Showing papers on "Turbofan published in 2000"

Acoustics and Trust of Separate-Flow Exhaust Nozzles With Mixing Devices for High-Bypass-Ratio Engines

Abstract: The NASA Glenn Research Center recently completed an experimental study to reduce the jet noise from modern turbofan engines. The study concentrated on exhaust nozzle designs for high-bypass-ratio engines. These designs modified the core and fan nozzles individually and simultaneously. Several designs provided an ideal jet noise reduction of over 2.5 EPNdB for the effective perceived noise level (EPNL) metric. Noise data, after correcting for takeoff thrust losses, indicated over a 2.0-EPNdB reduction for nine designs. Individually modifying the fan nozzle did not provide attractive EPNL reductions. Designs in which only the core nozzle was modified provided greater EPNL reductions. Designs in which core and fan nozzles were modified simultaneously provided the greatest EPNL reduction. The best nozzle design had a 2.7-EPNdB reduction (corrected for takeoff thrust loss) with a 0.06-point cruise thrust loss. This design simultaneously employed chevrons on the core and fan nozzles. In comparison with chevrons, tabs appeared to be an inefficient method for reducing jet noise. Data trends indicate that the sum of the thrust losses from individually modifying core and fan nozzles did not generally equal the thrust loss from modifying them simultaneously. Flow blockage from tabs did not scale directly with cruise thrust loss and the interaction between fan flow and the core nozzle seemed to strongly affect noise and cruise performance. Finally, the nozzle configuration candidates for full-scale engine demonstrations are identified.

Translating independently mounted air inlet system for aircraft turbofan jet engine

Abstract: An arrangement for an aircraft propulsion system to determinably reduce or eliminate inlet cowl induced bending of a turbofan aircraft jet engine case during certain flight conditions. The arrangement includes a turbofan aircraft jet engine having an engine case with a fan case portion, a nacelle structure housing the engine and having a forward nacelle portion and a rearward nacelle portion, and a pylon structure to support the engine and the nacelle structure while permitting the forward nacelle structure to be translated forwardly from an operational position to a servicing position to permit access to the engine and its components. The forward nacelle portion is independently supported from the pylon and has an inner skin portion that sealingly engages the fan case portion of the engine case and an outer skin portion that is determinably locked to the rearward nacelle portion. The sealing arrangement of the forward nacelle portion and the fan case portion is designed to determinably tailor the load transfer between the forward nacelle portion and the fan case portion to completely eliminate or to reduce such load transfer according to the particular propulsion system arrangement.

Turbofan engine control design using robust multivariable control technologies

Abstract: A unified robust multivariable approach to propulsion control design has been developed at NASA Glenn Research Center. The critical elements of this unified approach are: a robust H/sub /spl infin// control synthesis formulation; a simplified controller scheduling scheme; and a new approach to the synthesis of integrator windup protection gains for multivariable controllers. This paper presents results from an application of these technologies to control design for linear models of an advanced turbofan engine. The objectives of the study were to transfer technology to industry and to identify areas of further development for the technology. The technology elements and industrial development of tools to implement the steps are described with respect to their application to a GE variable-cycle turbofan engine. A set of three-input/three-output three-state linear engine models was used over a range of power levels covering engine operation from idle to maximum unaugmented power. Results from simulation evaluation are discussed and insight is provided into how the design parameter choices affect the results.

Auxiliary power and thrust unit

Abstract: An improvement to an aircraft is provided in the form of an aircraft auxiliary power and thrust unit located in the tail cone of the aircraft. The unit includes a turbofan engine, an air intake opening, an inlet duct extending between the air intake opening and the turbofan engine, a transmission assembly, and various auxiliary equipment. The engine includes a forward-facing main turbine shaft. The air intake opening is located in the tail cone at a body station location forward of the engine. The transmission assembly includes a drive shaft mounted axially to the main turbine shaft and extends forward through the inlet duct through a sealed opening in the inlet duct. The auxiliary equipment is also located in the tail cone, forward of the turbofan engine. The transmission assembly is releasably connected to the auxiliary equipment. In a first operating mode, the engine is operated at a low setting to power the auxiliary equipment. In a second operating mode, the turbofan engine is used to provide thrust and operate auxiliary equipment.

Turbojet and turbofan engine performance increases through turbine burners

Abstract: In a conventional turbojet and turbofan engine, fuel is burned in the main combustor before the heated highpressure gas expands through the turbine. A turbine-burner concept was proposed in a previous paper in which combustion is continued inside the turbine to increase the efe ciency and speciec thrust of the turbojet engine. This concept is extended to include not only continuous burning in the turbine but also “discrete” interstage turbine burners as an intermediate option. A thermodynamic cycle analysis is performed to compare the relative performances of the conventional engineand theturbine-burner engine with differentcombustion options for both turbojet and turbofan conegurations. Turbine-burner engines are shown to provide signie cantly higher specie c thrust with no or only small increases in thrust specie c fuel consumption compared to conventional engines. Turbine-burner engines also widen the operational range of e ight Mach number and compressor pressure ratio. The performance gain of turbine-burner engines over conventional engines increases with compressor pressure ratio, fan bypass ratio, and eight Mach number.

Development of Methods for Analysis and Optimization of Complex Jet Engine Systems

Abstract: This thesis describes the development of GESTPAN (GEneral Stationary and Transient Propulsion ANalysis), a generalized system for the design, steady-state and transient simulation of gas turbine systems. Some of the main achievements in the thesis are related to the development of new algorithms or integration of existing numerics tailored to simplify the structure and use of generalized gas turbine simulation systems. In particular, a method for performing system design utilizing the analysis equations, i.e. an inverse design method, has been developed. Furthermore, attention is drawn to a number of advantages of using an implicit high order differential algebraic system solver for transient gas turbine system analysis. The simulation studies carried out with the GESTPAN system have focused on the performance optimization of the Selective Bleed variable cycle engine. In particular, a method for controlling the engine during mode transition was developed. Work with the implementation of a hybridized optimization method suitable for mission optimization of variable cycle engines is also described. The method couples the cycle selection and the control optimization of the engine variable geometry. Simulations performed with the method indicate that previously published designs of the Selective Bleed Variable cycle engine can be downsized considerably. Early work carried out in the research project concentrated on developing a method for optimizing the performance of variable geometry compressors integrated in gas turbine systems. Although the method was limited to subsonic operation of compressors, it was successfully used to simulate the core driven fan stage of the double bypass variable cycle engine.

Some Aspects of Modelling Compressor Behavior in Gas Turbine Performance Calculations

Abstract: Performance calculation procedures for gas turbine engines are usually based on the performance characteristics of the engine components, and especially the turbo components are of major interest. In this paper methods of modelling compressors in gas turbine performance calculations are discussed. The basic methodologies based on Mach number similarity are summarized briefly including some second order effects. Under extreme engine partload conditions, as for example subidle or windmilling, the operating points in the compressor map are located in a region which is usually not covered by rig tests. In addition the parameters usually used in compressor maps are no longer appropriate. For these operating conditions a method is presented to extrapolate compressor maps towards very low spool speed down to the locked rotor. Instead of the efficiency more appropriate parameters as for example specific work or specific torque are suggested. A compressor map prepared with the proposed methods is presented and discussed. As another relevant topic the performance modelling of fans for low bypass ratio turbofans is covered. Due to the flow splitter downstream of such a fan the core and bypass stream may be throttled independently during engine operation and bypass ratio becomes a third independent parameter in the map. Because testing a fan on the rig for various bypass ratios is a very costly task, a simplified method has been developed which accounts for the effects of bypass ratio.

System identification of jet engines

Abstract: System identification plays an important role in advanced control systems for jet engines, in which controls are performed adaptively using data from the actual engine and the identified engine. An identification technique for jet engine using the Constant Gain Extended Kalman Filter (CGEKF) is described. The filter is constructed for a two-spool turbofan engine. The CGEKF filter developed here can recognize parameter change in engine components and estimate unmeasurable variables over whole flight conditions. These capabilities are useful for an advanced Full Authority Digital Electric Control (FADEC). Effects of measurement noise and bias, effects of operating point and unpredicted performance change are discussed. Some experimental results using the actual engine are shown to evaluate the effectiveness of CGEKF filter.

The optimization of gas turbine engine design by genetic algorithms1

Abstract: The main aim of this paper is the development of an integrated solver for the optimization of the cycle, the size and the weight of a gas turbine engine. The solver is based on the Genetic Algorithm techniques. The selected engine configuration is a two spool, separated flow turbofan with a low pressure compressor linked to the fan. The problem is to find the values of some parameters of engine cycle optimizing some engine performance and geometric parameters. The considered cycle parameters are: the by pass ratio, the presure ratio of fan, low and high pressure compressor and the turbine inlet temperature. The engine parameters are the specific thrust, the Specific Fuel Consumption (SFC), the cold and hot specific exhaust area (A8/m0 e Ai9/m0), the engine lenght, the thrust to weight ratio, etc. The paper shows some examples of solver application by carrying out both single target and multi target optimization processes.

Introduction of Intercooling in a High Bypass Jet Engine

Abstract: In this paper an exercise to introduce intercooling in a high bypass civil turbofan is outlined. The engine selected as the basic propulsion system is a three spool high bypass turbofan with a bypass ratio 6.4. The air leaving the IP compressor is cooled in the bypass duct prior to entering the HP compressor.This preliminary investigation appears to indicate that the main benefit to be gained is an increase in the net thrust from the engine without increasing the turbine inlet temperature. To keep engine diameter constant, the bypass ratio has not been changed. This results in a requirement to significantly increase the pressure ratio to reduce the SFC levels to an acceptable value.A sizing exercise has been carried out to understand the weight and volume penalties imposed by heat pipe intercooling hardware. The preliminary sizing exercise indicates that the weight penalty is very large. The performance of the aircraft using the intercooled engines is also investigated and some improvements in performance are predicted.Overall this investigation is considered to be positive so that further investigations should be considered. It appears that an intercooled engine can produce a somewhat higher thrust at a given turbine entry temperature at similar SFC levels of current engines, or, if a small increase in SFC is acceptable, the increase in thrust is quite important.Copyright © 2000 by ASME

Aerodynamic Test Results of Controlled Pressure Ratio Engine (COPE) Dual Spool Air Turbine Rotating Rig

Abstract: The Controlled Pressure Ratio Engine (COPE) is a fourth generation variable cycle engine combining the attributes of a high temperature turbojet (high dry specific thrust and low Max power SFC) with those of a turbofan (low specific thrust and low part power SFC). Variation in turbine flow function is achieved by the Controlled Area Turbine (CAT) Nozzle concept, which utilizes an innovative cam driven scheme to achieve desired flow function changes while minimizing loss in aerodynamic performance. The single stage high pressure turbine is coupled with a two stage vaneless counter-rotating low pressure turbine. The COPE Turbine System Aero/Heat Transfer Design Validation Program, jointly conducted by GE Aircraft Engines and Allison Advanced Development Company under the direction of the Air Force Research Laboratory at Wright-Patterson Air Force Base, has succeeded in demonstrating advanced turbine technologies that will be utilized on the XTE76, XTE77, and Joint Strike Fighter engines. The various phases of this program evaluated variable area nozzle performance, high pressure turbine performance under the influence of varying flow function, and dual spool testing of the vaneless, counter-rotating low pressure turbine. Evaluation of the three phases demonstrated the aerodynamic capability of these turbine technologies, meeting pre-test predictions in overall and component efficiencies.Copyright © 2000 by ASME

Switchably coupled turbofan engines for high speed aircraft

Abstract: A jet engine for supersonic aircraft which enables supersonic flight and high fuel efficiency to be achieved while suppressing noise at take-off. The engine comprises a front engine consisting of a turbofan engine, and a rear engine or engines consisting of a turbofan or turbojet engine disposed to the rear of said front engine, the rear engine(s) being coupled switchably by a tube to the bypass duct of the front engine, the air inlet(s) of the rear engine(s) being coupled to the bypass duct during supersonic flight, or during acceleration to supersonic speed, whereby the pressure and temperature of the bypass air from the front engine is raised by the rear engine(s), and the air inlet(s) of the rear engine(s) being separated from the bypass duct during take-off, in such a manner that the bypass air is discharged without the temperature or pressure thereof being raised further.

Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications - Summary of the NASA Langley Innovative Thrust Reverser Test Program

Abstract: The NASA Langley Configuration Aerodynamics Branch has conducted an experimental investigation to study the static performance of innovative thrust reverser concepts applicable to high-bypass-ratio turbofan engines. Testing was conducted on a conventional separate-flow exhaust system configuration, a conventional cascade thrust reverser configuration, and six innovative thrust reverser configurations. The innovative thrust reverser configurations consisted of a cascade thrust reverser with porous fan-duct blocker, a blockerless thrust reverser, two core-mounted target thrust reversers, a multi-door crocodile thrust reverser, and a wing-mounted thrust reverser. Each of the innovative thrust reverser concepts offer potential weight savings and/or design simplifications over a conventional cascade thrust reverser design. Testing was conducted in the Jet-Exit Test Facility at NASA Langley Research Center using a 7.9 percent-scale exhaust system model with a fan-to-core bypass ratio of approximately 9.0. All tests were conducted with no external flow and cold, high-pressure air was used to simulate core and fan exhaust flows. Results show that the innovative thrust reverser concepts achieved thrust reverser performance levels which, when taking into account the potential for system simplification and reduced weight, may make them competitive with, or potentially more cost effective than current state-of-the-art thrust reverser systems.

Acoustics and Thrust of Separate Flow Exhaust Nozzles With Mixing Devices Investigated for High Bypass Ratio Engines

Abstract: Typical installed separate-flow exhaust nozzle system. The jet noise from modern turbofan engines is a major contributor to the overall noise from commercial aircraft. Many of these engines use separate nozzles for exhausting core and fan streams. As a part of NASA s Advanced Subsonic Technology (AST) program, the NASA Glenn Research Center at Lewis Field led an experimental investigation using model-scale nozzles in Glenn s Aero-Acoustic Propulsion Laboratory. The goal of the investigation was to develop technology for reducing the jet noise by 3 EPNdB. Teams of engineers from Glenn, the NASA Langley Research Center, Pratt & Whitney, United Technologies Research Corporation, the Boeing Company, GE Aircraft Engines, Allison Engine Company, and Aero Systems Engineering contributed to the planning and implementation of the test.

Steady Performance Measurements of a Turbofan Engine With Inlet Distortions Containing Co- and Counter-Rotating Swirl From an Intake Diffuser for Hypersonic Flight

Abstract: The influence of distorted inlet flow on the steady and unsteady performance of a turbofan engine, which is a component of an air-breathing combined propulsion system for a hypersonic transport aircraft, is reported in this paper. The performance and stability of this propulsion system depend on the behavior of the turbofan engine. The complex shape of the intake duct causes inhomogeneous flow at the engine inlet plane, where total pressure and swirl distortions are present. The S-bend intakes are installed axisymmetrically left and right into the hypersonic aircraft, generating axisymmetric mirror-inverted flow patterns. Since all turbo engines of the propulsion system have the same direction of rotation, one distortion corresponds to a corotating swirl at the low pressure compressor (LPC) inlet while the mirror-inverted image counterpart represents a counterrotating swirl. Therefore the influence of the distortions on the performance and stability of the CO' and COUNTER' rotating turbo engine are different. The distortions were generated separately by an appropriate simulator at the inlet plane of a LARZAC 04 engine. The results of low-frequency measurements at different engine planes yield the relative variations of thrust and specific fuel consumption and hence the steady engine performance. High-frequency measurements were used to investigate the different influence of CO and COUNTER inlet distortions on the development of LPC instabilities.

Noise Reduction Potential of Large, Over-the-Wing Mounted, Advanced Turbofan Engines

Abstract: As we look to the future, increasingly stringent civilian aviation noise regulations will require the design and manufacture of extremely quiet commercial aircraft. Indeed, the noise goal for NASA's Aeronautics Enterprise calls for technologies that will help to provide a 20 EPNdB reduction relative to today's levels by the year 2022. Further, the large fan diameters of modem, increasingly higher bypass ratio engines pose a significant packaging and aircraft installation challenge. One design approach that addresses both of these challenges is to mount the engines above the wing. In addition to allowing the performance trend towards large, ultra high bypass ratio cycles to continue, this over-the-wing design is believed to offer noise shielding benefits to observers on the ground. This paper describes the analytical certification noise predictions of a notional, long haul, commercial quadjet transport with advanced, high bypass engines mounted above the wing.

Active Flow Control For Reduction of Unsteady Stator-Rotor Interaction In a Turbofan Simulator

Abstract: ii

Full 3D Analysis of the GE90 Turbofan Primary Flowpath

Abstract: The multistage simulations of the GE90 turbofan primary flowpath components have been performed. The multistage CFD code, APNASA, has been used to analyze the fan, fan OGV and booster, the 10-stage high-pressure compressor and the entire turbine system of the GE90 turbofan engine. The code has two levels of parallel, and for the 18 blade row full turbine simulation has 87.3 percent parallel efficiency with 121 processors on an SGI ORIGIN. Grid generation is accomplished with the multistage Average Passage Grid Generator, APG. Results for each component are shown which compare favorably with test data.

Identifying Optimal Fan Compressor Pressure Ratios for the Mixed-Stream Turbofan Engine

Abstract: This paper describes a heuristic approach for determining the optimal fan compressor pressure ratio for a e xed set of mixed-stream turbofan engine parameters. During the aircraft design process, it is important to select an enginedesign thatminimizes fuel consumption while producing the thrust required by the variousaircraftmission e ight conditions. Although the most fuel-efe cient values of such parameters as bypass ratio and high-pressure compressor (core) pressure ratio depend on the integrated effects of the various e ight altitudes, Mach numbers, and thrusts required throughout the mission, it is possible to heuristically locate the most efe cient fan compressor pressure ratio to complement other engine parameters independent of the mission. At this optimal value, thrustspecie c fuel consumption is approximately minimized and specie c thrust is approximately maximized at all e ight conditions. Exploiting the observed engine performance characteristics, optimal fan compressor pressure ratios were successfully located for a wide range of bypass ratios and core compressor pressure ratios across a variety of on- and off-design e ight conditions.

Advanced Subsonic Technology (AST) Separate-Flow High-Bypass Ratio Nozzle Noise Reduction Program Test Report

Abstract: NASA s model-scale nozzle noise tests show that it is possible to achieve a 3 EPNdB jet noise reduction with inwardfacing chevrons and flipper-tabs installed on the primary nozzle and fan nozzle chevrons. These chevrons and tabs are simple devices and are easy to be incorporated into existing short duct separate-flow nonmixed nozzle exhaust systems. However, these devices are expected to cause some small amount of thrust loss relative to the axisymmetric baseline nozzle system. Thus, it is important to have these devices further tested in a calibrated nozzle performance test facility to quantify the thrust performances of these devices. The choice of chevrons or tabs for jet noise suppression would most likely be based on the results of thrust loss performance tests to be conducted by Aero System Engineering (ASE) Inc. It is anticipated that the most promising concepts identified from this program will be validated in full scale engine tests at both Pratt & Whitney and Allied-Signal, under funding from NASA s Engine Validation of Noise Reduction Concepts (EVNRC) programs. This will bring the technology readiness level to the point where the jet noise suppression concepts could be incorporated with high confidence into either new or existing turbofan engines having short-duct, separate-flow nacelles.

The Emergence of the Turbofan Engine

Abstract: If you have looked out the window of an airplane lately, you may have noticed that jet engines are gradually getting shorter and fatter. You will see 737s, the most common airliner in service, with two types of engines of distinctly different shapes. The older models have long, stovepipe-shaped engines under the wings, where the newer ones (or older ones which have been retrofitted with new engines) have rounder, shorter powerplants, with a large shell or nacelle around the outside and a smaller cylinder protruding from the rear. Boeing’s latest, the 777, has relatively short but immense engines — each with diameter equivalent to the fuselage of the 737. This change represents the maturing of the turbofan engine, which in the early 1960s superseded the older turbojet engine. Strictly speaking, for the past thirty-five years we have been living in the fan age more than the jet age.

Investigation of Inlet Guide Vane Wakes in a F109 Turbofan Engine with and without Flow Control

Abstract: ii Investigation of Inlet Guide Vane Wakes in a F109 Turbofan Engine with and without Flow Control by Jeffrey D. Kozak Dr. W.F. Ng, Chairman Department of Mechanical Engineering Virginia Tech, 2000 (ABSTRACT) A series of experiments were conducted in a F109 turbofan engine to investigate the unsteady wake profiles of an Inlet Guide Vane (IGV) at a typical spacing to the downstream fan at subsonic and transonic relative blade velocities. The sharp trailingedge vanes were designed to produce a wake profile consistent with modern IGV. Time averaged baseline measurements were first performed with the IGV located upstream of the aerodynamic influence of the fan. Unsteady experiments were performed with an IGV-fan spacing of 0.43 fan chords. High-frequency on-vane pressure measurements showed strong peak-to-peak amplitudes at the blade passing frequency (BPF) of 4.7 psi at the transonic fan speeds. High-frequency total pressure measurements of the IGV wake were taken between the IGV and fan. Results showed that the total pressure loss coefficient of the time averaged IGV wake is reduced by 30% for the subsonic fan, and increased by a factor of 2 for the transonic fan compared to the baseline. Time resolved wake profiles for subsonic fan speeds show constructive and destructive interactions over each blade pass generated by the fan potential flow field. Time resolved wake profiles for the transonic fan speeds show that shock interactions with the IGV surface result in the wake shedding off of the vane at the BPF. Furthermore, the effectiveness of trailing edge blowing (TEB) flow control was investigated. TEB is the method of injecting air aft of the IGV to reduce the low pressure regions (deficits) in the viscous wakes shed by the vanes. Minimizing the IGV wakes reduces the forcing function on the downstream fan

Jet propulsion engine

Abstract: A jet engine comprises a front turbofan engine 2, and at least one rear turbofan or turbojet engine 4. The rear engine(s) may be coupled by a connecting tube 3, to the bypass duct 5, of the front engine. Means are provided for switching between a first state in which the air inlet(s) of the rear engine(s) are coupled to the bypass duct (during supersonic flight or during acceleration to supersonic speed) and a second state in which the air inlet(s) of the rear engine(s) are disconnected from the bypass duct (during take-off). The means for switching between the two states may involve rotational, or linear transverse, displacement of the rear engine(s) (figs. 1-3, 6-7), rotation of the connecting tube (figures 4 and 5), or linear transverse displacement of the front engine (fig.8).

Turbine boosted fan

Abstract: The utility model relates to a turbine boosted blower used for a smoke exhauster, which is composed of a motor and a motor frame Both ends of an output shaft of the motor is fixedly provided with a turbofan and a turbine boosted fan which adopts secondary turbine with multiple blades and large torsion angle The utility model adopts the working principle of an aviation turbine fan and solves the problem of relationship between noise and air quantity in the smoke exhauster, enlarging air quantity without increasing noise The secondary turbine ensures wind pressure that the smoke exhauster needs, so the utility model makes smoke exhausters provided with the turbine boosted blower satisfy three indexes of smoke exhausters