Showing papers on "Turbofan published in 1996"

Variable fan exhaust area nozzle for aircraft gas turbine engine with thrust reverser

Abstract: A structure to provide a variable nozzle for the fan air flow of a turbofan aircraft gas turbine engine with the nozzle being defined by the exit throat area defined by the aft edge of a trailing edge portion of the fan cowl and the core engine housing. A trailing edge cowl portion is slidably positioned in a rearwardly opening annular cavity in the aft end of the fan cowl for reciprocal axial translation between a deployed position to provide an enlarged exit throat area to provide enhanced performance of the aircraft engine during take off and climb to a cruising altitude and a stowed position to provide an optimum exit throat area for cruise condition of the engine. The trailing edge portion is sealably secured within the core cowl cavity to preclude any air leakage from the fan air stream of the fan duct around the trailing edge portion. A unique dual slider and track arrangement is provided to permit the variable nozzle to be translated aft simultaneously with and/or independently of a translating cowl portion of the fan cowl which may form a portion of a blocker door/cascade thrust reverser structure. The trailing edge portion includes sound attenuation treatment on its inner surface for exposure which when deployed enhances the sound attenuation capability of the nacelle system. The variable nozzle may also be used to advantage with other thrust reverser systems such as a four pivoting door structure positioned in the fan cowl.

Operation of Gas Turbine Engines in Volcanic Ash Clouds

Abstract: Results are reported for a technology program designed to determine the behavior of gas turbine engines when operating in particle-laden clouds. There are several ways that such clouds may be created, i.e., explosive volcanic eruption, sand storm, military conflict, etc. The response of several different engines, among them the Pratt & Whitney JT3D turbofan, the Pratt & Whitney J57 turbojet, a Pratt & Whitney engine of the JT9 vintage, and an engine of the General Electric CF6 vintage has been determined. The particular damage mode that will be dominant when an engine experiences a dust cloud depends upon the particular engine (the turbine inlet temperature at which the engine is operating when it encounters the dust cloud), the concentration of foreign material in the cloud, and the constituents of the foreign material (the respective melting temperature of the various constituents). Further, the rate at which engine damage will occur depends upon all of the factors given above, and the damage is cumulative with continued exposure. An important part of the Calspan effort has been to identify environmental warning signs and to determine which of the engine parameters available for monitoring by the flight crew can provide an early indication of impending difficulty. On the basis of current knowledge, if one knows the location of a particle-laden cloud, then that region should be avoided. However, if the cloud location is unknown, which is generally the case, then it is important to know how to recognize when an encounter has occurred and to understand how to operate safely, which is another part of the Calspan effort.

Turbofan engine with reduced noise

Abstract: A noise reduction kit for modifying a two (2) spool axial flow turbofan engine with multi-stage compressors and fan driven by multi-stage reaction turbines, and a thrust of at least about 18,000 lbs. at sea level. There is a fan at the upstream end of the core engine for generating axial fan air flow through bypass ducts terminating at a common nozzle, the common nozzle having a mixing plane area for each of the fan air flow and for the exhaust gas in a range between 700 and 800 square inches. A target thrust reverser includes opposing doors rotatable into position to block and divert the flow of exhaust gas for generating reverse thrust. Mixing means for radially diverting fan air and permitting radially outward expansion of exhaust gas is provided coaxially downstream to the core engine. An acoustically dampened light bulb-shaped nose cone is provided for coaxial attachment to an upstream end of the core engine. The axial flow front fan is axially separated from the inlet guide vane is extended relatively forwardly. A bleed air valve, selectively operational to bleed air from the core engine when the valve is open, is ducted for directing bleed air to the vicinity of the common nozzle.

Turbofan engine with a low pressure turbine driven supercharger in a bypass duct operated by a fuel rich combustor and an afterburner

Abstract: A multiple bypass turbofan engine includes a core Brayton Cycle gas generator with a fuel rich burning combustor and is provided with a variable supercharged bypass duct around the gas generator with a supercharging means in the supercharged bypass duct powered by a turbine not mechanically connected to the gas generator. The engine further includes a low pressure turbine driven forward fan upstream and forward of an aft fan and drivingly connected to a low pressure turbine by a low pressure shaft, the low pressure turbine being aft of and in serial flow communication with the core gas generator. A fan bypass duct is disposed radially outward of the core engine assembly and has first and second inlets disposed between the forward and aft fans. An inlet duct having an annular duct wall is disposed radially inward of the bypass duct and connects the second inlet to the bypass duct. A supercharger means for compressing air is drivingly connected to the low pressure turbine and is disposed in the inlet duct. A secondary combustor or augmentor is disposed in an exhaust duct downstream of and in fluid flow communication with the bypass duct and the gas generator.

Improved NASA-ANOPP Noise Prediction Computer Code for Advanced Subsonic Propulsion Systems

Abstract: 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.

Two-stage mixer ejector suppressor

Abstract: A two-stage mixer ejector concept ("TSMEC") is disclosed for suppressing the noise emanated by jet aircraft. This TSMEC was designed to help older engines meet the stringent new federal noise regulations, known as "Stage III". In an illustrated embodiment, the TSMEC comprises: a lobed engine nozzle attached to the rear of a turbofan engine; a short shroud that straddles the exit end of the engine nozzle; two rings of convergent/divergent primary and secondary lobes within the engine nozzle shroud and; and a ring of arcuate gaps that precede the shroud and the second nozzle ring. The lobes are complimentary shaped to rapidly mix the turbine's hot exhaust flow with cooler air including entrained ambient air, at supersonic speed. This drastically reduces the jet exhaust's velocity and increases jet mixing, thereby reducing the jet noise.

Performance Benefits for Wave Rotor-Topped Gas Turbine Engines

Abstract: 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.Copyright © 1996 by ASME

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.

Variable pressure and variable air flow turbofan engines

Abstract: A variable cycle gas turbine engine has, in axial flow serial relationship, a fan assembly, a core engine, and an exhaust assembly. The core engine includes a compressor, a combustor, a high pressure turbine and a first shaft interconnecting the compressor and the high pressure turbine. The fan assembly is connected to a low pressure turbine by a second shaft which is coaxial with the first shaft. The fan assembly includes a plurality of variable pitch inlet guide vanes, variable pitch stator vanes, and variable pitch outlet guide vanes, the pitch of which are adjusted to regulate the airflow through the gas turbine engine and the pressure ratio across the fan assembly as controlled by the exhaust assembly.

Turbofan with thrust reversal doors not submitted to bypass air in their inactive position

Abstract: A turbofan, preferentially with separated-flows or short nacelle, comprises two thrust reversal doors normally housed in casings formed in the nacelle and located totally outside the inner enclosure of the nacelle. When a thrust reversal effect is sought, the doors are directed in the prolongation of the fan channel, behind the rear end of the nacelle. Thus, the doors are not submitted to the engine bypass air of the jet engine when the aircraft is in flight and the inner enclosure of the nacelle presents neither leak nor discontinuity and may be totally anti-noise treated.

Nacelle acoustic design by ray tracing in three dimensions

Abstract: High frequency, broadband inlet radiation accounts for a major portion of the community noise of transport aircraft with high bypass ratio engines. A ray tracing technique has been developed to model the propagation of this noise from an assumed source location (an annular region on the fan face) to the far field. The procedure is 3-dimensional, accounts for the detailed distribution of acoustic lining, and computes diffracted waves directly. Flow effects are neglected, except for some boundary layer influences. The method has been validated in several model and full scale engine tests. One of these tests was a new acoustic design for a production high bypass-ratio turbofan inlet This inlet was designed using the ray tracing method. The technique shows that acoustic lining is critically important in certain areas of the inlet and unimportant in others. One critical area is the lip region on the side of the inlet opposite to the direction of interest (the crown for flyover noise). Acoustic lining on the bottom of the inlet is not predicted to be important for high frequency, broadband noise. The method also shows that the scarf design is quite effective.

Engine noise suppression ejector nozzle

Abstract: A moderately high bypass ratio turbofan engine nozzle (36) is provided including an outer structure (46) and one or more ejectors (38). The ejectors (38) include inner and outer doors (82), (80) for closing off an ejector passage (76) extending through the outer structure (46). The ejectors (38) are sized to entrain exterior air (40) at aspiration ratios of generally less than 60%. The exterior air (40) is mixed with engine exhaust (42), resulting in a lower combined airflow velocity which in turn reduces jet exhaust noise. Mixing components formed as turning vanes (108) are located in the ejector passage (76) for encouraging further mixing of the airflows (40), (42). In preferred embodiments, the nozzle (36) further includes a translatable centerbody (52), an aft flap assembly (112), and a control system (122) for controlling the movements of the nozzle components. A method of suppressing aircraft moderately high bypass ratio turbofan engine exhaust noise including entraining exterior air to engine exhaust at an aspiration ratio of less than about 60%.

Noise reduction kit for turbofan engine

Abstract: In a noise reduction kit for modifying a two spool axial flow turbofan engine (20) with multi-stage compressors and fan driven by multi-stage reaction turbines, and a thrust of at least about 18,000 lbs. at sea level, there is a fan (21,22) at the upstream end (66) of the core engine (20) for generating axial fan air flow through bypass ducts (26a,26b,26c;27a,27b,27c) terminating at a common nozzle (28), the common nozzle (28) having a mixing plane area (29) for each of the fan air flow and for the exhaust gas in a range between 700 and 800 square inches. A target thrust reverser (41) includes opposing doors (43,44) rotatable into position to block and divert the flow of exhaust gas for generating reverse thrust. A material layer (300) is located in a spacing between a tip of blades (302) for at least some of the fans (21) and a duct for the fans, thereby to reduce a normal clearance between the tip of blades (302) for a fan (21) and the duct (305). Mixing means for radially diverting fan air and permitting radially outward expansion of exhaust gas is provided coaxially downstream to the core engine (20). An acoustically dampened light bulb-shaped nose cone (51) is provided for coaxial attachment to an upstream end of the core engine (20). The axial flow front fan is axially separated from the inlet guide vane which is extended relatively forwardle. A bleed air valve (202), selectively operational to bleed air from the core engine (20) when the valve (202) is open, is ducted for directing bleed air to the vicinity of the common nozzle (28).

Multi-spool turbofan engine with turbine bleed

Abstract: A multi-spool turbofan engine has a plurality of circumferentially spaced poppet valves with diverters secured thereto for precisely controlling bleed of combustion gas aft of the high pressure turbine whereby the high pressure spool operates at high idle RPM so as to power accessories and the low pressure spool operates at low RPM so as to minimize noise and fuel consumption.

Thrust reverser for a turbofan jet engine

Abstract: A thrust reverser for a turbofan-type turbojet engine in which the axial length of the fan housing is less than that of the turbojet engine cowling such that the movable thrust reverser baffles may be attached to the jet engine cowling and moved to an extended position wherein the cold flow gases emanating from the cold flow exhaust duct are redirected to provide a thrust reversing force. The baffles are movably attached to the engine cowling by a plurality of linkrods forming a deformable parallelogram kinematic linkage. Each of the linkrods has opposite ends pivotally attached to the movable baffle and to the engine cowling, respectively. An actuator is connected between the movable baffle and the engine cowling to move the baffle between an extended position, wherein it redirects the cold flow gases to provide a thrust reversing force, and a retracted position wherein the outer surface of the movable baffle is substantially flush with the outer surface of the engine cowling so as to minimize disturbances of the cold flow gases passing over the surface of the cowling. The actuator may be either located forwardly of the baffle or rearwardly of the baffle.

Health Monitoring of Turbine Engine Gas Path Components and Measurement Instruments

Abstract: A performance simulation program has been used to simulate deteriorated performance of a new augmented turbofan engine developed for fighter aircraft and generate a fault pattern library.This fault pattern library forms a basis for both understanding degradation trends of engine usage and in developing an engine health monitoring system. An efficient health monitoring method is proposed to identify the engine faults along with measurement uncertainties and faulty instruments, and to reduce false alarms. A pattern matching method is used to discriminate the engine faults by matching the measurement patterns throughout the fault pattern library.The comparison of this approach to conventional gas path analysis has demonstrated that this approach has comparable ability to monitor engine gas path performance degradation, and provides some capability to handle measurement uncertainties and faults. It also provides a good base for future capability in conjunction with other engine inspection and/or monitoring methods.Copyright © 1996 by ASME

Turbofan engine with a floating pod

Abstract: In a turbofan engine, the mechanical connection between the pod (12) and the fan stator case (10) is limited to a centering system (22) in the immediate vicinity of the arms (20) by which the fan stator case is mounted on the central portion (16) of the engine. The aerodynamic and inertia loads applied to the pod (12), particularly on take-off and landing, consequently only lead to an acceptable deformation of the fan stator case (10). Thus, the gap between said case and the fan (18) can be kept to a very small value, which ensures maximum engine efficiency. Moreover, the centering system (22) prevents an excessive radial displacement (J') between the front of the fan stator case (10) and the adjacent portion of the pod.

Active Noise Control of Low Speed Fan Rotor-Stator Modes

Abstract: This report describes the Active Noise Cancellation System designed by General Electric and tested in the NASA Lewis Research Center's 48 inch Active Noise Control Fan. The goal of this study was to assess the feasibility of using wall mounted secondary acoustic sources and sensors within the duct of a high bypass turbofan aircraft engine for active noise cancellation of fan tones. The control system is based on a modal control approach. A known acoustic mode propagating in the fan duct is cancelled using an array of flush-mounted compact sound sources. Controller inputs are signals from a shaft encoder and a microphone array which senses the residual acoustic mode in the duct. The canceling modal signal is generated by a modal controller. The key results are that the (6,0) mode was completely eliminated at 920 Hz and substantially reduced elsewhere. The total tone power was reduced 9.4 dB. Farfield 2BPF SPL reductions of 13 dB were obtained. The (4,0) and (4,1) modes were reduced simultaneously yielding a 15 dB modal PWL decrease. Global attenuation of PWL was obtained using an actuator and sensor system totally contained within the duct.

Mounting for fixing turbofan engine on airframe

Abstract: The airframe (40) has three mounting points (44,46,48) arranged at the apices of a first triangle. The turbofan casing (18) has three mounting points (50,52,54) arranged at the apices of a second triangle. Supports (36,38) connect the turbofan casing mounting points with the casing of the turbine hot ducting (14). First connectors (56,58,60) connect a first point (44) on the aircraft structure with the three mounting points on the turbofan casing. Second connectors (62,64) connect a second point (46) on the aircraft structure with the mounting points (52,54) on the blower casing. The three aircraft structure mounting points are arranged in a first plane.

Indicators of Incipient Surge for Three Turbofan Engines Using Standard Equipment and Instrumentation

Abstract: Engine controller data have been interrogated for indications of incipient surge for three turbofan engines; a large Pratt and Whitney, a large General Electric, and a small Williams International. Versions of these engines are currently operating in the field and all have compression ratios of 18 or greater. The Pratt and Whitney engine was surged only at full power while the other two were surged at partial power and at full power. The interest in this work was in detecting the presence of warning signatures for a current inventory of engines. A constraint was imposed on the experiments to use only existing engine instrumentation. The frequency response of the controller and the engine instrumentation limited the high frequency detection capability to about 100 Hz for the large engines and about 200 Hz for the small engine, For the large engines, it was not possible to detect a surge warning but for the small engine a sufficient warning of incipient surge was detected.Copyright © 1996 by ASME

Time-Averaged Active Controller for Turbofan Engine Fan Noise Reduction

Abstract: An active control system has been developed and used experimentally on an operational turbofan engine to reduce inlet tonal noise at the fan blade passage frequency. Both single channel and multichannel systems were tested. The control approach used is the feedforward derivative measurement time-averaged adaptive algorithm. This algorithm is capable of adapting to a nonstationary system, requiring no prior knowledge of the system. A reference sensor mounted in the casing of the fan provides the fan blade passing frequency reference signal to the controller. Microphones with a large sensing area are placed in the acoustic far field of the engine. The control algorithms minimize a cost function, the sum of the error signal mean-square-values, by activating loudspeakers that are mounted around the inlet of the engine. The control signals are obtained by filtering the reference signal with finite impulse response digital filters. The control algorithm finds the optimum filter weights by using statistical estimates of the cost function to be minimized. A single channel control system was tested on an axisymmetric dominant mode case and produced reductions of up to 17 dB at the error microphone. The controller maintained reductions in the fan tone even as the engine speed was increased. In a traverse of the radiated sound field of the engine the fan blade passage tone was reduced over a sector of 30 deg around the error sensor with a considerable spillover increase in the fan tone outside of this sector. Two six-channel control systems were developed and tested on both axisymmetric and spinning mode dominant cases. Reductions of up to 19.7 dB were achieved over large sectors around the error microphones with significantly reduced spillover.

Dynamic and Transient Performance of Turbofan/Turboshaft Convertible Engine With Variable Inlet Guide Vanes

Abstract: 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.

Multidisciplinary and Multiobjective Optimization in Conceptual Design for Mixed-Stream Turbofan Engines

Abstract: : Despite major advances in design tools such as engine cycle analysis software and computer aided design, conceptual gas turbine engine design is essentially a trial-and-error process based on the experience of engineers. Modern optimization concepts, such as multidisciplinary optimization (MDO), and multiobjective optimization (MOO), linked with sequential quadratic programming (SQP) methods and genetic algorithms (GA), were applied to the conceptual engine design process to automate the conceptual design phase. Robust integrated computer codes were created to find the optimal values of eight engine parameters in order to minimize fuel usage, aircraft cost and engine annulus area over a given mission. The engine cycle selected for study was the mixed stream, low bypass turbofan. SQP and GA optimization algorithms were integrated with on-design and off- design engine cycle analysis and mission analysis computer codes created by the authors to obtain the optimized conceptual engine design for an imaginary short range interceptor and the Global Strike Aircraft U.S. Air Force concept. The process used a nonspecific approach that can be applied to a wide variety of missions and aircraft. All the codes were written in Matlab, and so operate under the same programming architecture and can be easily upgraded or modified.

Turbofan with a floatury nacelle

Abstract: In a double flow turbojet, the mechanical link between the nacelle (12) and the frame of the fan (10) is limited to a centring system (22) situated immediately next to arms (20) by which the fan frame is mounted on the central part (16) of the engine. The aerodynamic and inertial loads applied to the nacelle (12), especially during take-off and landing, cause only a limited and acceptable deformation of the frame of the fan (10). The play between the frame and the fan (18) may be kept to a very small value, which ensure the maximum efficiency of the engine. The centring system (22) avoids excessive radial stress on the frame of the fan.

Performance Benefits for Wave Rotor-Topped Gas Turbine Engines

Abstract: 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.

Quiet High-Speed Fan

Abstract: A calibration of the acoustic and aerodynamic prediction methods was performed and a baseline fan definition was established and evaluated to support the quiet high speed fan program. A computational fluid dynamic analysis of the NASA QF-12 Fan rotor, using the DAWES flow simulation program was performed to demonstrate and verify the causes of the relatively poor aerodynamic performance observed during the fan test. In addition, the rotor flowfield characteristics were qualitatively compared to the acoustic measurements to identify the key acoustic characteristics of the flow. The V072 turbofan source noise prediction code was used to generate noise predictions for the TFE731-60 fan at three operating conditions and compared to experimental data. V072 results were also used in the Acoustic Radiation Code to generate far field noise for the TFE731-60 nacelle at three speed points for the blade passage tone. A full 3-D viscous flow simulation of the current production TFE731-60 fan rotor was performed with the DAWES flow analysis program. The DAWES analysis was used to estimate the onset of multiple pure tone noise, based on predictions of inlet shock position as a function of the rotor tip speed. Finally, the TFE731-60 fan rotor wake structure predicted by the DAWES program was used to define a redesigned stator with the leading edge configured to minimize the acoustic effects of rotor wake / stator interaction, without appreciably degrading performance.

A Full-Scale Testing of an Annular Turbo-Fan Combustor Including Emission Characteristics

Abstract: Korea Aerospace Research Institute has performed a joint research with the Central Institute of Aviation Motors of Russia on the design, manufacturing, and testing of an annular combustor for a turbo-fan engine with the thrust of 8,000 lbf. In order to reduce smoke generation, a rather large amount of air is introduced to the primary zone. Presented in this paper is a description of the full-scale annular combustor, the test facility, the test procedure, and the test results. The measured parameters include the pressure loss and its dependence on flow velocity, the combustion efficiency by gas analysis, the exit temperature pattern factor for a wide range of air excess ratio, the lean blow-off limit, and the emission characteristics. The main test conditions are the ‘ground idle’ condition and the ‘altitude cruise’ condition. It was confirmed that the exit temperature profile is closely related to the location of dilution holes on the flame tube. Lean blow-off limit is rather narrow since the combustor was designed to provide a large amount of air to the primary zone with an aim of smoke reduction.Copyright © 1996 by ASME