhaust duct modal analysis. This will allow the fan source to be defined in terms of characteristic modes at the fan source plane and predict the propagation to the arbitrarily-located ANC source plane. The actuator velocities can then be determined to generate the anti-phase mode. The resulting combined fan source/ANC pressure can then be calculated at any desired wall sensor position. The actuator velocities can be determined manually or using a simulation of a control system feedback loop. This will provide a very useful ANC system design and evaluation tool.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198514 , NAS 1.26:198514 , E-10380
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    Active Control of Fan Noise: Feasibility Study (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An extension of a prior study has been completed to examine the potential reduction of aircraft flyover noise by the method of active noise control (ANC). It is assumed that the ANC system will be designed such that it cancels discrete tones radiating from the engine fan inlet or fan exhaust duct, at least to the extent that they no longer protrude above the surrounding broadband noise levels. Thus, without considering the engineering details of the ANC system design, tone levels am arbitrarily removed from the engine component noise spectrum and the flyover noise EPNL levels are compared with and without the presence of tones. The study was conducted for a range of engine cycles, corresponding to fan pressure ratios of 1.3, 1.45, 1.6, and 1.75. This report is an extension of an effort reported previously. The major conclusions drawn from the prior study, which was restricted to fan pressure ratios of 1.45 and 1.75, are that, for a fan pressure ratio of 1.75, ANC of tones gives about the same suppression as acoustic treatment without ANC. For a fan pressure ratio of 1.45, ANC appears to offer less effectiveness from passive treatment. In the present study, the other two fan pressure ratios are included in a more detailed examination of the benefits of the ANC suppression levels. The key results of this extended study are the following observations: (1) The maximum overall benefit obtained from suppression of BPF alone was 2.5 EPNdB at high fan speeds. The suppression benefit increases with increase in fan pressure ratio (FPR), (2) The maximum overall benefit obtained from suppression of the first three harmonics was 3 EPNdB at high speeds. Suppression benefit increases with increase in FPR, (3) At low FPR, only about 1.0 EPNdB maximum reduction was obtained. Suppression is primarily from reduction of BPF at high FPR values and from the combination of tones at low FPR, (4) The benefit from ANC is about the same as the benefit from passive treatment at fan pressure ratios of 1.75 and 1.60. At the two lower fan pressure ratios, the effectivness of treatment is much greater than that of ANC, and (5) No significant difference in ANC suppression behavior was found from the QCSEE engine database analysis compared to that of the E3 engine database, for the FPR = 1.3 engine cycle. The effects of ANC on EPNL noise reduction are difficult to generalize. It was found that the reduction obtained in any particular case depended upon the frequency of the tones and their shift with rpm, the amount of ANC suppression received by each tone (which depended on its protrusion from the background), and the NOY-value of the tone relative to the NOY-value of other tones and the peak broadband levels, because PNL is determined from the sum of the NOY-values.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198512 , NAS 1.26:198512 , E-10378
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    Effects of Shrouded Stator Cavity Flows on Multistage Axial Compressor Aerodynamic Performance (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experiments were performed on a low-speed multistage axial-flow compressor to assess the effects of shrouded stator cavity flows on aerodynamic performance. Five configurations, which involved changes in seal-tooth leakage rates and/or elimination of the shrouded stator cavities, were tested. Data collected enabled differences in overall individual stage and the third stage blade element performance parameters to be compared. The results show conclusively that seal-tooth leakage ran have a large impact on compressor aerodynamic performance while the presence of the shrouded stator cavities alone seemed to have little influence. Overall performance data revealed that for every 1% increase in the seal-tooth clearance to blade-height ratio the pressure rise dropped up to 3% while efficiency was reduced by 1 to 1.5 points. These observed efficiency penalty slopes are comparable to those commonly reported for rotor and cantilevered stator tip clearance variations. Therefore, it appears that in order to correctly predict overall performance it is equally important to account for the effects of seal-tooth leakage as it is to include the influence of tip clearance flows. Third stage blade element performance data suggested that the performance degradation observed when leakage was increased was brought about in two distinct ways. First, increasing seal-tooth leakage directly spoiled the near hub performance of the stator row in which leakage occurred. Second, the altered stator exit now conditions caused by increased leakage impaired the performance of the next downstream stage by decreasing the work input of the downstream rotor and increasing total pressure loss of the downstream stator. These trends caused downstream stages to progressively perform worse. Other measurements were acquired to determine spatial and temporal flow field variations within the up-and-downstream shrouded stator cavities. Flow within the cavities involved low momentum fluid traveling primarily in the circumferential direction at about 40% of the hub wheel speed. Measurements indicated that the flow within both cavities was much more complex than first envisioned. A vortical flow structure in the meridional plane, similar to a driven cavity, existed within the upstream cavity Furthermore, other spatial and temporal variations in Row properties existed. the most prominent being caused by the upstream potential influence of the downstream blade. This influence caused the fluid within cavities near the leading edges of either stator blades in space or rotor blades in time to be driven radially inward relative to fluid near blade mid-pitch. This influence also produced large unsteady velocity fluctuations in the downstream cavity because of the passing of the downstream rotor blade.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198536 , E-10465 , NAS 1.26:198536
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    Linear Instability of a Uni-Directional Transversely Sheared Mean Flow (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The effect of spanwise-periodic mean-flow distortions (i.e. streamwise-vortex structures) on the evolution of small-amplitude, single-frequency instability waves in an otherwise two-dimensional shear flow is investigated. The streamwise-vortex structures are taken to be just weak enough so that the spatially growing instability waves behave (locally) like linear perturbations about a uni-directional transversely sheared mean flow. Numerical solutions are computed and discussed for both the mean flow and the instability waves. The influence of the streamwise-vortex wavelength on the properties of the most rapidly growing instability wave is also discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198535 , NAS 1.26:198535 , E-10458
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    Concentration Measurements in a Cold Flow Model Annular Combustor Using Laser Induced Fluorescence (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A nonintrusive concentration measurement method is developed for determining the concentration distribution in a complex flow field. The measurement method consists of marking a liquid flow with a water soluble fluorescent dye. The dye is excited by a two dimensional sheet of laser light. The fluorescent intensity is shown to be proportional to the relative concentration level. The fluorescent field is recorded on a video cassette recorder through a video camera. The recorded images are analyzed with image processing hardware and software to obtain intensity levels. Mean and root mean square (rms) values are calculated from these intensity levels. The method is tested on a single round turbulent jet because previous concentration measurements have been made on this configuration by other investigators. The previous results were used to comparison to qualify the current method. These comparisons showed that this method provides satisfactory results. 'Me concentration measurement system was used to measure the concentrations in the complex flow field of a model gas turbine annular combustor. The model annular combustor consists of opposing primary jets and an annular jet which discharges perpendicular to the primary jets. The mixing between the different jet flows can be visualized from the calculated mean and rms profiles. Concentration field visualization images obtained from the processing provide further qualitative information about the flow field.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-182252 , E-9864 , NAS 1.26:182252
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    Influence of Geometry and Flow Variation on Jet Mixing and NO Formation in a Model Staged Combustor Mixer with Eight Orifices (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A series of non-reacting parametric experiments was conducted to investigate the effect of geometric and flow variations on mixing of cold jets in an axis-symmetric, heated cross flow. The confined, cylindrical geometries tested represent the quick mix region of a Rich-Burn/Quick-Mix/Lean-Burn (RQL) combustor. The experiments show that orifice geometry and jet to mainstream momentum-flux ratio significantly impact the mixing characteristic of jets in a cylindrical cross stream. A computational code was used to extrapolate the results of the non-reacting experiments to reacting conditions in order to examine the nitric oxide (NO) formation potential of the configurations examined. The results show that the rate of NO formation is highest immediately downstream of the injection plane. For a given momentum-flux ratio, the orifice geometry that mixes effectively in both the immediate vicinity of the injection plane, and in the wall regions at downstream locations, has the potential to produce the lowest NO emissions. The results suggest that further study may not necessarily lead to a universal guideline for designing a low NO mixer. Instead, an assessment of each application may be required to determine the optimum combination of momentum-flux ratio and orifice geometry to minimize NO formation. Experiments at reacting conditions are needed to verify the present results.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-194473 , NAS 1.26:194473 , E-8614
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    The Effect of Vane-Blade Spacing on Transonic Turbine Stage Performance (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This paper presents the results of a computational study on the effect of axial spacing between the vane and blade rows of a transonic turbine stage. The study was performed on the mid-span section of a high-pressure turbine stage using a quasi-3D, unsteady Navier-Stokes solver that provides a fully interactive vane-blade unsteady flow solution. Three different cases were considered, corresponding to axial spacings of 20%, 40%, and 60% of the vane axial chord. The calculated vane and blade pressure distributions for the 40 percent case were found to compare favorably with experimental measurements acquired in a short-duration shock tunnel. In addition, the analysis shows a marked increase in the amplitude of the unsteady pressure fluctuations on the vane and blade surfaces as the spacing decreases. Time-averaged stage adiabatic efficiency predictions for each case are presented to show the effect of spacing on aerodynamic performance.
    Keywords: Aircraft Propulsion and Power
    Type: Loss Mechanisms and Unsteady Flows in Turbomachines; AGARD-CP-571
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  • 36
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    Optimization of Orifice Geometry for Cross-Flow Mixing in a Cylindrical Duct (1996)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Mixing of gaseous jets in a cross-flow has significant applications in engineering, one example of which is the dilution zone of a gas turbine combustor. Despite years of study, the design of the jet injection in combustors is largely based on practical experience. The emergence of NO(x) regulations for stationary gas turbines and the anticipation of aero-engine regulations requires an improved understanding of jet mixing as new combustor concepts are introduced. For example, the success of the staged combustor to reduce the emission of NO(x) is almost entirely dependent upon the rapid and complete dilution of the rich zone products within the mixing section. It is these mixing challenges to which the present study is directed. A series of experiments was undertaken to delineate the optimal mixer orifice geometry. A cross-flow to core-flow momentum-flux ratio of 40 and a mass flow ratio of 2.5 were selected as representative of a conventional design. An experimental test matrix was designed around three variables: the number of orifices, the orifice length-to- width ratio, and the orifice angle. A regression analysis was performed on the data to arrive at an interpolating equation that predicted the mixing performance of orifice geometry combinations within the range of the test matrix parameters. Results indicate that the best mixing orifice geometry tested involves eight orifices with a long-to-short side aspect ratio of 3.5 at a twenty-three degree inclination from the center-line of the mixing section.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198482 , UCICL-ARTR-93-4 , NAS 1.26:198482 , E-10247
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    Subsonic Jet Noise Reduced With Improved Internal Exhaust Gas Mixers (1996)
    In:  CASI
    Details
    Publication Date: 2018-06-05
    Description: Aircraft noise pollution is becoming a major environmental concern for the world community. The Federal Aviation Administration (FAA) is responding to this concern by imposing more stringent noise restrictions for aircraft certification then ever before to keep the U.S. industry competitive with the rest of the world. At the NASA Lewis Research Center, attempts are underway to develop noise-reduction technology for newer engines and for retrofitting existing engines so that they are as quiet as (or quieter than) required. Lewis conducted acoustic and Laser Doppler Velocimetry (LDV) tests using Pratt & Whitney's Internal Exhaust Gas Mixers (IEGM). The IEGM's mix the core flow with the fan flow prior to their common exhaust. All tests were conducted in Lewis' Aero-Acoustic Propulsion Laboratory--a semihemispheric dome open to the ambient atmosphere. This was the first time Laser Doppler Velocimetry was used in such a facility at Lewis. Jet exhaust velocity and turbulence and the internal velocity fields were detailed. Far-field acoustics were also measured. Pratt & Whitney provided 1/7th scale model test hardware (a 12-lobe mixer, a 20-lobe mixer, and a splitter) for 1.7 bypass ratio engines, and NASA provided the research engineers, test facility, and test time. The Pratt & Whitney JT8D-200 engine power conditions were used for all tests.
    Keywords: Aircraft Propulsion and Power
    Type: Research and Technology 1995; NASA-TM-107111
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    STOVL Jet Efflux Characteristics and their Influence on Propulsion-Induced Effects (1996)
    In:  Other Sources
    Details
    Publication Date: 2019-07-18
    Description: Jet efflux characteristics are a determining factor in STOVL aircraft aero/propulsion induced effects. Subcritical jets may have core lengths which range up to 6 diameters. The shorter core length jets tend to entrain ambient air more rapidly, inducing larger hover lift losses, and decay more rapidly, reducing adverse ground erosion. In transition flight, shorter core length jets show a larger decrease in the lift loss and a slight decrease in nose-up pitching moment. Supercritical pressure ratio jets tend to have longer, higher pressure core lengths with a greater hazard for ground erosion. The decay in the fully developed region is similar for both subcritical and supercritical pressure ratio jets. For subsonic jets the decay is inversely proportional to the distance from the jet exit. In ground effect the supercritical jet induces an oscillating pressure distribution on the ground with reflected shocks and expansions which can increase ground erosion and, at low ground heights, cause non-monotonic lift loss variations.
    Keywords: Aircraft Propulsion and Power
    Type: International Powered Lift Conference; Nov 18, 1996 - Nov 21, 1996; West Palm Beach, FL; United States
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    Two-Dimensional Computational Model for Wave Rotor Flow Dynamics (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A two-dimensional (theta,z) Navier-Stokes solver for multi-port wave rotor flow simulation is described. The finite-volume form of the unsteady thin-layer Navier-Stokes equations are integrated in time on multi-block grids that represent the stationary inlet and outlet ports and the moving rotor passages of the wave rotor. Computed results are compared with three-port wave rotor experimental data. The model is applied to predict the performance of a planned four-port wave rotor experiment. Two-dimensional flow features that reduce machine performance and influence rotor blade and duct wall thermal loads are identified. The performance impact of rounding the inlet port wall, to inhibit separation during passage gradual opening, is assessed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107192 , ARL-TR-924 , E-10164 , NAS 1.15:107192 , Turbo Expo 1996; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    An Experimental Determination of Losses in a 3-Port Wave Rotor (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Wave rotors, used in a gas turbine topping cycle, offer a potential route to higher specific power and lower specific fuel consumption. In order to exploit this potential properly, it is necessary to have some realistic means of calculating wave rotor performance, taking losses into account, so that wave rotors can be designed for good performance. This in turn requires a knowledge of the loss mechanisms. The experiment reported here was designed as a statistical experiment to identify the losses due to finite passage opening time, friction, and leakage. For simplicity, the experiment used a 3-port, flow divider, wave cycle, but the results should be applicable to other cycles. A 12 inch diameter rotor was used, with two different lengths, 9 inches and 18 inches, and two different passage widths, 0.25 inch and 0.54 inch, in order to vary friction and opening time. To vary leakage, moveable end-walls were provided so that the rotor to end-wall gap could be adjusted. The experiment is described, and the results are presented, together with a parametric fit to the data. The fit shows that there will be an optimum passage width for a given wave rotor, since, as the passage width increases, friction losses decrease, but opening-time losses increase, and vice-versa. Leakage losses can be made small at reasonable gap sizes.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198456 , E-10079 , NAS 1.26:198456 , Turbo Expo 1996; Jun 10, 1996 - Jun 13, 1996; Birmingham, England; United States
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    A CFD Study of Turbojet and Single-Throat Ramjet Ejector Interaction (1996)
    In:  Other Sources
    Details
    Publication Date: 2019-07-17
    Description: Supersonic ejector-diffuse systems have application in driving an advanced airbreathing propulsion system, consisting of turbojet engines acting as the primary and a single throat ramjet acting as the secondary. The turbojet engines are integrated into the single throat ramjet to minimize variable geometry and eliminate redundant propulsion components. The result is a simple, lightweight system that is operable from takeoff to high Mach numbers. At this high Mach number (approximately Mach 3.0), the turbojets are turned off and the high speed ramjet/scramjet take over and drive the vehicle to Mach 6.0. The turbojet-ejector-ramjet system consists of nonafterburning turbojet engines with ducting canted at 20 degrees to supply supersonic flow (downstream of CD nozzle) to the horizontal ramjet duct at a supply total pressure and temperature. Two conditions were modelled by a 2-D full Navier Stokes code at Mach 2.0. The code modelled the Fabri choke as well as the non-Fabri non critical case, using a computational throat to supply the back pressure. The results, which primarily predict the secondary mass flow rate and the mixed conditions at the ejector exit were in reasonable agreement with the 1-D cycle code (TBCC).
    Keywords: Aircraft Propulsion and Power
    Type: Paper-28 , HBCUs Research Conference Agenda and Abstracts; 39; NASA-CP-10189
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    Shock-free supersonic elliptic nozzles and method of forming same (1996)
    In:  CASI
    Details
    Publication Date: 2019-08-28
    Description: A method of forming a shock-free supersonic elliptic nozzle, in which the nozzle to be designed is divided into three sections, a circular-to-elliptic section which begins at a circular nozzle inlet, an elliptic subsonic section downstream from the circular-to-elliptic section and a supersonic section downstream from the elliptic subsonic section. The maximum and minimum radii for each axial point in the circular-to-elliptic section and the elliptic subsonic section are then separately determined, the maximum and minimum radii being the radii for the widest part of an elliptic cross-section and the narrowest part of the elliptic cross-section, respectively. The maximum and minimum radii for each axial point in the supersonic section are determined based on the Method of Characteristics, Then, each of the three sections are based on the maximum and minimum radii for each axial point in the section. The resulting nozzle is acoustically superior.
    Keywords: Aircraft Propulsion and Power
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    Calculation of tip clearance effects in a transonic compressor rotor (1996)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    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 concerning 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 computer 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: NASA-TM-107216 , NAS 1.15:107216 , E-10242 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham, UK; United States
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    Macroscopic balance model for wave rotors (1996)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A mathematical model for multi-port wave rotors is described. The wave processes that effect energy exchange within the rotor passage are modeled using one-dimensional gas dynamics. Macroscopic mass and energy balances relate volume-averaged thermodynamic properties in the rotor passage control volume to the mass, momentum, and energy fluxes at the ports. Loss models account for entropy production in boundary layers and in separating flows caused by blade-blockage, incidence, and gradual opening and closing of rotor passages. The mathematical model provides a basis for predicting design-point wave rotor performance, port timing, and machine size. Model predictions are evaluated through comparisons with CFD calculations and three-port wave rotor experimental data. A four-port wave rotor design example is provided to demonstrate model applicability. The modeling approach is amenable to wave rotor optimization studies and rapid assessment of the trade-offs associated with integrating wave rotors into gas turbine engine systems.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107114 , ARL-TR-925 , AIAA Paper 96-0243 , E-10022 , NAS 1.15:107114 , 34th Aerospace Sciences Meeting and Exhibit; Jan 15, 1996 - Jan 18, 1996; Reno, NV; United States
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    Numerical Simulation of Flow in a Whirling Annular Seal and Comparison With Experiments (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The turbulent flow field inside a whirling annular seal was simulated by using SCISEAL, a three-dimensional computational fluid dynamics code. The rotor center described a circular synchronous whirl. A rotating frame transformation was used to make the problem quasi-steady. The flow field at an axial Reynolds number of 24,000 and a Taylor number of 6600 was simulated. The standard kappa-epsilon model with wall functions and the low-Reynolds-number model were used to treat turbulence. An experimentally measured velocity field was used at the inlet boundary. Numerical predictions of the velocities and the stator wall pressures compared well with experimental data. Both turbulence models yielded nearly the same results. The capability of the SCISEAL code to analyze this complex flow field was demonstrated; the isotropic turbulence models performed adequately on this nonisotropic turbulence flow.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107117 , E-10026 , NAS 1.15:107117 , Sixth International Symposium on Transport Phenomena and Dynamics of Rotating Machinery; Feb 25, 1996 - Feb 29, 1996; Honolulu, HI; United States
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    Estimating Engine Airflow in Gas-Turbine Powered Aircraft with Clean and Distorted Inlet Flows (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The P404-GF-400 Powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the impact of inlet-generated total-pressure distortion on estimating levels of engine airflow. Five airflow estimation methods were studied. The Reference Method was a fan corrected airflow to fan corrected speed calibration from an uninstalled engine test. In-flight airflow estimation methods utilized the average, or individual, inlet duct static- to total-pressure ratios, and the average fan-discharge static-pressure to average inlet total-pressure ratio. Correlations were established at low distortion conditions for each method relative to the Reference Method. A range of distorted inlet flow conditions were obtained from -10 deg. to +60 deg. angle of attack and -7 deg. to +11 deg. angle of sideslip. The individual inlet duct pressure ratio correlation resulted in a 2.3 percent airflow spread for all distorted flow levels with a bias error of -0.7 percent. The fan discharge pressure ratio correlation gave results with a 0.6 percent airflow spread with essentially no systematic error. Inlet-generated total-pressure distortion and turbulence had no significant impact on the P404-GE400 engine airflow pumping. Therefore, a speed-flow relationship may provide the best airflow estimate for a specific engine under all flight conditions.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-198052 , NAS 1.26:198052 , H-2127 , High-Angle-of-Attack Technology; Sep 17, 1996 - Sep 19, 1996; Hampton, VA; United States
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    Computation of Radiation Heat Transfer in Aeroengine Combustors (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In this report the highlights of the research completed for the NASA are summarized. This research has been completed in the form of two Ph.D. theses by Chai (1994) and Parthasarathy (1996). Readers are referred to these theses for a complete details of the work and lists of references. In the following sections, first objectives of this research are introduced, then the finite-volume method for radiation heat transfer is described, and finally computations of radiative heat transfer in non-gray participating media is presented.
    Keywords: Aircraft Propulsion and Power
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    Evaluation of Water Injection Effect on NO(x) Formation for a Staged Gas Turbine Combustor (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: NO(x) emission control by water injection on a staged turbine combustor (STC) was modeled using the KIVA-2 code with modification. Water is injected into the rich-burn combustion zone of the combustor by a single nozzle. Parametric study for different water injection patterns was performed. Results show NO(x) emission will decrease after water being injected. Water nozzle location also has significant effect for NO formation and fuel ignition. The chemical kinetic model is also sensitive to the excess water. Through this study, a better understanding of the physics and chemical kinetics is obtained, this will enhance the STC design process.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-204498 , NAS 1.26:204498 , AIAA Paper 96-0706 , Aerospace Sciences; Jan 15, 1996 - Jan 18, 1996; Reno, NV; United States
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    Velocity and drop size measurements in a confined, swirl-stabilized, combusting spray (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Drop size and velocity measurements in a confined, swirl-stabilized, reacting spray are presented. The configuration consisted of a center-mounted research air-assist atomizer surrounded by a coflowing air stream. A quartz tube surrounded the burner and provided the confinement. Both the air-assist and coflow streams had swirl imparted to them in the same direction with 45-degree-angle swirlers. The fuel and air entered the combustor at ambient temperature. The gas-phase measurements reported were obtained from the velocity drops with a mean diameter of four microns. Heptane fuel was used for all the experiments. Measurements of drop size and velocity, gas-phase velocity and drop number flux are reported for axial distances of 23, 5, 10, 15, 25, and 50 mm downstream of the nozzle. The measurements were performed using a two-component phase/Doppler particle analyzer. Profiles across the entire flowfield are presented.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107243 , NAS 1.15:107243 , AIAA Paper 96-3164 , E-10291 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Oscillating cascade aerodynamics at large mean incidence (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low aspect ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies of up to 1.2 for out-of-phase oscillations at a Mach number of 0.5 and chordal incidence angles of 0 deg and 10 deg; the Reynolds number was 0.9 x l0(exp 6). For the 10 deg chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107247 , NAS 1.15:107247 , E-10300 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Brimingham; United Kingdom
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    Flow Coupling Effects in Jet-in-Crossflow Flowfields (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The combustor designer is typically required to design liner orifices that effectively mix air jets with crossflow effluent. CFD combustor analysis is typically used in the design process; however the jets are usually assumed to enter the combustor with a uniform velocity and turbulence profile. The jet-mainstream flow coupling is usually neglected because of the computational expense. This CFD study was performed to understand the effect of jet-mainstream flow coupling, and to assess the accuracy of jet boundary conditions that are commonly used in combustor internal calculations. A case representative of a plenum-fed quick-mix section of a Rich Burn/Quick Mix/Lean Burn combustor (i.e. a jet-mainstream mass-flow ratio of about 3 and a jet-mainstream momentum-flux ratio of about 30) was investigated. This case showed that the jet velocity entering the combustor was very non-uniform, with a low normal velocity at the leading edge of the orifice and a high normal velocity at the trailing edge of the orifice. Three different combustor-only cases were analyzed with uniform inlet jet profile. None of the cases matched the plenum-fed calculations. To assess liner thickness effects, a thin-walled case was also analyzed. The CFD analysis showed the thin-walled jets had more penetration than the thick-walled jets.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107257 , NAS 1.15:107257 , AIAA Paper 96-2762 , E-10318 , AIAA/ASME/SAE/ASEE Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Task 7: Endwall treatment inlet flow distortion analysis (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The overall objective of this study was to develop a 3-D numerical analysis for compressor casing treatment flowfields, and to perform a series of detailed numerical predictions to assess the effectiveness of various endwall treatments for enhancing the efficiency and stall margin of modern high speed fan rotors. Particular attention was given to examining the effectiveness of endwall treatments to counter the undesirable effects of inflow distortion. Calculations were performed using three different gridding techniques based on the type of casing treatment being tested and the level of complexity desired in the analysis. In each case, the casing treatment itself is modeled as a discrete object in the overall analysis, and the flow through the casing treatment is determined as part of the solution. A series of calculations were performed for both treated and untreated modern fan rotors both with and without inflow distortion. The effectiveness of the various treatments were quantified, and several physical mechanisms by which the effectiveness of endwall treatments is achieved are discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-195468 , NAS 1.26:195468 , E-9647
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    Quantification of Gear Tooth Damage by Optimal Tracking of Vibration Signatures (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a technique for quantifying the wear or damage of gear teeth in a transmission system. The procedure developed in this study can be applied as a part of either an onboard machine health-monitoring system or a health diagnostic system used during regular maintenance. As the developed methodology is based on analysis of gearbox vibration under normal operating conditions, no shutdown or special modification of operating parameters is required during the diagnostic process. The process of quantifying the wear or damage of gear teeth requires a set of measured vibration data and a model of the gear mesh dynamics. An optimization problem is formulated to determine the profile of a time-varying mesh stiffness parameter for which the model output approximates the measured data. The resulting stiffness profile is then related to the level of gear tooth wear or damage. The procedure was applied to a data set generated artificially and to another obtained experimentally from a spiral bevel gear test rig. The results demonstrate the utility of the procedure as part of an overall health-monitoring system.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107100 , E-10029 , NAS 1.15:107100 , 6th International Symposium on Transport Phenomena and Dynamics of Rotating Machinery; Feb 25, 1996 - Feb 29, 1996; Honolulu, HA; United States
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    Dynamic Modeling of Starting Aerodynamics and Stage Matching in an Axi-Centrifugal Compressor (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A DYNamic Turbine Engine Compressor Code (DYNTECC) has been modified to model speed transients from 0-100% of compressor design speed. The impetus for this enhancement was to investigate stage matching and stalling behavior during a start sequence as compared to rotating stall events above ground idle. The model can simulate speed and throttle excursions simultaneously as well as time varying bleed flow schedules. Results of a start simulation are presented and compared to experimental data obtained from an axi-centrifugal turboshaft engine and companion compressor rig. Stage by stage comparisons reveal the front stages to be operating in or near rotating stall through most of the start sequence. The model matches the starting operating line quite well in the forward stages with deviations appearing in the rearward stages near the start bleed. Overall, the performance of the model is very promising and adds significantly to the dynamic simulation capabilities of DYNTECC.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107338 , E-10477 , NAS 1.15:107338 , ARL-TR-1107 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    Computation of Kinetics for the Hydrogen/Oxygen System Using the Thermodynamic Method (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A new method for predicting chemical rate constants using thermodynamics has been applied to the hydrogen/oxygen system. This method is based on using the gradient of the Gibbs free energy and a single proportionality constant D to determine the kinetic rate constants. Using this method the rate constants for any gas phase reaction can be computed from thermodynamic properties. A modified reaction set for the H/O system is determined. A11 of the third body efficiencies M are taken to be unity. Good agreement was obtained between the thermodynamic method and the experimental shock tube data. In addition, the hydrogen bromide experimental data presented in previous work is recomputed with M's of unity.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM 170340 , NAS 1.15:170340 , WSS-96F-090 , E-10481 , Oct 18, 1996 - Oct 19, 1996; Los Angeles, CA; United States
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    Performance Improvement Through Indexing of Turbine Airfoils (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An experimental/analytical study has been conducted to determine the performance improvements achievable by circumferentially indexing succeeding rows of turbine stator airfoils. A series of tests was conducted to experimentally investigate stator wake clocking effects on the performance of the space shuttle main engine (SSME) alternate turbopump development (ATD) fuel turbine test article (TTA). The results from this study indicate that significant increases in stage efficiency can be attained through application of this airfoil clocking concept. Details of the experiment and its results are documented in part 1 of this paper. In order to gain insight into the mechanisms of the performance improvement, extensive computational fluid dynamics (CFD) simulations were executed. The subject of the present paper is the initial results from the CFD investigation of the configurations and conditions detailed in part 1 of the paper. To characterize the aerodynamic environments in the experimental test series, two-dimensional (2D), time accurate, multistage, viscous analyses were performed at the TTA midspan. Computational analyses for five different circumferential positions of the first stage stator have been completed. Details of the computational procedure and the results are presented. The analytical results verify the experimentally demonstrated performance improvement and are compared with data whenever possible. Predictions of time-averaged turbine efficiencies as well as gas conditions throughout the flow field are presented. An initial understanding of the turbine performance improvement mechanism based on the results from this investigation is described.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-111805 , NAS 1.15:111805 , Gas Turbines; Jun 04, 1995 - Jun 08, 1995; Houston, TX; United States|Journal of Turbomachinery
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    Mixed composition materials suitable for vacuum web sputter coating (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Ion beam sputter deposition techniques were used to investigate simultaneous sputter etching of two component targets so as to produce mixed composition films. Although sputter deposition has been largely confined to metals and metal oxides, at least one polymeric material, poly-tetra-fluorethylene, has been demonstrated to produce sputtered fragments which repolymerize upon deposition to produce a highly cross-linked fluoropolymer resembling that of the parent target Fluoropolymer-filled silicon dioxide and fluoropolymer-filled aluminum oxide coatings have been deposited by means of ion beam sputter coat deposition resulting in films having material properties suitable for aerospace and commercial applications. The addition of fluoropolymer to silicon dioxide films was found to increase the hydrophobicity of the resulting mixed films; however, adding fluoropolymer to aluminum oxide films resulted in a reduction in hydrophobicity, thought to be caused by aluminum fluoride formation.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107264 , NAS 1.15:107264 , E-10329 , Annual Technical Conference; May 08, 1996 - May 10, 1996; Philadelphia, PA; United States
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    Effects of Inlet Flow Conditions on Crossflow Jet Mixing (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: An experimental investigation of the effects of mainstream turbulence, mainstream swirl and non-symmetric mass addition has been conducted for the isothermal mixing of multiple jets injected into a confined rectangular crossflow. Jet penetration and mixing in the near field was studied using planar Mie scattering to measure time-averaged mixture fraction distributions. Orifice configurations were used that were optimized for mixing performance based on previous experimental and computational results for a homogeneous approach flow. Mixing effectiveness, determined using a spatial unmixedness parameter based on the variance of the mean jet concentration distributions, was found to be minimally affected by free-stream turbulence but significantly influenced by the addition of swirl to the mainstream. The results for non-symmetric mass addition indicate that the concentration distribution of the flowfield can be tailored if desired.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107258 , NAS 1.15:107258 , AIAA Paper 96-2881 , E-10319 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Forced response testing of an axi-centrifugal turboshaft engine (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Dynamic data from tests of a T55-L-712 engine are presented. Engine stall/surge data were analyzed using digital signal processing techniques. In addition, forced response testing (system identification studies) was done at various engine speeds. Forced response testing was done using eight jet ejectors approximately equally circumferentially spaced about the compressor front face. This paper presents some preliminary results for the ground idle (approximately 60% of design speed) point. Brief descriptions of the jet injection system, the test matrix, and analysis techniques used are presented. Results of these analyses indicate a substantial transfer of energy across the compressor first stage at some frequencies and that the ejectors are effective in modifying the local flow conditions in front of the first compressor stage.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107282 , NAS 1.15:107282 , AIAA Paper 96-2573 , E-10357 , ARL-TR-1151 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Minimum Weight Design of a Generic Axisymmetric Inlet (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A new minimum weight design method for high-speed axisymmetric inlets was demonstrated on a generic inlet. The method uses Classical Beam Theory and shell buckling to determine the minimum required equivalent isotropic thickness for a stiffened shell based on prescribed structural design requirements and load conditions. The optimum spacing and equivalent isotropic thickness of ring frame supports are computed to prevent buckling. The method thus develops a preliminary structural design for the inlet and computes the structural weight. Finite element analyses were performed on the resulting inlet design to evaluate the analytical results. Comparisons between the analytical and finite element stresses and deflections identified areas needing improvement in the analytical method. The addition of the deflection due to shear and a torsional buckling failure mode to the new method brought its results in line with those from the finite element analyses. Final validation of the new method will be made using data from actual inlets.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107288 , E-10363 , NAS 1.15:107288 , AIAA Paper 96-2550 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Dimension Determination of Precursive Stall Events in a Single Stage High Speed Compressor (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: This paper presents a study of the dynamics for a single-stage, axial-flow, high speed compressor core, specifically, the NASA Lewis rotor stage 37. Due to the overall blading design for this advanced core compressor, each stage has considerable tip loading and higher speed than most compressor designs, thus, the compressor operates closer to the stall margin. The onset of rotating stall is explained as bifurcations in the dynamics of axial compressors. Data taken from the compressor during a rotating stall event is analyzed. Through the use of a box-assisted correlation dimension methodology, the attractor dimension is determined during the bifurcations leading to rotating stall. The intent of this study is to examine the behavior of precursive stall events so as to predict the entrance into rotating stall. This information may provide a better means to identify, avoid or control the undesirable event of rotating stall formation in high speed compressor cores.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107268 , E-10335 , NAS 1.15:107268 , Technical Conference on Nonlinear Dynamics and Full Spectrum Processing; Jul 10, 1995 - Jul 14, 1995; Mystic, CT; United States
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    BladeCAD: An Interactive Geometric Design Tool for Turbomachinery Blades (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A new metthodology for interactive design of turbomachinery blades is presented. Software implementation of the meth- ods provides a user interface that is intuitive to aero-designers while operating with standardized geometric forms. The primary contribution is that blade sections may be defined with respect to general surfaces of revolution which may be defined to represent the path of fluid flow through the turbomachine. The completed blade design is represented as a non-uniform rational B-spline (NURBS) surface and is written to a standard IGES file which is portable to most design, analysis, and manufacturing applications.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107262 , E-10327 , NAS 1.15:107262 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    High Stability Engine Control (HISTEC) (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Future aircraft turbine engines, both commercial and military, must be able to successfully accommodate expected increased levels of steady-state and dynamic engine-face distortion. The current approach of incorporating a sufficient component 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, high-stability, integrated engine control system that uses measurement-based, real-time 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, consisting of a Distortion Estimation System and a Stability Management Control, has been designed and developed. The Distortion Estimation System uses a small number of high-response pressure sensors at the engine face to calculate indicators of the type and extent of distortion in real time. The Stability Management Control, through direct control of the fan and compressor pressure ratio, accommodates the distortion by transiently increasing the amount of stall margin available based on information from the Distortion Estimation System. Simulation studies have shown the HISTEC distortion tolerant control is able to successfully estimate and accommodate time-varying distortion. Currently, hardware and software systems necessary for flight demonstration of the HISTEC concept are being designed and developed. The HISTEC concept will be flight tested in early 1997.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107272 , E-10339 , NAS 1.15:107272 , AIAA Paper 96-2586 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    A numerical investigation of premixed combustion in wave rotors (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Wave rotor cycles which utilize premixed combustion processes within the passages are examined numerically using a one-dimensional CFD-based simulation. Internal-combustion wave rotors are envisioned for use as pressure-gain combustors in gas turbine engines. The simulation methodology is described, including a presentation of the assumed governing equations for the flow and reaction in the channels, the numerical integration method used, and the modeling of external components such as recirculation ducts. A number of cycle simulations are then presented which illustrate both turbulent-deflagration and detonation modes of combustion. Estimates of performance and rotor wall temperatures for the various cycles are made, and the advantages and disadvantages of each are discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107242 , NAS 1.15:107242 , E-10288 , ASME-96-GT-116 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    Vortex generator installation studies on steady state and dynamic inlet distortion (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The theoretical and experimental work carried out under the NASA/MOD Joint Aeronautical Program has shown that CFD vortex generator installations designs successfully managed inlet duct flow distortion and that significant benefits in flow unsteadiness at the engine face were also present. The main conclusions to date from the collaborative effort between NASA/Lewis and DRA/Bedford are as follows: (1) vortex generator installations can be designed to be effective over a wide range of inlet operating conditions using Computational Fluid Dynamics and formal optimization procedures, (2) reductions in steady state engine face distortion of up to 80% have been measured in the M2129 inlet S-duct using CFD designed vortex generator installations, (3) reductions in flow unsteadiness of up to 80% have been measured in the W129 inlet S-duct using CFD designed vortex generator installations, and (4) the Reduced Navier-Stokes code RNS3D is a useful tool to design vortex generator installations to manage engine face distortions over a wide range of inlet operating conditions.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107220 , NAS 1.15:107220 , AIAA Paper 96-3279 , E-10252 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Mixing of Multiple Jets with a Confined Subsonic Crossflow in a Cylindrical Duct (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: This paper summarizes NASA-supported experimental and computational results on the mixing of a row of jets with a confined subsonic crossflow in a cylindrical duct. The studies from which these results were derived investigated flow and geometric variations typical of the complex 3-D flowfield in the combustion chambers in gas turbine engines. The principal observations were that the momentum-flux ratio and the number of orifices were significant variables. Jet penetration was critical, and jet penetration decreased as either the number of orifices increased or the momentum-flux ratio decreased. It also appeared that jet penetration remained similar with variations in orifice size, shape, spacing, and momentum-flux ratio when the number of orifices was proportional to the square-root of the momentum-flux ratio. In the cylindrical geometry, planar variances are very sensitive to events in the near wall region, so planar averages must be considered in context with the distributions. The mass-flow ratios and orifices investigated were often very large (mass-flow ratio greater than 1 and ratio of orifice area-to-mainstream cross-sectional area up to 0.5), and the axial planes of interest were sometimes near the orifice trailing edge. Three-dimensional flow was a key part of efficient mixing and was observed for all configurations. The results shown also seem to indicate that non-reacting dimensionless scalar profiles can emulate the reacting flow equivalence ratio distribution reasonably well. The results cited suggest that further study may not necessarily lead to a universal 'rule of thumb' for mixer design for lowest emissions, because optimization will likely require an assessment for a specific application.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107185 , NAS 1.15:107185 , ASME-96-GT-482 , E-9996 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    Performance Benefits for Wave Rotor-Topped Gas Turbine Engines (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The benefits of wave rotor-topping in turboshaft engines, subsonic high-bypass turbofan engines, auxiliary power units, and ground power units are evaluated. The thermodynamic cycle performance is modeled using a one-dimensional steady-state code; wave rotor performance is modeled using one-dimensional design/analysis codes. Design and off-design engine performance is calculated for baseline engines and wave rotor-topped engines, where the wave rotor acts as a high pressure spool. The wave rotor-enhanced engines are shown to have benefits in specific power and specific fuel flow over the baseline engines without increasing turbine inlet temperature. The off-design steady-state behavior of a wave rotor-topped engine is shown to be similar to a conventional engine. Mission studies are performed to quantify aircraft performance benefits for various wave rotor cycle and weight parameters. Gas turbine engine cycles most likely to benefit from wave rotor-topping are identified. Issues of practical integration and the corresponding technical challenges with various engine types are discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107193 , E-10168 , NAS 1.15:107193 , ARL-TR-1065 , 41st Turbo Expo 1996; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    A Passive Cavity Concept for Improving the Off-Design Performance of Fixed-Geometry Exhaust Nozzles (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to study a passive cavity concept for improving the off-design performance of fixed-geometry exhaust nozzles. Passive cavity ventilation (through a porous surface) was applied to divergent flap surfaces and tested at static conditions in a sub-scale, nonaxisymmetric, convergent-divergent nozzle. As part of a comprehensive investigation, force, moment and pressure measurements were taken and focusing schlieren flow visualization was obtained for a baseline configuration and D passive cavity configurations. All tests were conducted with no external flow and high-pressure air was used to simulate jet-exhaust flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable shock-induced boundary-layer separation at off-design conditions, which came about through the natural tendency of overexpanded exhaust flow to satisfy conservation requirements by detaching from the nozzle divergent flaps. Passive cavity ventilation added the ability to control off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. Separation alleviation offers potential for installed nozzle performance benefits by reducing drag at forward flight speeds, even though it may reduce off-design static thrust efficiency as much as 3.2 percent. Encouraging stable separation of the exhaust flow offers significant performance improvements at static, low NPR and low Mach number flight conditions by improving off-design static thrust efficiency as much as 2.8 percent. By designing a fixed-geometry nozzle with fully porous divergent flaps, where both cavity location and percent open porosity of the flaps could be varied, passive flow control would make it possible to improve off-design nozzle performance across a wide operating range. In addition, the ability to encourage separation on one flap while alleviating it on the other makes it possible to generate thrust vectoring in the nozzle through passive flow control.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-111589 , NAS 1.15:111589 , AIAA Paper 96-2541 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Seals Code Development Workshop (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Seals Workshop of 1995 industrial code (INDSEAL) release include ICYL, GCYLT, IFACE, GFACE, SPIRALG, SPIRALI, DYSEAL, and KTK. The scientific code (SCISEAL) release includes conjugate heat transfer and multidomain with rotordynamic capability. Several seals and bearings codes (e.g., HYDROFLEX, HYDROTRAN, HYDROB3D, FLOWCON1, FLOWCON2) are presented and results compared. Current computational and experimental emphasis includes multiple connected cavity flows with goals of reducing parasitic losses and gas ingestion. Labyrinth seals continue to play a significant role in sealing with face, honeycomb, and new sealing concepts under investigation for advanced engine concepts in view of strict environmental constraints. The clean sheet approach to engine design is advocated with program directions and anticipated percentage SFC reductions cited. Future activities center on engine applications with coupled seal/power/secondary flow streams.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CP-10181 , E-9945 , NAS 1.55:10181 , Seals Code Development Workshop; Jun 14, 1995 - Jun 15, 1995; Cleveland, OH; United States
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    Noise control using a plate radiator and an acoustic resonator (1996)
    In:  CASI
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    Publication Date: 2019-07-12
    Description: An active noise control subassembly for reducing noise caused by a source (such as an aircraft engine) independent of the subassembly. A noise radiating panel is bendably vibratable to generate a panel noise canceling at least a portion of the source noise. A piezoceramic actuator plate is connected to the panel. A front plate is spaced apart from the panel and the first plate, is positioned generally between the source noise and the panel, and has a sound exit port. A first pair of spaced-apart side walls each generally abut the panel and the front plate so as to generally enclose a front cavity to define a resonator.
    Keywords: Aircraft Propulsion and Power
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    Comparisons of Rig and Engine Dynamic Events in the Compressor of an Axi-Centrifugal Turboshaft Engine (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Steady state and dynamic data were acquired in a T55-L-712 compressor rig. In addition, a T55-L-12 engine was instrumented and similar data were acquired. Rig and engine stall/surge data were analyzed using modal techniques. This paper compares rig and engine preliminary results for the ground idle (approximately 60% of design speed) point. The results of these analyses indicate both rig and engine dynamic event are preceded by indications of traveling wave energy in front of the compressor face. For both rig and engine, the traveling wave energy contains broad band energy with some prominent narrow peaks and, while the events are similar in many ways, some noticeable differences exist between the results of the analyses of rig data and engine data.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107339 , NAS 1.15:107339 , ASME-96-GT-239 , ARL-TR-1108 , E-10478 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    Application of the kappa-omega Turbulence Model to Quasi-Three-Dimensional Turbomachinery Flows (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Many computational fluid dynamics codes for turbomachinery use the Baldwin-Lomax (B-L) turbulence model. It is easy to implement in two dimensions and works well for predicting overall turbomachinery performance. However, it is awkward to implement in three dimensions, often has difficulty finding the length scale, has a crude transition model, and neglects freestream turbulence, surface roughness, and mass injection. The kappa-omega model developed by Wilcox is an appealing alternative for several reasons. First, it is the only two-equation model that can be integrated to the wall without requiring damping functions or the distance to the wall, and hence, should behave well numerically. Second, the effects of freestream turbulence, surface roughness, and mass injection are easily included in the model. Finally, transition can be simulated using the low Reynolds number version of the model. Menter applied the kappa-model to external flows and showed very good results for flows with adverse pressure gradients. Liu and Zheng described their implementation of the kappa-model in a cascade code that included an area change term to account for endwall convergence. They validated the model for a flat plate, and compared the B-L and kappa-models to measured surface pressures for a low-pressure turbine cascade. Since they did not use the low Reynolds number version of the model, their results showed problems resulting from early transition. In this Note the low Reynolds number kappa-model was incorporated in the author's quasi-three-dimensional turbomachinery analysis code. The code includes the effects of rotation, radius change, and stream-surface thickness variation, and also includes the B-L turbulence model. The kappa-omega model was implemented using many of Menter's recommendations and an implicit approximate-factorization scheme described by Baldwin and Barth. The model was tested for a transonic compressor with rotation and variable stream-surface radius and height, and for a transonic turbine vane with transition and heat transfer. Results were compared to the B-L model and to experimental data.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-112767 , NAS 1.15:112767 , AIAA Paper 96-0248 , Aerospace Sciences; Jan 15, 1996 - Jan 19, 1996; Reno, NV; United States|J. Propulsion: Technical Notes; 12; 6; 1176-1179
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    Comparison of Predicted Low Speed Fan Rotor/Stator Interaction Modes to Measured (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The V072 Rotor Wake/Stator Interaction Code is widely used as a state-of-the-art prediction code. This paper validates the code by comparing experimentally measured mode levels to those predicted by V072. The experimental mode levels were measured by the Rotating Rake system installed on the 48 inch Active Noise Control Fan at NASA Lewis Research Center. V072 predicted mode levels by inputting the actual wake profiles of the ANCF rotor measured by a 2-component hotwire. The mode levels were also predicted from the V072 wake models. V072 reasonably predicts the mode levels within the design limits of the code.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107462 , NAS 1.15:107462 , AIAA Paper 97-1609 , E-10748 , Aeroacoustic; May 08, 1997 - May 12, 1997; Atlanta, GA; United States
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    Experimental Characterization of Gas Turbine Emissions at Simulated Flight Altitude Conditions (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: NASA's Atmospheric Effects of Aviation Project (AEAP) is developing a scientific basis for assessment of the atmospheric impact of subsonic and supersonic aviation. A primary goal is to assist assessments of United Nations scientific organizations and hence, consideration of emissions standards by the International Civil Aviation Organization (ICAO). Engine tests have been conducted at AEDC to fulfill the need of AEAP. The purpose of these tests is to obtain a comprehensive database to be used for supplying critical information to the atmospheric research community. It includes: (1) simulated sea-level-static test data as well as simulated altitude data; and (2) intrusive (extractive probe) data as well as non-intrusive (optical techniques) data. A commercial-type bypass engine with aviation fuel was used in this test series. The test matrix was set by parametrically selecting the temperature, pressure, and flow rate at sea-level-static and different altitudes to obtain a parametric set of data.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-112751 , NAS 1.15:112751 , AD-A314312 , AEDC-TR-96-3
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    AST Combustion Workshop: Diagnostics Working Group Report (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A workshop was convened under NASA's Advanced Subsonics Technologies (AST) Program. Many of the principal combustion diagnosticians from industry, academia, and government laboratories were assembled in the Diagnostics/Testing Subsection of this workshop to discuss the requirements and obstacles to the successful implementation of advanced diagnostic techniques to the test environment of the proposed AST combustor. The participants, who represented the major relevant areas of advanced diagnostic methods currently applied to combustion and related fields, first established the anticipated AST combustor flowfield conditions. Critical flow parameters were then examined and prioritized as to their importance to combustor/fuel injector design and manufacture, environmental concerns, and computational interests. Diagnostic techniques were then evaluated in terms of current status, merits and obstacles for each flow parameter. All evaluations are presented in tabular form and recommendations are made on the best-suited diagnostic method to implement for each flow parameter in order of applicability and intrinsic value.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107354 , NAS 1.15:107354 , E-10508 , Aug 10, 1994 - Aug 11, 1994; Cleveland, OH; United States
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    Thrust Vectoring on the NASA F-18 High Alpha Research Vehicle (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Investigations into a multiaxis thrust-vectoring system have been conducted on an F-18 configuration. These investigations include ground-based scale-model tests, ground-based full-scale testing, and flight testing. This thrust-vectoring system has been tested on the NASA F-18 High Alpha Research Vehicle (HARV). The system provides thrust vectoring in pitch and yaw axes. Ground-based subscale test data have been gathered as background to the flight phase of the program. Tests investigated aerodynamic interaction and vane control effectiveness. The ground-based full-scale data were gathered from static engine runs with image analysis to determine relative thrust-vectoring effectiveness. Flight tests have been conducted at the NASA Dryden Flight Research Center. Parameter identification input techniques have been developed. Individual vanes were not directly controlled because of a mixer-predictor function built into the flight control laws. Combined effects of the vanes have been measured in flight and compared to combined effects of the vanes as predicted by the cold-jet test data. Very good agreement has been found in the linearized effectiveness derivatives.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-4771 , NAS 1.15:4771 , H-2139 , High-Angle-of-Attack Technology Conference; Sep 17, 1996 - Sep 19, 1996; Hampton, VA; United States
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    Comparative Study of Advanced Turbulence Models for Turbomachinery (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: A computational study has been undertaken to study the performance of advanced phenomenological turbulence models coded in a modular form to describe incompressible turbulent flow behavior in two dimensional/axisymmetric and three dimensional complex geometry. The models include a variety of two equation models (single and multi-scale k-epsilon models with different near wall treatments) and second moment algebraic and full Reynolds stress closure models. These models were systematically assessed to evaluate their performance in complex flows with rotation, curvature and separation. The models are coded as self contained modules that can be interfaced with a number of flow solvers. These modules are stand alone satellite programs that come with their own formulation, finite-volume discretization scheme, solver and boundary condition implementation. They will take as input (from any generic Navier-Stokes solver) the velocity field, grid (structured H-type grid) and computational domain specification (boundary conditions), and will deliver, depending on the model used, turbulent viscosity, or the components of the Reynolds stress tensor. There are separate 2D/axisymmetric and/or 3D decks for each module considered. The modules are tested using Rocketdyn's proprietary code REACT. The code utilizes an efficient solution procedure to solve Navier-Stokes equations in a non-orthogonal body-fitted coordinate system. The differential equations are discretized over a finite-volume grid using a non-staggered variable arrangement and an efficient solution procedure based on the SIMPLE algorithm for the velocity-pressure coupling is used. The modules developed have been interfaced and tested using finite-volume, pressure-correction CFD solvers which are widely used in the CFD community. Other solvers can also be used to test these modules since they are independently structured with their own discretization scheme and solver methodology. Many of these modules have been independently tested by Professor C.P. Chen and his group at the University of Alabama at Huntsville (UAH) by interfacing them with own flow solver (MAST).
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-203937 , NAS 1.26:203937 , RI/RD96-182 , CDR-DR-3
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    High-Performance Parallel Analysis of Coupled Problems for Aircraft Propulsion (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: This research program dealt with the application of high-performance computing methods to the numerical simulation of complete jet engines. The program was initiated in January 1993 by applying two-dimensional parallel aeroelastic codes to the interior gas flow problem of a bypass jet engine. The fluid mesh generation, domain decomposition and solution capabilities were successfully tested. Attention was then focused on methodology for the partitioned analysis of the interaction of the gas flow with a flexible structure and with the fluid mesh motion driven by these structural displacements. The latter is treated by a ALE technique that models the fluid mesh motion as that of a fictitious mechanical network laid along the edges of near-field fluid elements. New partitioned analysis procedures to treat this coupled three-component problem were developed during 1994 and 1995. These procedures involved delayed corrections and subcycling, and have been successfully tested on several massively parallel computers, including the iPSC-860, Paragon XP/S and the IBM SP2. For the global steady-state axisymmetric analysis of a complete engine we have decided to use the NASA-sponsored ENG10 program, which uses a regular FV-multiblock-grid discretization in conjunction with circumferential averaging to include effects of blade forces, loss, combustor heat addition, blockage, bleeds and convective mixing. A load-balancing preprocessor tor parallel versions of ENG10 was developed. During 1995 and 1996 we developed the capability tor the first full 3D aeroelastic simulation of a multirow engine stage. This capability was tested on the IBM SP2 parallel supercomputer at NASA Ames. Benchmark results were presented at the 1196 Computational Aeroscience meeting.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-203477 , NAS 1.26:203477 , CU-CAS-96-29
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    Wave Engine Topping Cycle Assessment (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The performance benefits derived by topping a gas turbine engine with a wave engine are assessed. The wave engine is a wave rotor that produces shaft power by exploiting gas dynamic energy exchange and flow turning. The wave engine is added to the baseline turboshaft engine while keeping high-pressure-turbine inlet conditions, compressor pressure ratio, engine mass flow rate, and cooling flow fractions fixed. Related work has focused on topping with pressure-exchangers (i.e., wave rotors that provide pressure gain with zero net shaft power output); however, more energy can be added to a wave-engine-topped cycle leading to greater engine specific-power-enhancement The energy addition occurs at a lower pressure in the wave-engine-topped cycle; thus the specific-fuel-consumption-enhancement effected by ideal wave engine topping is slightly lower than that effected by ideal pressure-exchanger topping. At a component level, however, flow turning affords the wave engine a degree-of-freedom relative to the pressure-exchanger that enables a more efficient match with the baseline engine. In some cases, therefore, the SFC-enhancement by wave engine topping is greater than that by pressure-exchanger topping. An ideal wave-rotor-characteristic is used to identify key wave engine design parameters and to contrast the wave engine and pressure-exchanger topping approaches. An aerodynamic design procedure is described in which wave engine design-point performance levels are computed using a one-dimensional wave rotor model. Wave engines using various wave cycles are considered including two-port cycles with on-rotor combustion (valved-combustors) and reverse-flow and through-flow four-port cycles with heat addition in conventional burners. A through-flow wave cycle design with symmetric blading is used to assess engine performance benefits. The wave-engine-topped turboshaft engine produces 16% more power than does a pressure-exchanger-topped engine under the specified topping constraints. Positive and negative aspects of wave engine topping in gas turbine engines are identified.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107371 , NAS 1.15:107371 , AIAA Paper 97-0707 , ARL-TR-1284 , E-10539 , Aerospace Sciences Meeting and Exhibit; Jan 06, 1997 - Jan 10, 1997; Reno, NV; United States
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    Imaging of Combustion Species in a Radially-Staged Gas Turbine Combustor (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Planar laser-induced fluorescence (PLIF) is used to characterize the complex flowfield of a unique fuel-lean, radially-staged, high pressure gas turbine combustor. PLIF images of OH are presented for two fuel injector configurations. PLIF images of NO, the first acquired at these conditions, are presented and compared with gas sample extraction probe measurements. Flow field imaging of nascent C2 chemiluminescence is also investigated. An examination is made of the interaction between adjoining lean premixed prevaporized (LPP) injectors. Fluorescence interferences at conditions approaching 2000 K and 15 atm are observed and attributed to polycyclic aromatic hydrocarbon (PAH) emissions. All images are acquired at a position immediately downstream of the fuel injectors with the combustor burning JP-5 fuel.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107373 , NAS 1.15:107373 , E-10542 , Joint Combustion and Propulsion Systems Hazards Subcommittees Meeting; Nov 04, 1996 - Nov 09, 1996; Monterey, CA; United States
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    Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107388 , NAS 1.15:107388 , ARL-TR-1252 , E-10584 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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    Modern Experimental Techniques in Turbine Engine Testing (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The paper describes application of two modern experimental techniques, thin-film thermocouples and pressure sensitive paint, to measurement in turbine engine components. A growing trend of using computational codes in turbomachinery design and development requires experimental techniques to refocus from overall performance testing to acquisition of detailed data on flow and heat transfer physics to validate these codes for design applications. The discussed experimental techniques satisfy this shift in focus. Both techniques are nonintrusive in practical terms. The thin-film thermocouple technique improves accuracy of surface temperature and heat transfer measurements. The pressure sensitive paint technique supplies areal surface pressure data rather than discrete point values only. The paper summarizes our experience with these techniques and suggests improvements to ease the application of these techniques for future turbomachinery research and code verifications.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107383 , E-10572 , NAS 1.15:107383 , International Congress on Fluid Dynamics and Propulsion; Dec 29, 1996 - Dec 31, 1996; Cairo; Egypt
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    Forced Response Testing of an Axi-Centrifugal Turboshaft Engine (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Dynamic data from tests of a T55-L,712 engine are presented. Engine stall/surge data were analyzed using digital signal processing techniques. In addition, forced response testing (system identification studies) was done at various engine speeds. Forced response testing was done using eight jet ejectors approximately equally circumferentially spaced about the compressor front face. This paper presents some preliminary results for the ground idle (approximately 60% of design speed) point. Brief descriptions of the jet injection system, the test matrix, and analysis techniques used are presented. Results of these analyses indicate a substantial transfer of energy across the compressor first stage at some frequencies and that the ejectors are effective in modifying the local flow conditions in front of the first compressor stage.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107337 , E-10474 , NAS 1.15:107337 , ARL-TR-1151 , AIAA Paper 96-2573 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Advanced Models for Aeroelastic Analysis of Propulsion Systems (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: This report describes an integrated, multidisciplinary simulation capability for aeroelastic analysis and optimization of advanced propulsion systems. This research is intended to improve engine development, acquisition, and maintenance costs. One of the proposed simulations is aeroelasticity of blades, cowls, and struts in an ultra-high bypass fan. These ducted fans are expected to have significant performance, fuel, and noise improvements over existing engines. An interface program was written to use modal information from COBSTAN and NASTRAN blade models in aeroelastic analysis with a single rotation ducted fan aerodynamic code.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-202646 , NAS 1.26:202646
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    Endwall Heat Transfer Measurements in a Transonic Turbine Cascade (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: Turbine blade endwall heat transfer measurements are given for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 x 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for freestream turbulence intensities of 0.25% and 7.0%. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136' of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for computational fluid dynamics (CFD) code and model verification. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107387 , E-10583 , NAS 1.15:107387 , ARL-TR-1253 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Birmingham; United Kingdom
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  • 86
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    Propulsion Selection for 85kft Remotely Piloted Atmospheric Science Aircraft (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: This paper describes how a 3 stage turbocharged gasoline engine was selected to power NASA's atmospheric science unmanned aircraft now under development. The airplane, whose purpose is to fly sampling instruments through targeted regions of the upper atmosphere at the exact location and time (season, time of day) where the most interesting chemistry is taking place, must have a round trip range exceeding 1000 km, carry a payload of about 500 lb to altitudes exceeding 80 kft over the site, and be able to remain above that altitude for at least 30 minutes before returning to base. This is a subsonic aircraft (the aerodynamic heating and shock associated with supersonic flight could easily destroy the chemical species that are being sampled) and it must be constructed so it will operate out of small airfields at primitive remote sites worldwide, under varying climate and weather conditions. Finally it must be low cost, since less than $50 M is available for its development. These requirements put severe constraints on the aircraft design (for example, wing loading in the vicinity of 10 psf) and have in turn limited the propulsion choices to already-existing hardware, or limited adaptations of existing hardware. The only candidate that could emerge under these circumstances was a propeller driven aircraft powered by spark ignited (SI) gasoline engines, whose intake pressurization is accomplished by multiple stages of turbo-charging and intercooling. Fortunately the turbocharged SI powerplant, owing to its rich automotive heritage and earlier intensive aero powerplant development during WWII, enjoys in addition to its potentially low development costs some subtle physical advantages (arising from its near-stochiometric combustion) that may make it smaller and lighter than either a turbine engine or a diesel for these altitudes. Just as fortunately, the NASA/industry team developing this aircraft includes the same people who built multi-stage turbocharged SI powerplants for unmanned military spyplanes in the early 1980's. Now adapting hardware developed for reconaissance at 65-70 kft to the interests of atmospheric science at 80-90 kft, their efforts should yield an aero powerplant that pushes the altitude limits of subsonic air breathing propulsion.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107302 , NAS 1.15:107302 , E-10390 , AUVSI 1996; Jul 16, 1996 - Jul 19, 1996; Orlando, FL; United States
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    Evaluation of Water Injection Effect on NO(x) Formation for a Staged Gas Turbine Combustor (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: NO(x) emission control by water injection on a staged turbine combustor (STC) was modeled using the KIVA-2 code with modification. Water is injected into the rich-burn combustion zone of the combustor by a single nozzle. Parametric study for different water injection patterns was performed. Results show NO(x) emission will decrease after water being injected. Water nozzle location also has significant effect for NO formation and fuel ignition. The chemical kinetic model is also sensitive to the excess water. Through this study, a better understanding of the physics and chemical kinetics is obtained, this will enhance the STC design process.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-204769 , NAS 1.26:204769 , AIAA Paper 96-0706 , Aerospace Sciences Meeting and Exhibit; Jan 15, 1996 - Jan 18, 1996; Reno, NV; United States
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    Continued Development of Expert System Tools for NPSS Engine Diagnostics (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The objectives of this grant were to work with previously developed NPSS (Numerical Propulsion System Simulation) tools and enhance their functionality; explore similar AI systems; and work with the High Performance Computing Communication (HPCC) K-12 program. Activities for this reporting period are briefly summarized and a paper addressing the implementation, monitoring and zooming in a distributed jet engine simulation is included as an attachment.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-CR-202436 , NAS 1.26:202436
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  • 89
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    Experimental investigation of unsteady flows at large incidence angles in a linear oscillating cascade (1996)
    In:  CASI
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    Publication Date: 2019-07-13
    Description: The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low aspect ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies up to 0.8 for out-of-phase oscillations at Mach numbers up to 0.8 and chordal incidence angles of 0 deg and 10 deg. For the 10 deg chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107283 , NAS 1.15:107283 , AIAA Paper 96-2823 , E-10358 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
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    Experimental Investigation of the Performance of a Mach-2.7 Two-dimensional Bifurcated Duct Inlet with 30 Percent Internal Contraction (1996)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: An experimental study was conducted to determine the performance of a two-dimensional, mixed-compression bifurcated duct inlet system designed for a free-stream Mach number of 2.7. Thirty percent of the supersonic area contraction occurred internally. A movable ramp was used to vary the contraction ratio for off-design operation. Boundary layer bleed regions were located on the cowl, centerbody, and sidewall surfaces. There were also provisions for vortex generators on the cowl and centerbody of the subsonic diffuser. Data were obtained over the Mach number range of 2.0 to 2.8 and at angles of yaw from 0 deg. to the maximum value prior to inlet un-start. The test at Mach 2.8 was to obtain data for an over- speed condition. The Reynolds number varied from 2.5 to 2.3 million/ft for Mach numbers above 2.5. At Mach numbers of 2.5 and lower, the Reynolds number was set at 2.5 million/ft. Bleed patterns, vortex generator patterns, and ramp position were varied, and three inlet configurations were selected for more extensive study. Two of these configurations had self-starting capability. The self-starting configuration that was developed produced 89 percent total pressure recovery at the compressor face station with 6.8 percent total bleed. The compressor face distortion was about 16 percent. Vortex generators were extremely effective in re-distributing flow but were not as effective in reducing distortion. Excellent flow symmetry was achieved between the separated halves of the inlet, and twin-duct instability was not observed. The ramp tip shock was steeper than expected. This caused the cowl lip shock to be reflected from the ramp instead of being cancelled at the shoulder. However, peak recovery at the throat was still obtained with the ramp near the design position.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-106728 , NAS 1.16:106728 , E-9099
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  • 91
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    Evolutionary conservation of a genetic pathway of programmed cell death (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 4-11 
    Details
    ISSN: 0730-2312
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: Genetic analysis of programmed cell death in Caenorhabditis elegans has led to the identification of 13 genes that constitute a developmental pathway of programmed cell death. Two of the three key genes in this pathway, ced-9, a cell death suppressor, and ced-3, a cell death inducer, were found to encode proteins that share structural and functional similarities with the mammalian proto-oncogene product Bcl-2 and interleukin-1β converting enzyme, respectively. These results suggest that the genetic pathway of programmed cell death may be evolutionarily conserved from worms to mammals. © 1996 Wiley-Liss, Inc.
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    BCL-2, a novel regulator of apoptosis (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 12-17 
    Details
    ISSN: 0730-2312
    Keywords: bcl-2 gene ; localization ; apoptosis ; antioxidants ; oxidative stress ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: The bcl-2 gene has a unique function among mammalian oncogenes as a negative regulator of apoptosis. Its expression pattern in embryonic and adult tissues is consistent with a role in maintaining in vivo survival of specific cell types.The biochemical function of bcl-2 is unknown, but its localization to mitochondrial and microsomal membranes suggests several possibilities, bcl-2 is protective against oxidative stress in mammalian cells and can be replaced by antioxidants in a factor-deprivation model of apoptosis. These results are consistent with a model of apoptotic death involving oxidative stress in a central pathway.The recent discovery of several bcl-2-related genes, some of which also inhibit apoptosis and others that unexpectedly promote apoptosis, has shed new light on several aspects of bcl-2 action. © 1996 Wiley-Liss, Inc.
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    Regulation of apoptosis by oncogenes (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 33-38 
    Details
    ISSN: 0730-2312
    Keywords: Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: No abstract.
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    BCL-2 family proteins: Regulators of cell death involved in the pathogenesis of cancer and resistance to therapy (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 23-32 
    Details
    ISSN: 0730-2312
    Keywords: BCL-2 gene ; Bcl-2 protein ; homologs ; homo- and heterotypic dimers ; cancer ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: The BCL-2 gene was first discovered because of its involvement in the t(14;18) chromosomal translocations commonly found in lymphomas, which result in deregulation of BCL-2 gene expression and cause inappropriately high levels of Bcl-2 protein production. Expression of the BCL-2 gene can also become altered in human cancers through other mechanisms, including loss of the p53 tumor suppressor which normally functions as a repressor of BCL-2 gene expression in some tissues. Bcl-2 is a blocker of programmed cell death and apoptosis that contributes to neoplastic cell expansion by preventing cell turnover caused by physiological cell death mechanisms, as opposed to accelerating rates of cell division. Overproduction of the Bcl-2 protein also prevents cell death induced by nearly all cytotoxic anticancer drugs and radiation, thus contributing to treatment failures in patients with some types of cancer. Several homologs of Bcl-2 have recently been discovered, some of which function as inhibitors of cell death and others as promoters of apoptosis that oppose the actions of the Bcl-2 protein. Many of these Bcl-2 family proteins can interact through formation of homo- and heterotypic dimers. In addition, several nonhomologous proteins have been identified that bind to Bcl-2 and that can modulate apoptosis. These protein-protein interactions may eventual serve as targets for pharmacologically manipulating the physiological cell death pathway for treatment of cancer and several other diseases. © 1996 Wiley-Liss, Inc.
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    Nuclear import of protein kinases and cyclins (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 61-82 
    Details
    ISSN: 0730-2312
    Keywords: protein kinases ; cyclins ; nuclear import ; NLS ; acidic domains ; cell cycle ; phosphatases ; p34cdc2 ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: Karyophilic and acidic clusters were found in most nonmembrane serine/threonine protein kinases whose primary structure was examined. These karyophilic clusters might mediate the anchoring of the kinase molecules to transporter proteins for their regulated nuclear import and might constitute the nuclear localization signals (NLS) of the kinase molecules. In contrast to protein transcription factors that are exclusively nuclear possessing strong karyophilic peptides composed of at least four arginines (R) and lysines (K) within an hexapeptide flanked by proline and glycine helix-breakers, protein kinases often contain one histidine and three K + R residues; this is proposed to specify a weak NLS structure resulting in the nuclear import of a fraction of the total cytoplasmic kinase molecules as well as in their weak retention in the different ionic strength nuclear environment. Putative NLS peptides in protein kinases may also contain hydrophobic or bulky aromatic amino acids proposed to further diminish their capacity to act as strong NLS. Most kinases lacking karyophilic clusters (c-Mos, v-Mos, sea star MAP, and yeast KIN28, SRA1, SRA3, TPK1, TPK2) also lack acidic clusters, which is in contrast to most kinases containing both acidic and karyophilic peptides; this and the presence of R/K clusters in the transporter proteins supports a role of acidic clusters on kinases in nuclear import. Cyclins B lack karyophilic signals and are proposed to be imported into nuclei via their association with Cdc2. © 1996 Wiley-Liss, Inc.
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    Calphostin C induces tyrosine dephosphorylation/inactivation of protein kinase Fa/GSK-3α in a pathway independent of tumor promoter phorbol ester-mediated down-regulation of protein kinase C (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 121-129 
    Details
    ISSN: 0730-2312
    Keywords: protein kinase FA/GSK-3α ; PKC inhibition ; calphostin C ; down-regulation ; carcinoma dedifferentiation/progression ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: The signal transduction mechanism of protein kinase FA/GSK-3α by tyrosine phosphorylation in A431 cells was investigated using calphostin C as an inhibitor for protein kinase C (PKC). Kinase Fa/GSK-3α could be tyrosine-dephosphorylated and inactivated to ∼ 10% of control in a concentration-dependent manner by 0.1-10 μM calphostin C (IC50, ∼ 1 μM), as demonstrated by immunoprecipitation of kinase Fa/GSK-3α from cell extracts, followed by phosphoamino acid analysis and by immunodetection in an antikinase Fa/GSK-3α immunoprecipitate kinase assay. In sharp contrast, down-regulation of PKC by 0.05 μM calphostin C (IC50, ∼ 0.05 μM for inhibiting PKC in cells) or by tumor promoter phorbol ester TPA was found to have stimulatory effect on the cellular activity of kinase Fa/GSK-3α, when processed under identical conditions. Furthermore, TPA-mediated down-regulation of PKC was found to have no effect on calphostin C-mediated tyrosine dephosphorylation/inactivation of kinase Fa/GSK-3α. Taken together, the results provide initial evidence that the PKC inhibitor calphostin C may induce tyrosine dephosphorylation/inactivation of kinase Fa/GSK-3α in a pathway independent of TPA-mediated down-regulation of PKC, representing a new mode of signal transduction for the regulation of this multisubstrate/multifunctional protein kinase by calphostin C in cells. Since kinase Fa/GSK-3α is a possible carcinoma dedifferentiation/progression-promoting factor, the results further suggest calphostin C as a potential anticancer drug involved in blocking carcinoma dedifferentiation/progression, possibly via inactivation of protein kinase FA/GSK-3α in tumor cells. © 1996 Wiley-Liss, Inc.
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    Inducible expression of cyclin D1 in T-47D human breast cancer cells is sufficient for Cdk2 activation and pRB hyperphosphorylation (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 363-378 
    Details
    ISSN: 0730-2312
    Keywords: cyclin D1 function ; CDK activity ; pRB phosphorylation ; G1 phase ; cell cycle control ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: The sequential transcriptional activation of cyclins, the regulatory subunits of cell cycle specific kinases, regulates progress through the cell cycle. In mitogen-stimulated cells cyclin D1 induction in early G1 is followed by induction of cyclin E, activation of the cyclin-dependent kinase Cdk2, and hyperphosphorylation of the retinoblastoma gene product (pRB) in mid-to-late G1 phase. T-47D breast cancer cells expressing cyclin D1 under the control of a metal-responsive metallothionein promoter were used to determine whether Cdk2 activation and pRB hyperphosphorylation are consequences of cyclin D1 induction. A 4-5-fold increase in cyclin D1 protein abundance was followed by approximately 2-fold increases in cyclin E protein abundance and Cdk2 activity and by hyperphosphorylation of pRB. These responses were apparent ∼ 3 h after the increase in cyclin D1 protein, and ∼ 3 h prior to the entry of cyclin D1-stimulated cells into S phase 12 h after zinc treatment. Cyclin D1 immunoprecipitates contained Cdk4 but no detectable Cdk2 and displayed pRb but not histone H1 kinase activity. Cdk2 activation was therefore likely to be due to increased abundance of cyclin E/Cdk2 complexes rather than formation of active cyclin D1/Cdk2 complexes. The sequence of events following zinc induction of cyclin D1 thus mimicked that following mitogen induction of cyclin D1. These data show that cyclin D1 induction is sufficient for Cdk2 activation and pRB hyperphosphorylation in T-47D human breast cancer cells, providing evidence that cyclin D1 induction is a critical event in G1 phase progression. © 1996 Wiley-Liss, Inc.
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  • 98
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    Enhanced expression of heregulin in c-erb B-2 and c-Ha-ras transformed mouse and human mammary epithelial cells (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 437-446 
    Details
    ISSN: 0730-2312
    Keywords: heregulin ; transformation ; erb B-2 ; c-Ha-ras ; mammary cells ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: Heregulin β1 was found to stimulate the anchorage-dependent, serum-free growth of nontransformed human MCF-10A mammary epithelial cells. Unlike epidermal growth factor, transforming growth factor α, or amphiregulin, heregulin β1 was also able to induce the anchorage-independent growth of MCF-10A cells. In contrast, the anchorage-dependent, serum-free growth of c-Ha-ras or c-erb B-2 transformed MCF-10A cells was unaffected by heregulin β1, whereas heregulin β1 was able to stimulate the anchorage-independent growth of these cells. c-Ha-ras or c-erb B-2 (c-neu) transformed MCF-10A or mouse NOG-8 mammary epithelial cells express elevated levels of 2.5, 5.0, 6.5, 6.8, and 8.5 kb heregulin mRNA transcripts and/or synthesize cell-associated 25, 29, 50, and 115 kDa isoforms of heregulin. Since the MCF-10A cells and transformants also express c-erb B-3, these data suggest that endogenous heregulin might function as an autocrine growth factor for Ha-ras or erb B-2 transformed mammary epithelial cells. © 1996 Wiley-Liss, Inc. This article is a US Government work and, as such, is in the public domain in the United States of America.
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    Ecto-alkaline phosphatase considered as levamisole-sensitive phosphohydrolase at physiological pH range during mineralization in cultured fetal calvaria cells (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 484-494 
    Details
    ISSN: 0730-2312
    Keywords: ecto-enzyme ; ALP inhibitor ; Ca incorporation ; glycosylphosphatidylinositol-anchored proteins ; PI-PLC ; bone differentiation ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: Alkaline phosphatase (ALP) activity expressed on the external surface of cultured fetal rat calvaria cells and its relationship with mineral deposition were investigated under pH physiological conditions. After replacement of culture medium by assay buffer and addition of p-nitrophenyl phosphate (pNPP), the rate of substrate hydrolysis catalyzed by whole cells remained constant for up to seven successive incubations of 10 min and was optimal over the pH range 7.6-8.2. It was decreased by levamisole by a 90% inhibition at 1 mM which was reversible within 10 min, dexamisole having no effect. Values of apparent Km for pNPP were close to 0.1 mM, and inhibition of pNPP hydrolysis by levamisole was uncompetitive (Ki = 45 μM). Phosphatidylinositol-specific phospholipase C (PI-PLC) produced the release into the medium of a p-nitrophenyl phosphatase (pNPPase) sensitive to levamisole at pH 7.8. The released activity whose rate was constant up to 75 min represented after 15 min 60% of the value of ecto-pNPPase activity. After 75 min of PI-PLC treatment the ecto-pNPPase activity remained unchanged despite the 30% decrease in Nonidet P-40-extractable ALP activity. High levels of 45Ca incorporation into cell layers used as index of mineral deposition were decreased by levamisole in a stereospecific manner after 4 h, an effect which was reversed within 4 h after inhibitor removal, in accordance with ecto-pNPPase activity variations. These results evidenced the levamisole-sensitive activity of a glycosylphosphatidylinositol-anchored pNPPase consistent with ALP acting as an ecto-enzyme whose functioning under physiological conditions was correlated to 45Ca incorporation and permit the prediction of the physiological importance of the enzyme dynamic equilibrium at the cell surface in cultured fetal calvaria cells. © 1996 Wiley-Liss, Inc.
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    Alternative splicing of smooth muscle myosin heavy chains and its functional consequences (1996)
    New York, N.Y. : Wiley-Blackwell
    Journal of Cellular Biochemistry 60 (1996), S. 521-528 
    Details
    ISSN: 0730-2312
    Keywords: myosin heavy chains ; smooth muscle ; alternative splicing ; contractility ; myosin light chains ; Life and Medical Sciences ; Cell & Developmental Biology
    Source: Wiley InterScience Backfile Collection 1832-2000
    Topics: Biology , Chemistry and Pharmacology , Medicine
    Notes: The aim of our study was to determine the relation between alternatively spliced myosin heavy chain (MHC) isoforms and the contractility of smooth muscle. The relative amount of MHC with an alternatively spliced insert in the 5′ (amino terminal) domain was determined on the protein level using a peptide-directed antibody (a25K/50K) raised against the inserted sequence (QGPSFAY). Smooth muscle MHC isoforms of both bladder and myometrium but not nonmuscle MHC reacted with a25/50K. Using a quantitative Western-blot approach the amount of 5′-inserted MHC in rat bladder was detected to be about eightfold higher than in normal rat myometrium. The amount of heavy chain with insert was found to be decreased by about 50% in the myometrium of pregnant rats. Although bladder contained significantly more 5′-inserted MHC than myometrium, apparent maximal shortening velocities (Vmax) were comparable, being 0.138 ± 0.012 and 0.114 ± 0.023 muscle length per second of skinned bladder and normal myometrium fibers, respectively. Phosphorylation of myosin light chain 20 induced by maximal Ca2+/calmodulin activation was the same in bladder and myometrial fibers. These results suggest that the amount of 5′-inserted MHC is not necessarily associated with contractile properties of smooth muscle. © 1996 Wiley-Liss, Inc.
    Additional Material: 7 Ill.
    Type of Medium: Electronic Resource
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