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Performance of aircraft gas turbine ball bearings, seals, and lubricants under simulated supersonic transport engine conditions
The research requirements for general aviation aircraft are presented. Emphasis is placed on improving the performance of airfoils, propellers, and engines. Additional requirements are expressed with respect to external noise reduction, internal noise reduction, and exhaust emission control. The requirement for anti-icing developments to create improved flight safety is discussed.
NASA is studying the characteristics of future aircraft fuels produced from either petroleum or nonpetroleum sources such as oil shale or coal. These future hydrocarbon based fuels may have chemical and physical properties that are different from present aviation turbine fuels. This research is aimed at determining what those characteristics may be, how present aircraft and engine components and materials would be affected by fuel specification changes, and what changes in both aircraft and engine design would be required to utilize these future fuels without sacrificing performance, reliability, or safety. This fuels technology program was organized to include both in-house and contract research on the synthesis and characterization of fuels, component evaluations of combustors, turbines, and fuel systems, and, eventually, full-scale engine demonstrations. A review of the various elements of the program and significant results obtained so far are presented.
Aviation related rotary (Wankel-type) engine tests, possible growth directions and relevant developments at Curtiss-Wright have been reviewed. Automotive rotary engines including stratified charge are described and flight test results of rotary aircraft engines are presented. The current 300 HP engine prototype shows basic durability and competitive performance potential. Recent parallel developments have separately confirmed the geometric advantages of the rotary engine for direct injected unthrottled stratified charge. Specific fuel consumption equal to or better than pre- or swirl-chamber diesels, low emission and multi-fuel capability have been shown by rig tests of similar rotary engine.
Results are presented from an experimental program aimed at optimizing the features of an advanced two-stage combustor that currently is being developed as part of the National Aeronautics and Space Administration/Pratt & Whitney Aircraft Energy Efficient Engine program. The combustor is designed to meet stringent goals for performance, emissions, durability and operational characteristics. An overview of the design selection process in light of these goals is discussed. Combustor rig test results, which highlight the techniques used to reduce emissions, development of the pilot and main power zone fuel injection systems, and methods of achieving efficient fuel staging, are presented.
High payoff technologies for a year 2000 regenerative cycle turboprop engine were identified for a 19 passenger commuter aircraft application. A series of engines incorporating eight levels of advanced technologies were studied and their impact on aircraft performance was evaluated. Four advanced technologies are recommended to achieve a potential reduction in fuel burn of 38.3 percent. At $1.00 per gallon fuel price, a potential direct operating cost (DOC) benefit of 12.5 percent is obtained. At $2.00 per gallon, the potential DOC benefit increases to 17.0 percent.
The input for global atmospheric chemistry models was generated for baseline High Speed Civil Transport (HSCT) configurations at Mach 1.6, 2.2, and 3.2. The input is supplied in the form of number of molecules of specific exhaust constituents injected into the atmosphere per year by latitude and by altitude (for 2-D codes). Seven exhaust constituents are currently supplied: NO, NO2, CO, CO2, H2O, SO2, and THC (Trace Hydrocarbons). An eighth input is also supplied, NO(x), the sum of NO and NO2. The number of molecules of a given constituent emitted per year is a function of the total fuel burned by a supersonic fleet and the emission index (EI) of the aircraft engine for the constituent in question. The EIs for an engine are supplied directly by the engine manufacturers. The annual fuel burn of a supersonic fleet is calculated from aircraft performance and economic criteria, both of which are strongly dependent on basic design parameters such as speed and range. The altitude and latitude distribution of the emission is determined based on 10 Intern. Air Transport Assoc. (IATA) regions chosen to define the worldwide route structure for future HSCT operations and the mission flight profiles.
The study was conducted to identify high payoff technologies for year 2000 small gas turbine applications and to provide a technology plan for guiding future research and technology efforts. A regenerative cycle turboprop engine was selected for a 19 passenger commuter aircraft application. A series of engines incorporating eight levels of advanced technologies were studied and their impact on aircraft performance was evaluated. The study indicated a potential reduction in fuel burn of 38.3 percent. At $1.00 per gallon fuel price, a potential DOC benefit of 12.5 percent would be achieved. At $2.00 per gallon, the potential DOC benefit would increase to 17.0 percent. Four advanced technologies are recommended and appropriate research and technology programs were established to reach the year 2000 goals.
Many aircraft employ ducts with centerline curvature or changing cross-sectional shape to join the engine with inlet and exhaust components. S-ducts convey air to the engine compressor from the intake and often decelerate the flow to achieve an acceptable Mach number at the engine compressor by increasing the cross-sectional area downstream. Circular-to-rectangular transition ducts are used on aircraft with rectangular exhaust nozzles to connect the engine and nozzle. To achieve maximum engine performance, the ducts should minimize flow total pressure loss and total pressure distortion at the duct exit. Changes in the curvature of the duct centerline or the duct cross-sectional shape give rise to streamline curvature which causes cross stream pressure gradients. Secondary flows can be caused by deflection of the transverse vorticity component of the boundary layer. This vortex tilting results in counter-rotating vortices. Additionally, the adverse streamwise pressure gradient caused by increasing cross-sectional area can lead to flow separation. Vortex pairs have been observed in the exit planes of both duct types. These vortices are due to secondary flows induced by pressure gradients resulting from streamline curvature. Regions of low total pressure are produced when the vortices convect boundary layer fluid into the main flow. The purpose of the present study is to predict the measured flow field in a diffusing S-duct and a circular-to-rectangular transition duct with a full Navier-Stokes computer program, PARC3D, and to compare the numerical predictions with new detailed experimental measurements. The work was undertaken to extend previous studies and to provide additional CFD validation data needed to help model flows with strong secondary flow and boundary layer separation. The S-duct computation extends the study of Smith et al, and Harloff et al, which concluded that the computation might be improved by using a finer grid and more advanced turbulence models. The present study compares results for both the Baldwin-Lomas and k-epsilon turbulence models and is conducted with a refined grid. For the transition duct, two inlet conditions were considered, the first with straight flow and the second with swirling flow. The first case permits examination of the effects of the geometric transition on the flow field, while the second case includes the rotational flow effect characteristic of a gas turbine engine.
The cost in mean effective pressure of generating air flow in the combustion chambers of single-cylinder compression-ignition engines was determined for the prechamber and the displaced-piston types of combustion chamber. For each type a wide range of air-flow quantities, speeds, and boost pressures was investigated. Supplementary tests were made to determine the effect of lubricating-oil temperature, cooling-water temperature, and compression ratio on the friction mean effective pressure of the single-cylinder test engine. Friction curves are included for two 9-cylinder, radial, compression-ignition aircraft engines. The results indicate that generating the optimum forced air flow increased the motoring losses approximately 5 pounds per square inch mean effective pressure regardless of chamber type or engine speed. With a given type of chamber, the rate of increase in friction mean effective pressure with engine speed is independent of the air-flow speed. The effect of boost pressure on the friction cannot be predicted because the friction was decreased, unchanged, or increased depending on the combustion-chamber type and design details. High compression ratio accounts for approximately 5 pounds per square inch mean effective pressure of the friction of these single-cylinder compression-ignition engines. The single-cylinder test engines used in this investigation had a much higher friction mean effective pressure than conventional aircraft engines or than the 9-cylinder, radial, compression-ignition engines tested so that performance should be compared on an indicated basis.
Circular-rectangular transition ducts are used between engine exhausts and nozzles with rectangular cross sections that are designed for high performance aircraft. NASA Glenn Research Center has made experimental investigations of a series of circular-rectangular transition ducts to provide benchmark flow data for comparison with numerical calculations. These ducts are all designed with superellipse cross sections to facilitate grid generation. In response to this challenge, the three-dimensional RNS3D code has been applied to one of these transition ducts. This particular duct has a length-to-inlet diameter ratio of 1.5 and an exit-plane aspect ratio of 3.0. The inlet Mach number is 0.35. Two GRC experiments and the code were run for this duct without inlet swirl. One GRC experiment and the code were also run with inlet swirl. With no inlet swirl the code was successful in predicting pressures and secondary flow conditions, including a pair of counter-rotating vortices at both sidewalls of the exit plane. All these phenomena have been reported from the two GRC experiments. However, these vortices were suppressed in the one experiment when inlet swirl was used; whereas the RNS3D code still predicted them. The experiment was unable to provide data near the sidewalls, the very region where the vortices were predicted.
Starting about 1920 it becomes easier to track the growth of bearing materials technology. Until 1955, with few exceptions, comparatively little progress was made in this area. AISI 52100 and some carburizing grades (AISI 4320, AISI 9310) were adequate for most applications. The catalyst to quantum advances in high-performance rolling-element bearing steels was the advent of the aircraft gas turbine engine. With improved bearing manufacturing and steel processing together with advanced lubrication technology, the potential improvements in bearing life can be as much as 80 times that attainable in the late 1950s or as much as 400 times that attainable in 1940. This paper summarizes the chemical, metallurgical and physical aspects of bearing steels and their effect on rolling bearing life and reliability. The single most important variable that has significantly increased bearing life and reliability is vacuum processing of bearing steel. Differences between through hardened, case carburized and corrosion resistant steels are discussed. The interrelation of alloy elements and carbides and their effect on bearing life are presented. An equation relating bearing life, steel hardness and temperature is given. Life factors for various steels are suggested and discussed. A relation between compressive residual stress and bearing life is presented. The effects of retained austenite and grain size are discussed.
An overview will be provided on NASA GRC activities in additive manufacturing (AM) technology development for advanced high performance, propulsion component designs in electric motors and turbine engine systems in future electrified aircraft. AM offers significant benefits over conventional manufacturing to fabricate components that are more compact, lighter weight, geometrically complex, innovatively cooled, integrated, multi-material, and multifunctional. For turbine engine applications, silicon carbide (SiC) based composites are being pursued using binder jetting. For electric motor applications, additive processes are being applied toward optimized component designs to include the housing, rotor, and stators with 3-D printing and wire imbedded coils and with direct printed 3-phase coils. The will be an emphasis on capabilities in additive manufacturing.
Battery pack gravimetric energy density is one of the most important, yet often miss-estimated design parameters for sizing all-electric aircraft. Proper accounting for thermal, structural, and operational safety margins are frequently lost when extrapolating performance from the cell level to the aircraft level. This paper summarizes the relevant engineering and certification details needed to better account for the penalties associated when assembling battery packs. The relationship between the cell and pack energy density is not linear, as is often assumed. Furthermore, the relationship varies depending on pack requirements, cell chemistry, and architecture. Parametric, high-fidelity models are used to determine optimal battery pack sizes over a range of conditions to better quantify technology scaling effects.
Battery pack gravimetric energy density is one of the most important, yet often miss- estimated design parameters for sizing all-electric aircraft. Proper accounting for thermal, structural, and operational safety margins are frequently lost when extrapolating performance from the cell level to the aircraft level. This paper summarizes the relevant engineering and certification details needed to better account for the penalties associated when assembling battery packs. The relationship between the cell and pack energy density is not linear, as is often assumed. Furthermore, the relationship varies depending on pack requirements, cell chemistry, and architecture. Parametric, high-fidelity models are used to determine optimal battery pack sizes over a range of conditions to better quantify technology scaling effects.
The neural network and regression methods of NASA Glenn Research Center s COMETBOARDS design optimization testbed were used to generate approximate analysis and design models for a subsonic aircraft operating at Mach 0.85 cruise speed. The analytical model is defined by nine design variables: wing aspect ratio, engine thrust, wing area, sweep angle, chord-thickness ratio, turbine temperature, pressure ratio, bypass ratio, fan pressure; and eight response parameters: weight, landing velocity, takeoff and landing field lengths, approach thrust, overall efficiency, and compressor pressure and temperature. The variables were adjusted to optimally balance the engines to the airframe. The solution strategy included a sensitivity model and the soft analysis model. Researchers generated the sensitivity model by training the approximators to predict an optimum design. The trained neural network predicted all response variables, within 5-percent error. This was reduced to 1 percent by the regression method. The soft analysis model was developed to replace aircraft analysis as the reanalyzer in design optimization. Soft models have been generated for a neural network method, a regression method, and a hybrid method obtained by combining the approximators. The performance of the models is graphed for aircraft weight versus thrust as well as for wing area and turbine temperature. The regression method followed the analytical solution with little error. The neural network exhibited 5-percent maximum error over all parameters. Performance of the hybrid method was intermediate in comparison to the individual approximators. Error in the response variable is smaller than that shown in the figure because of a distortion scale factor. The overall performance of the approximators was considered to be satisfactory because aircraft analysis with NASA Langley Research Center s FLOPS (Flight Optimization System) code is a synthesis of diverse disciplines: weight estimation, aerodynamic analysis, engine cycle analysis, propulsion data interpolation, mission performance, airfield length for landing and takeoff, noise footprint, and others.
This paper explores the design of a hybrid electric aircraft propulsion system that uses a turbofan to power an electric system. In such a system, the gas turbine will experience a loss of power generation as altitude increases, however the electric system will not. This difference results in designs that may over size the electric system at high altitude or under size at low altitude. Two studies are performed within this paper. The first looks at extracting power from the engine for use with electric aircraft propulsion at cruise and the second reviews a design of an engine that uses thrust assist for takeoff. Both studies look at the effects of changing altitude on the amount of power extraction or insertion that can be taken from the turbofan as dictated by operability limits. Results of the paper show that low-pressure compressor surge margin and high-pressure compressor speed can be pushed to unaccepted limits with large scale power extraction or insertion, however these issues can be mitigated by adding power extraction or insertion at off design operating points to compensate. Additionally, the benefits of thrust assist are quantified for this configuration demonstrating a reduction in thrust specific fuel consumption at cruise of over 5%.
This paper explores the design of a hybrid electric aircraft propulsion system that uses a turbofan to power an electric system. In such a system, the gas turbine will experience a loss of power generation as altitude increases, however the electric system will not. This difference results in designs that may over size the electric system at high altitude or under size at low altitude. Two studies are performed within this paper. The first looks at extracting power from the engine for use with electric aircraft propulsion at cruise and the second reviews a design of an engine that uses thrust assist for takeoff. Both studies look at the effects of changing altitude on the amount of power extraction or insertion that can be taken from the turbofan as dictated by operability limits. Results of the paper show that low-pressure compressor surge margin and high-pressure compressor speed can be pushed to unaccepted limits with large scale power extraction or insertion, however these issues can be mitigated by adding power extraction or insertion at off design operating points to compensate. Additionally, the benefits of thrust assist are quantified for this configuration demonstrating a reduction in thrust specific fuel consumption at cruise of over 5%.