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Showing papers on "Turbofan published in 2010"


Journal ArticleDOI
TL;DR: In this paper, the performance of a transonic fan operating within non-uniform inlet flow remains a key concern for the design and operability of a turbofan engine and the results of the unsteady computations agree well with the measurement data.
Abstract: The performance of a transonic fan operating within non-uniform inlet flow remains a key concern for the design and operability of a turbofan engine. This paper applies computational methods to improve the understanding of the interaction between a transonic fan and an inlet total pressure distortion. The test case studied is the NASA rotor 67 stage operating with a total pressure distortion covering a 120-degree sector of the inlet flow-field. Full-annulus, unsteady, three-dimensional CFD has been used to simulate the test rig installation and the full fan assembly operating with inlet distortion. Novel post-processing methods have been applied to extract the fan performance and features of the interaction between the fan and the non-uniform inflow. The results of the unsteady computations agree well with the measurement data. The local operating condition of the fan at different positions around the annulus has been tracked and analysed, and this is shown to be highly dependent on the swirl and mass flow redistribution that the rotor induces ahead of it due to the incoming distortion. The upstream flow effects lead to a variation in work input that determines the distortion pattern seen downstream of the fan stage. In addition, the unsteady computations also reveal more complex flow-features downstream of the fan stage, which arise due to the three-dimensionality of the flow and unsteadiness.Copyright © 2010 by ASME

118 citations


Proceedings ArticleDOI
01 Oct 2010
TL;DR: In this paper, the authors describe the control algorithms and control design process for a generic commercial aircraft engine simulation of a 40,000 lb thrust class, two spool, high bypass ratio turbofan engine.
Abstract: This paper describes the control algorithms and control design process for a generic commercial aircraft engine simulation of a 40,000 lb thrust class, two spool, high bypass ratio turbofan engine. The aircraft engine is a complex nonlinear system designed to operate over an extreme range of environmental conditions, at temperatures from approximately -60 to 120+ F, and at altitudes from below sea level to 40,000 ft, posing multiple control design constraints. The objective of this paper is to provide the reader an overview of the control design process, design considerations, and justifications as to why the particular architecture and limits have been chosen. The controller architecture contains a gain-scheduled Proportional Integral controller along with logic to protect the aircraft engine from exceeding any limits. Simulation results illustrate that the closed loop system meets the Federal Aviation Administration s thrust response requirements

114 citations


Proceedings ArticleDOI
01 Oct 2010
TL;DR: A new high-fidelity simulation of a generic 40,000 pound thrust class commercial turbofan engine with a representative controller with a significant feature not found in other non-proprietary models is the inclusion of transient stall margin debits.
Abstract: A new high-fidelity simulation of a generic 40,000 lb thrust class commercial turbofan engine with a representative controller, known as CMAPSS40k, has been developed. Based on dynamic flight test data of a highly instrumented engine and previous engine simulations developed at NASA Glenn Research Center, this non-proprietary simulation was created especially for use in the development of new engine control strategies. C-MAPSS40k is a highly detailed, component-level engine model written in MATLAB/Simulink (The MathWorks, Inc.). Because the model is built in Simulink, users have the ability to use any of the MATLAB tools for analysis and control system design. The engine components are modeled in C-code, which is then compiled to allow faster-than-real-time execution. The engine controller is based on common industry architecture and techniques to produce realistic closed-loop transient responses while ensuring that no safety or operability limits are violated. A significant feature not found in other non-proprietary models is the inclusion of transient stall margin debits. These debits provide an accurate accounting of the compressor surge margin, which is critical in the design of an engine controller. This paper discusses the development, characteristics, and capabilities of the C-MAPSS40k simulation

112 citations


Posted Content
TL;DR: In this article, the authors analyse scaling patterns in terms of changes in the ratios among product characteristics of 143 designs in civil aircraft and show that two allegedly dominant designs, the piston propeller DC3 and the turbofan Boeing 707, are shown to have triggered a scaling trajectory at the level of the respective firms.
Abstract: Using entropy statistics we analyse scaling patterns in terms of changes in the ratios among product characteristics of 143 designs in civil aircraft. Two allegedly dominant designs, the piston propeller DC3 and the turbofan Boeing 707, are shown to have triggered a scaling trajectory at the level of the respective firms. Along these trajectories different variables have been scaled at different moments in time: this points to the versatility of a dominant design which allows a firm to react to a variety of user needs. Scaling at the level of the industry took off only after subsequently reengineered models were introduced, like the piston propeller Douglas DC4 and the turbofan Boeing 767. The two scaling trajectories in civil aircraft corresponding to the piston propeller and the turbofan paradigm can be compared with a single, less pronounced scaling trajectory in helicopter technology for which we have data during the period 1940-1996. Management and policy implications can be specified in terms of the phases of codification at the firm and the industry level.

104 citations


Journal ArticleDOI
TL;DR: The fine particulate matter (PM) emissions from nine commercial aircraft engine models were determined by plume sampling during the three field campaigns of the Aircraft Particle Emissions Experiment (APEX) as discussed by the authors.

89 citations


Patent
28 Jan 2010
TL;DR: In this article, a variable drive system for a propeller or fan of a gas turbine engine is described, which can be applied to a turboprop or turbofan engine having a gearing between the shaft and propeller.
Abstract: A system and method for variable drive of a propeller or fan of a gas turbine engine. The gas turbine engine has a combustor and a turbine arranged to be driven by a combustion product from the combustor. The variable drive system comprises a primary shaft arranged for transmission of torque from said turbine to the propeller; an electric generator arranged to be driven by said turbine; and an electric motor arranged to be driven by the output of said generator. A clutch is mounted between the propeller and the primary rotor and is operable to mechanically disconnect the shaft from the propeller so that the propeller can be driven by any or any combination of the turbine and/or electric motor. The invention may be applied to a turboprop or turbofan engine having a gearing between the shaft and propeller or fan and may be particularly suited to unmanned aerial vehicle proulsion.

81 citations


Patent
Morris G. Anderson1
10 Sep 2010
TL;DR: In this paper, a mixer for a turbofan engine (100) includes a centerbody (144) and a mixer nozzle (146), which is configured to direct at least a portion of the bypass air to impinge on the center body.
Abstract: A mixer for a turbofan engine (100) includes a centerbody (144) and a mixer nozzle (146). The mixer nozzle (146) surrounds at least a portion of the centerbody (144) and is spaced apart to define a core flow path between the mixer nozzle (146) and the centerbody (144). The mixer nozzle (146) is configured, when bypass air flows through the turbofan engine (100), to direct at least a portion of the bypass air to impinge on the centerbody (144). The mixer nozzle (146) includes a plurality of circumferentially spaced mixer lobes (206) that extend axially in a rearward direction and have a cross-section shape defined by a set of equations.

66 citations


Patent
13 May 2010
TL;DR: In this paper, an integrated, single piece mixer-center body ventilation apparatus for use with a turbofan jet engine is described. But the work is limited to the case of a single-stage turbojet.
Abstract: An integrated, single piece mixer-center body ventilation apparatus is disclosed for use with a turbofan jet engine. The apparatus may incorporate a circumferential forward body portion adapted to be coupled to an aft end of a core engine turbine case of the jet engine, and a center body tube portion integrally formed with the forward body portion and having an axially opening vent exit. The forward body portion may have a plurality of inner mixer flow paths in communication with scalloped projecting portions. The inner mixer flow paths direct a pressurized core exhaust flow through the mixer device and mix the pressurized core exhaust flow with a portion of a pressurized fan exhaust flow, to thus significantly cool the pressurized core exhaust flow.

66 citations


Patent
26 Feb 2010
TL;DR: A turbine engine component of a turbofan engine fitted with a bypass air valve (20) includes at least one turbine engine components having a surface (31) with at least 1 aperture (37) and a flow transfer location (34) comprising an area proximate to a turbine exhaust stream flow (28) as discussed by the authors.
Abstract: A turbine engine component of a turbofan engine (10) fitted with a bypass air valve (20) includes at least one turbine engine component having a surface (31) with at least one aperture (37), said turbine engine component located from between a bypass fan duct (32) and a turbine exhaust nozzle (24) of the turbofan engine (10) The bypass air valve (20) includes a liner concentrically disposed about the turbine engine component and parallel to a centerline (30) of the turbofan engine (10) The liner has a surface including at least one aperture (42) and at least one impermeable region (44) Means are provided for actuating the liner about the turbine engine components The flow transfer location (34) comprises an area proximate to a turbine exhaust stream flow (28)

64 citations


Journal ArticleDOI
TL;DR: In this paper, an integrated noise and performance assessment methodology for advanced propfan powered aircraft configurations is presented. The approach is based on first principles and combines a coupled aircraft and propulsion system mission and performance analysis tool with 3-D unsteady, full wheel CRP CFD computations and aero-acoustic simulations.
Abstract: Due to their inherent noise challenge and potential for significant reductions in fuel burn, counter-rotating propfans (CRPs) are currently being investigated as potential alternatives to high-bypass turbofan engines. This paper introduces an integrated noise and performance assessment methodology for advanced propfan powered aircraft configurations. The approach is based on first principles and combines a coupled aircraft and propulsion system mission and performance analysis tool with 3-D unsteady, full wheel CRP CFD computations and aero-acoustic simulations. Special emphasis is put on computing CRP noise due to interaction tones. The method is capable of dealing with parametric studies and exploring noise reduction technologies. An aircraft performance, weight and balance and mission analysis was first conducted on a candidate CRP powered aircraft configuration. Guided by data available in the literature, a detailed aerodynamic design of a pusher CRP was carried out. Full wheel unsteady 3-D RANS simulations were then used to determine the time varying blade surface pressures and unsteady flow features necessary to define the acoustic source terms. A frequency domain approach based on Goldstein’s formulation of the acoustic analogy for moving media and Hanson’s single rotor noise method were extended to counter-rotating configurations. The far field noise predictions were compared to measured data of a similar CRP configuration and demonstrated good agreement between the computed and measured interaction tones. The underlying noise mechanisms have previously been described in the literature but, to the authors’ knowledge, this is the first time that the individual contributions of front-rotor wake interaction, aft-rotor upstream influence, hub-endwall secondary flows and front-rotor tip-vortices to interaction tone noise are dissected and quantified. Based on this investigation, the CRP was re-designed for reduced noise incorporating a clipped rear-rotor and increased rotor-rotor spacing to reduce upstream influence, tip-vortex, and wake interaction effects. Maintaining the thrust and propulsive efficiency at takeoff conditions, the noise was calculated for both designs. At the interaction tone frequencies, the re-designed CRP demonstrated an average reduction of 7.25 dB in mean SPL computed over the forward and aft polar angle arcs. On the engine/aircraft system level, the re-designed CRP demonstrated a reduction of 9.2 EPNdB and 8.6 EPNdB at the FAR 36 flyover and sideline observer locations, respectively. The results suggest that advanced open rotor designs can possibly meet Stage 4 noise requirements.Copyright © 2010 by ASME

41 citations


Journal ArticleDOI
TL;DR: In this paper, the authors present a computational study, validated by mean-flow experiments, of a dual-stream nozzle simulating the exit conditions of a supersonic turbofan engine with noise-suppressing fan flow deflectors.
Abstract: We present a computational study, validated by mean-flow experiments, of a dual-stream nozzle simulating the exit conditions of a supersonic turbofan engine with noise-suppressing fan flow deflectors. The study is conducted for eight nozzle configurations and two operating conditions: a cold condition at which mean velocity surveys were conducted and against which the computational code was validated and a hot condition that corresponds to the takeoff engine cycle and at which acoustic data were collected. The code predictions successfully replicate the mean velocity fields and the inflectional layers of the experimental flows. The code is then extended to the conditions of the actual engine cycle. The computations reveal a similar velocity profile for the hot and cold conditions when the axial distance is normalized by the potential core length. For both conditions, the vane deflectors reduce the turbulent kinetic energy k on the underside of the jet. An overall noise source strength is modeled as the axial integral of k 7/2 . A significant correlation is found between the reduction in the noise source strength and the reduction in the peak level of the overall sound pressure level.

Patent
25 Jan 2010
TL;DR: In this article, a system and method for controlling the temperature of engine bleed air from a turbofan gas turbine engine is presented, including a fan air valve and a controller.
Abstract: A system and method are provided for controlling the temperature of engine bleed air from a turbofan gas turbine engine. The system includes a fan air valve and a fan air valve controller. The fan air valve is adapted to receive a flow of fan air from a turbofan gas turbine engine intake fan. The fan air valve is coupled to receive valve position commands and is configured, in response to the valve position commands, to move to a valve position to thereby control the engine bleed air temperature. The fan air valve controller is configured to implement a linear quadratic regulator (LQR) control. The fan air valve controller is adapted to receive a plurality of sensor signals, each sensor signal representative of one or more system parameters, and is configured, in response to the sensor signals, to supply the valve position commands to the fan air valve.

Journal ArticleDOI
TL;DR: In this article, the performance of an intercooled turbofan engine is analyzed by multidisciplinary optimization in terms of the relevant constraints such as compressor exit blade height, and the results indicate that a state-of-the-art high pressure compressor efficiency can be achieved.
Abstract: The performance of an intercooled turbofan engine is analysed by multidisciplinary optimization. A model for making preliminary simplified analysis of the mechanical design of the engine is coupled to an aircraft model and an engine performance model. A conventional turbofan engine with technology representative for a year 2013 entry of service is compared with a corresponding intercooled engine. A mission fuel burn reduction of 3.4% is observed. The results are analysed in terms of the relevant constraints such as compressor exit blade height. It is shown that the gas path of an intercooled engine for medium range commericial transport applications, having an overall pressure ratio greater than 70 in top of climb, may still be optimized to fulfil a compressor exit blade height constraint. This indicates that a state of the art high pressure compressor efficiency can be achieved. Empirical data and a parametric CFD study is used to verify the intercooler heat transfer and pressure loss characteristics.

Journal ArticleDOI
TL;DR: In this article, a novel simulation theory for a complete fixed-wing aircraft is presented, which addresses a shortfall in multi-disciplinary integration in aircraft flight, including economic operations, preliminary design and environmental emissions.
Abstract: This contribution presents a novel simulation theory for a complete fixed-wing aircraft. Novel methods are presented for flight mechanics (fuel planning), turbofan engine simulation (in direct and inverse mode), thermo-physics integration (tire temperature on the ground and fuel temperature in flight) and aircraft noise. At the fundamental level, the framework presented addresses a shortfall in multi-disciplinary integration in aircraft flight, including economic operations, preliminary design and environmental emissions. Validation strategies are introduced for component-level analysis and system integration. Results are presented for geometry models, specific air range and optimal cruise conditions, payload-range performance, fuel temperature of a wing tank, tire heating during normal take-off, aircraft propulsive (jet/nozzle) and non propulsive (landing gear) noise. Selected results are shown for the Boeing B777-300 and the Airbus A380-861.

Patent
06 Dec 2010
TL;DR: In this paper, a turbofan gas turbine engine with a unique power off-take shaft and gear system is described, where a drive gear is provided near the front end of the high pressure shaft.
Abstract: A turbofan gas turbine engine is provided having a unique power off-take shaft and gear system. Other gas turbine engine types are also contemplated herein. Two power off-takes are provided, one each for the low pressure spool and high pressure spool. The power off-takes extend across a core flow path of the turbofan engine between the low and high pressure shafts to a fan frame of the turbofan. A drive gear is provided near the front end of the high pressure shaft, and another drive gear is provided on the low pressure shaft near the drive gear for the high pressure shaft. Both gears are located in a sump of the gas turbine engine. The power off-take shafts are coupled to the drive gears. Two power devices are coupled to the power off-take shafts and are located in the fan frame. The power devices can be electric generators or motors.

Journal ArticleDOI
TL;DR: In this article, an aerodynamic model for a large blended-wing-body transport aircraft with blown flap effects was formulated using empirical and vortex lattice methods and then integrated with a Trent 500 turbofan engine model.
Abstract: While operating at low airspeeds with nominal static margins, the controls on a blended-wing-body aircraft begin to saturate, and the dynamic performance gets sluggish. Augmentation of aerodynamic controls with the propulsion system is therefore considered in this research. Two aspects were of interest: namely, thrust vectoring and flap blowing. An aerodynamic model for a large blended-wing-body transport aircraft with blown flap effects was formulated using empirical and vortex lattice methods and then integrated with a Trent 500 turbofan engine model. To enhance control effectiveness, both internally and externally blown flaps were simulated. For a full-span internally blown flap arrangement using intermediate compressor flow, the amount of engine bleed and the resulting blowing coefficients were limited. However, even with a reduced bleed mass flow, the pitch control effectiveness increases by 15.9% at 85% fan revolutions per minute. For an externally blown flap arrangement using bypass air, much higher blowing coefficients can be achieved. For instance, at 100% fan revolutions per minute, there is a 44% increase in pitch control authority at low dynamic pressures. The main benefit occurs during takeoff, where both thrust vectoring and flap blowing help in achieving early pitch rotation, reducing takeoff field length and liftoff speed considerably. With central flap blowing and a limited thrust vectoring of 10°, the liftoff range reduces by 48%, and liftoff speed reduces by almost 26%.


Proceedings ArticleDOI
22 Dec 2010
TL;DR: The Numerical Propulsion System Simulation (NPSS) as mentioned in this paper is an object-oriented framework allowing the gas turbine engine analyst considerable flexibility in cycle conceptual design and performance estimation.
Abstract: The Numerical Propulsion System Simulation (NPSS) code was created through a joint United States industry and National Aeronautics and Space Administration (NASA) effort to develop a state-of-the-art aircraft engine cycle analysis simulation tool. Written in the computer language C++, NPSS is an object-oriented framework allowing the gas turbine engine analyst considerable flexibility in cycle conceptual design and performance estimation. Furthermore, the tool was written with the assumption that most users would desire to easily add their own unique objects and calculations without the burden of modifying the source code. The purpose of this paper is twofold: first, to present an introduction to the discipline of thermodynamic cycle analysis to those who may have some basic knowledge in the individual areas of fluid flow, gas dynamics, thermodynamics, and turbomachinery theory but not necessarily how they are collectively used in engine cycle analysis. Second, this paper will show examples of performance modeling of gas turbine engine cycles specifically using Numerical Propulsion System Simulation concepts and model syntax. Current practices in industry and academia will also be discussed. While NPSS allows both steady-state and transient simulations and is written to facilitate higher orders of analysis fidelity, the pedagogical example will focus primarily on steady-state analysis of an aircraft mixed flow turbofan at the 0-D and 1-D level. Ultimately it is hoped that this paper will provide a starting point by which both the novice cycle analyst and the experienced engineer looking to transition to a superior tool can use NPSS to analyze any kind of practical gas turbine engine cycle in detail.

01 Jun 2010
TL;DR: In this article, an integrated noise and performance assessment methodology for advanced propfan powered aircraft configurations is presented. The approach is based on first principles and combines a coupled aircraft and propulsion system mission and performance analysis tool with 3-D unsteady, full wheel CRP CFD computations and aero-acoustic simulations.
Abstract: Due to their inherent noise challenge and potential for significant reductions in fuel burn, counter-rotating propfans (CRPs) are currently being investigated as potential alternatives to high-bypass turbofan engines. This paper introduces an integrated noise and performance assessment methodology for advanced propfan powered aircraft configurations. The approach is based on first principles and combines a coupled aircraft and propulsion system mission and performance analysis tool with 3-D unsteady, full wheel CRP CFD computations and aero-acoustic simulations. Special emphasis is put on computing CRP noise due to interaction tones. The method is capable of dealing with parametric studies and exploring noise reduction technologies. An aircraft performance, weight and balance and mission analysis was first conducted on a candidate CRP powered aircraft configuration. Guided by data available in the literature, a detailed aerodynamic design of a pusher CRP was carried out. Full wheel unsteady 3-D RANS simulations were then used to determine the time varying blade surface pressures and unsteady flow features necessary to define the acoustic source terms. A frequency domain approach based on Goldstein’s formulation of the acoustic analogy for moving media and Hanson’s single rotor noise method were extended to counter-rotating configurations. The far field noise predictions were compared to measured data of a similar CRP configuration and demonstrated good agreement between the computed and measured interaction tones. The underlying noise mechanisms have previously been described in the literature but, to the authors’ knowledge, this is the first time that the individual contributions of front-rotor wake interaction, aft-rotor upstream influence, hub-endwall secondary flows and front-rotor tip-vortices to interaction tone noise are dissected and quantified. Based on this investigation, the CRP was re-designed for reduced noise incorporating a clipped rear-rotor and increased rotor-rotor spacing to reduce upstream influence, tip-vortex, and wake interaction effects. Maintaining the thrust and propulsive efficiency at takeoff conditions, the noise was calculated for both designs. At the interaction tone frequencies, the re-designed CRP demonstrated an average reduction of 7.25 dB in mean SPL computed over the forward and aft polar angle arcs. On the engine/aircraft system level, the re-designed CRP demonstrated a reduction of 9.2 EPNdB and 8.6 EPNdB at the FAR 36 flyover and sideline observer locations, respectively. The results suggest that advanced open rotor designs can possibly meet Stage 4 noise requirements.Copyright © 2010 by ASME

Patent
29 Sep 2010
TL;DR: The cross sectional flow area of a fan discharge nozzle on one side of a central plane of an associated gas turbine engine power plant is greater than the corresponding flow area on an opposite side of the central plane to compensate for the blockage of fan airflow by a pylon as mentioned in this paper.
Abstract: The cross sectional flow area of a fan discharge nozzle on one side of a central plane of an associated gas turbine engine power plant is greater than the corresponding flow area of the fan discharge nozzle on an opposite side of the central plane to compensate for the blockage of fan airflow by a pylon.

Patent
30 Jun 2010
TL;DR: In this paper, a variable area nozzle (26) was used to regulate the core flow by positioning the variable area to meet the thrust demand of a turbofan. But the controller was not able to control the gas path temperature of core flow.
Abstract: A control system (40) for a turbofan (10) comprises a variable area nozzle (26) for regulating core flow (Fc) through the turbofan, an actuator (27), a temperature sensor (42), a flight controller (44) and a nozzle control (47). The actuator (27) is coupled to the variable area nozzle (26) to regulate the core flow by positioning the variable area nozzle. The temperature sensor (42) is positioned in the turbofan to sense a gas path temperature of the core flow (Fc). The flight controller (44) is connected to the turbofan to make a thrust demand based on a flight condition of the turbofan (10). The nozzle control (47) is connected to the flight controller (44) and the actuator for directing the actuator based on the gas path temperature and the flight condition, such that the gas path temperature is controlled by adjusting the variable area nozzle (26) to regulate the core flow (Fc) while the turbofan meets the thrust demand.

01 Sep 2010
TL;DR: In this paper, the thermodynamic cycles of dry and reheated turbojets as well as turbofans are examined at supersonic flight Mach numbers, and point performance calculations are done for altitude/Mach number combinations on a line in the middle of a typical flight envelope with constant equivalent airspeed EAS.
Abstract: : High speed propulsion employing turbojets, turbofans and variable cycle engines is interpreted here as propulsion for supersonic air vehicles with flight Mach numbers up to the technical limits of the gas turbine. This limit is somewhere between flight Mach numbers of 3 to 4. If the mission asks for higher vehicle speeds then other propulsion concepts need to be considered, eventually in combination with gas turbines dedicated to take off, acceleration and the return segments of the mission. First the thermodynamic cycles of dry and reheated turbojets as well as turbofans are examined at supersonic flight Mach numbers. All point performance calculations are done for altitude/Mach number combinations on a line in the middle of a typical flight envelope with constant equivalent airspeed EAS. For the flight condition of Mach 2 at 11km altitude it is shown that for a given thrust the size of a dry turbofan is significantly bigger than that of an engine with afterburner. However, all the engines must not only be able to operate at their supersonic design condition but also at all the combinations of Mach number and altitude on the flight path from take off to maximum speed. This off-design requirement influences the selection of the aerodynamic compressor design point and consequently also the size of the turbomachines. A short section about variable cycle engines explains with an example how such a machine operates with the various settings of flow diverter valves, mixer and nozzle area. It is shown that in addition to these adjustable geometry elements the core driven fan stage needs variable inlet guide vanes. Finally two components that are not found in engines designed for subsonic flight are described in some detail with examples: the afterburner and the variable area convergent divergent nozzle.

Patent
20 Dec 2010
TL;DR: In this paper, a method and tooling for partial disassembly of a bypass turbofan engine where the longitudinal axis of the bypass turbine remains generally horizontal during disassembly is presented.
Abstract: A method and tooling for partial disassembly of a bypass turbofan engine wherein the longitudinal axis of the bypass turbofan engine remains generally horizontal during disassembly. The low pressure turbine module is removed with a low pressure turbine module horizontal removal tool. An extended bearing nut tool may be supported by a stabilization member and may remove a bearing nut. An extended high pressure turbine shaft stretching tool may stretch a high pressure turbine shaft to release a high pressure turbine shaft nut. An extended bearing pulling tool may be used to pull a bearing while the low pressure turbine shaft remains in place. A modified measurement bridge may be used to measure the position of certain components while the low pressure turbine shaft remains in place. A nozzle jig may be used to assemble nozzles and feather seals to create a nozzle module. And an arcuate datum may be used to make certain measurements from the aft end of the high pressure turbine shaft while the low pressure turbine shaft remains in place.

Proceedings ArticleDOI
13 Sep 2010
TL;DR: The purpose of these NASA UHB engine concept studies is to determine if the fuel consumption and noise benefits of engines having lower fan pressure ratios translate into overall aircraft system-level benefits for a 737 class vehicle.
Abstract: Considerable interest surrounds the design of the next generation of single-aisle commercial transports in the Boeing 737 and Airbus A320 class. Aircraft designers will depend on advanced, next-generation turbofan engines to power these airplanes. The focus of this study is to apply singleand multi-objective optimization algorithms to the conceptual design of ultrahigh bypass (UHB) turbofan engines for this class of aircraft, using NASA’s Subsonic Fixed Wing Project goals as multidisciplinary objectives for optimization. The independent propulsion design parameters investigated are aerodynamic design point fan pressure ratio, overall pressure ratio, fan drive system architecture (i.e., director geardriven), bypass nozzle architecture (i.e., fixedor variable-geometry), and the highand lowpressure compressor work split. NASA Project goal metrics – fuel burn, noise, and emissions – are among the parameters treated as dependent objective functions. These optimized solutions provide insight to the UHB engine design process and provide independent information to NASA program management to help guide its technology development efforts. This assessment leverages results from earlier NASA system concept studies conducted in 2008 and 2009, in which UHB turbofans were examined for a notional, nextgeneration, single-aisle transport. The purpose of these NASA UHB engine concept studies is to determine if the fuel consumption and noise benefits of engines having lower fan pressure ratios (and correspondingly higher bypass ratios) translate into overall aircraft system-level benefits for a 737 class vehicle.

Proceedings ArticleDOI
10 Oct 2010
TL;DR: In this article, the potential benefits of introducing heat-exchanged cores in future turbofan engine designs were investigated using a multidisciplinary design tool, TERA2020, which comprises of various modules covering a wide range of disciplines: engine performance, engine aerodynamic and mechanical design, aircraft design and performance, emissions prediction and environmental impact, engine and airframe noise, as well as production, maintenance and direct operating costs.
Abstract: Reduction of CO2 emissions is strongly linked with the improvement of engine specific fuel consumption, as well as the reduction of engine nacelle drag and weight. Conventional turbofan designs however that reduce CO2 emissions — such as increased OPR designs — can increase the production of NOx emissions. In the present work, funded by the European Framework 6 collaborative project NEWAC, an aero engine multidisciplinary design tool, TERA2020, has been utilised to study the potential benefits from introducing heat-exchanged cores in future turbofan engine designs. The tool comprises of various modules covering a wide range of disciplines: engine performance, engine aerodynamic and mechanical design, aircraft design and performance, emissions prediction and environmental impact, engine and airframe noise, as well as production, maintenance and direct operating costs. Fundamental performance differences between heat-exchanged cores and a conventional core are discussed and quantified. Cycle limitations imposed by mechanical considerations, operational limitations and emissions legislation are also discussed. The research work presented in this paper concludes with a full assessment at aircraft system level that reveals the significant potential performance benefits for the inter-cooled and intercooled recuperated cycles. An intercooled core can be designed for a significantly higher OPR and with reduced cooling air requirements, providing a higher thermal efficiency than could otherwise be practically achieved with a conventional core. Variable geometry can be implemented to optimise the use of the intercooler for a given flight mission. An intercooled recuperated core can provide high thermal efficiency at low OPR values and also benefit significantly from the introduction of a variable geometry low pressure turbine. The necessity of introducing novel lean-burn combustion technology, to reduce NOx emissions, at cruise as well as for the landing and take-off cycle, is demonstrated for both heat-exchanged cores and conventional designs. Significant benefits in terms of NOx reduction are predicted from the introduction of a variable geometry low pressure turbine in an intercooled core with lean-burn combustion technology.Copyright © 2010 by ASME

Journal ArticleDOI
TL;DR: In this paper, a computational viewpoint on the problems of design and numerical simulation for the nozzles of modern aircraft turbofan engines is presented, including simulation of near and far field of a nozzle, for generation of input perturbations and for processing the far-field noise.

Proceedings ArticleDOI
10 Oct 2010
TL;DR: In this paper, the application of the compliant foil bearing to turbofan engines was examined for the 120 kN (approx. 25000 lb) thrust class. But the application was limited to the military turboprocessor.
Abstract: Over the past several years the term oil-free turbomachinery has been used to describe a rotor support system for high speed turbomachinery that does not require oil for lubrication, damping, or cooling. The foundation technology for oil-free turbomachinery is the compliant foil bearing. This technology can replace the conventional rolling element bearings found in current engines. Two major benefits are realized with this technology. The primary benefit is the elimination of the oil lubrication system, accessory gearbox, tower shaft, and one turbine frame. These components account for 8 to 13 percent of the turbofan engine weight. The second benefit that compliant foil bearings offer to turbofan engines is the capability to operate at higher rotational speeds and shaft diameters. While traditional rolling element bearings have diminished life, reliability, and load capacity with increasing speeds, the foil bearing has a load capacity proportional to speed. The traditional applications for foil bearings have been in small, lightweight machines. However, recent advancements in the design and manufacturing of foil bearings have increased their potential size. An analysis, grounded in experimentally proven operation, is performed to assess the scalability of the modern foil bearing. This analysis was coupled to the requirements of civilian turbofan engines. The application of the foil bearing to larger, high bypass ratio engines nominally at the 120 kN (approx.25000 lb) thrust class has been examined. The application of this advanced technology to this system was found to reduce mission fuel burn by 3.05 percent.

Journal ArticleDOI
TL;DR: In this paper, the application of Computational Aero-Acoustics (CAA) to the prediction of acoustic propagation in turbofan ducts is discussed and evidence is presented to indicate that current CAA is able to represent the effect of in-duct liners on far-field measured rig and engine data.

01 Mar 2010
TL;DR: In this paper, a 48-microphone planar phased array system was used to acquire noise source localization data on a full-scale Williams International FJ44 turbofan engine, where data were acquired with the array at three different locations relative to the engine.
Abstract: A 48-microphone planar phased array system was used to acquire noise source localization data on a full-scale Williams International FJ44 turbofan engine. Data were acquired with the array at three different locations relative to the engine, two on the side and one in front of the engine. At the two side locations the planar microphone array was parallel to the engine centerline; at the front location the array was perpendicular to the engine centerline. At each of the three locations, data were acquired at eleven different engine operating conditions ranging from engine idle to maximum (take off) speed. Data obtained with the array off to the side of the engine were spatially filtered to separate the inlet and nozzle noise. Tones occurring in the inlet and nozzle spectra were traced to the low and high speed spools within the engine. The phased array data indicate that the Inflow Control Device (ICD) used during this test was not acoustically transparent; instead, some of the noise emanating from the inlet reflected off of the inlet lip of the ICD. This reflection is a source of error for far field noise measurements made during the test. The data also indicate that a total temperature rake in the inlet of the engine is a source of fan noise.

Patent
15 Nov 2010
TL;DR: A fan cowl support for an aircraft having an engine pylon, an engine fan case, and an engine cowl is described in this paper, where at least a portion of the engine fan is connected to the support.
Abstract: A fan cowl support for an aircraft having an engine pylon, an engine fan case and an engine fan cowl. The fan cowl support includes a support having a forward end connected to the engine fan case and an aft end connected to the engine pylon. At least a portion of the engine fan cowl is connected to the support.