Showing papers on "Turbofan published in 2007"

User's Guide for the Commercial Modular Aero-Propulsion System Simulation (C-MAPSS)

Abstract: This report is a Users Guide for the NASA-developed Commercial Modular Aero-Propulsion System Simulation (C-MAPSS) software, which is a transient simulation of a large commercial turbofan engine (up to 90,000-lb thrust) with a realistic engine control system. The software supports easy access to health, control, and engine parameters through a graphical user interface (GUI). C-MAPSS provides the user with a graphical turbofan engine simulation environment in which advanced algorithms can be implemented and tested. C-MAPSS can run user-specified transient simulations, and it can generate state-space linear models of the nonlinear engine model at an operating point. The code has a number of GUI screens that allow point-and-click operation, and have editable fields for user-specified input. The software includes an atmospheric model which allows simulation of engine operation at altitudes from sea level to 40,000 ft, Mach numbers from 0 to 0.90, and ambient temperatures from -60 to 103 F. The package also includes a power-management system that allows the engine to be operated over a wide range of thrust levels throughout the full range of flight conditions.

Turbofan engine assembly and method of assembling same

Abstract: A turbofan engine assembly (10) is provided. A turbofan engine assembly includes a core gas turbine engine (13) including a high-pressure compressor (14), a combustor (16) disposed downstream from the high-pressure compressor, and a high-pressure turbine (18) coupled to the high-pressure compressor using a shaft. A turbofan engine assembly (10) further includes counter-rotating booster compressor (22) coupled to the core gas turbine engine, and the counter-rotating booster compressor includes a first rotor section configured to rotate in a first direction and a second rotor section configured to rotate in an opposite second direction. The turbofan engine assembly (10) further includes a single stage fan assembly (12) coupled to the first rotor section, and a drive shaft (31) coupled between the low-pressure turbine (20) and the fan assembly. A turbofan engine assembly (10) further includes a gearbox (100) coupled between the drive shaft and the second rotor section such that the low-pressure turbine drives the gearbox and such that the gearbox drives the second rotor section.

Epicyclic gear train for turbo fan engine

Abstract: A two-stage turbofan system for use in a gas turbine engine comprises a first-stage fan shaft, a second-stage fan shaft, a stationary torque tube and a gear system. The second-stage fan shaft connects with a drive shaft in the gas turbine engine such that the second-stage fan shaft is driven at the speed of the drive shaft. The stationary torque tube is connected with a fan case in the gas turbine engine. The gear system is connected to the second-stage fan shaft and the torque tube. The first-stage fan shaft extends from the gear system such that the first-stage fan shaft is driven at a speed reduced from that of the drive shaft.

Engine mounting configuration for a turbofan gas turbine engine

Abstract: An engine mounting configuration reacts engine thrust at an aft mount. The engine mounting configuration reduces backbone bending of the engine, intermediate case distortion and frees-up space within the core nacelle.

Turbofan engine cowl assembly and method of operating the same

Abstract: A thrust reverser assembly (100) for a turbofan engine assembly (10), the engine assembly including a core gas turbine engine (20), a core cowl (22) which circumscribes the core gas turbine engine (20), a nacelle (24) positioned radially outward from the core cowl, a fan nozzle duct (26) having an area defined between the core cowl and a portion of the nacelle, and a thrust reverser positioned within a portion of the nacelle. The thrust reverser assembly includes a first cowl (104), a second cowl (102) repositionable with respect to the first cowl for varying the area of the fan nozzle duct, and at least one flow directing member (133, 137) coupled to the second cowl to selectively impede airflow from flowing through a cavity (156) defined between the first cowl and the second cowl.

Challenges in the silent aircraft engine design

Abstract: The Silent Aircraft Initiative goal is to design an aircraft that is imperceptible above background noise outside the airport boundary. The aircraft that fulfils this objective must also be economically competitive with conventional aircraft of the future and therefore fuel consumption and mechanical reliability are key considerations for the design. To meet these ambitious targets, a multi-fan embedded turbofan engine with boundary layer ingestion has been proposed. This configuration includes several new technologies including a variable area nozzle, a complex high-power transmission system, a Low Pressure turbine designed for low-noise, an axial-radial HP compressor, advanced acoustic liners and a low-speed fan optimized for both cruise and off-design operation. These technologies, in combination, enable a low-noise and fuel efficient propulsion system but they also introduce significant challenges into the design. These challenges include difficulties in predicting the noise and performance of the new components but there are also challenges in reducing the design risks and proving that the new concepts are realizable. This paper presents the details of the engine configuration that has been developed for the Silent Aircraft application. It describes the design approach used for the critical components and discusses the benefits of the new technologies. The new technologies are expected to offer significant benefits in noise reduction without compromising fuel burn. However, more detailed design and further research are required to fully control the additional risks generated by the system complexity.

Epicyclic gear train for variable cycle engine

Abstract: A two-stage turbofan system for use in a gas turbine engine comprises a first-stage fan shaft, a second-stage fan shaft, a stationary torque tube and a gear system. The second-stage fan shaft connects with a drive shaft in the gas turbine engine such that the second-stage fan shaft is driven at the speed of the drive shaft. The stationary torque tube is connected with a fan case in the gas turbine engine. The gear system is connected to the second-stage fan shaft and the torque tube. The first-stage fan shaft extends from the gear system such that the first-stage fan shaft is driven at a speed reduced from that of the drive shaft.

Exergetic analysis of an aircraft turbofan engine

Abstract: The main objective of the present study is to perform an exergy analysis of a turbofan kerosene-fired engine with afterburner (AB) at sea level and an altitude of 11 000 m. The main components of this engine include a fan, a compressor, a combustion chamber, a turbine, an AB and an exhaust. Exergy destructions in each of the engine components are determined, while exergy efficiency values for both altitudes are calculated. The AB unit is found to have the highest exergy destruction with 48.1% of the whole engine at the sea level, followed by the exhaust, the combustion chamber and the turbine amounting to 29.7, 17.2 and 2.5%, respectively. The corresponding exergy efficiency values for the four components on the product/fuel basis are obtained to be 59.9, 65.6, 66.7 and 88.5%, while those for the whole engine at the sea level and an altitude of 11 000 m are calculated to be 66.1 and 54.2%. Copyright © 2007 John Wiley & Sons, Ltd.

Turbofan duct mode measurements using a continuously rotating microphone rake

Abstract: An experimental measurement system was developed and implemented at the NASA Glenn Research Center in the 1990s to measure fan duct acoustic modes. The system is a radial array of microphones inserted into the duct that continuously rotates about the duct centerline. This Rotating Rake provides a complete map of the acoustic duct modes at fan harmonics in a ducted fan. It has been used on a variety of test articles: from a low-speed, concept test fan rig, to a full-scale production turbofan engine. The Rotating Rake has been critical in developing and evaluating a number of noise reduction concepts as well as providing experimental databases for verification of several aero-acoustic codes.

Geared counter-rotating gas turbofan engine

Abstract: A gas turbine engine includes a gear system driven by a first and second counter rotating low pressure shaft and a fan driven by the gear system.

Gas turbine engine with variable geometry fan exit guide vane system

Abstract: A turbofan engine includes a variable geometry fan exit guide vane (FEGV) system having a multiple of circumferentially spaced radially extending fan exit guide vanes. Rotation of the fan exit guide vanes between a nominal position and a rotated position selectively changes a fan bypass flow path to permit efficient operation at various flight conditions.

Noise Reduction Technologies for Turbofan Engines

Abstract: Significant progress continues to be made with noise reduction for turbofan engines. NASA has conducted and sponsored research aimed at reducing noise from commercial aircraft. Since it takes many years for technologies to be developed and implemented, it is important to have aggressive technology goals that lead the target entry into service dates. Engine noise is one of the major contributors to the overall sound levels as aircraft operate near airports. Turbofan engines are commonly used on commercial transports due to their advantage for higher performance and lower noise. The noise reduction comes from combinations of changes to the engine cycle parameters and low noise design features. In this paper, an overview of major accomplishments from recent NASA research programs for engine noise will be given.

Feasibility of In-Duct Beamforming

Abstract: *This paper discusses the possibility of source loca tion by phased array beamforming on fan and stator of turbofan engines using an intake wall-mounted microphone array. To demonstrate the feasibility, beamforming techniques were applied to existing measured data of a circular microphone array in the intake of a f an rig. This array is normally used for azimuthal mode detection. From the measured data set a test case was selected with low engine speed, and with an intake liner between the fan and the array. Beamforming methods with both stationary and rotating focus are applied , and the contributions of tonal noise and broadband noise were separated. Thus, tonal and broadband noise sources can be identified on both fan and stator. The free-field Green’s func tion is used for the definition of the steering vectors, thereby ignoring duct wall reflec tions. By a simulation study it is shown that the presence of an intake liner is a necessary condition for obtaining meaningful beamforming results. Further simulations showed that significantly improved beamforming results can be obtained with an in-duct “cage” arra y consisting of a number of parallel microphone rings.

Turbofan jet engine

Abstract: To provide a turbofan jet engine for a supersonic aircraft, which enables a supersonic cruise with small additional drag due to engine installation while minimizing the jet noise at takeoff. In the turbofan jet engine, two fans, i.e., a front fan and an aft fan, are disposed, and an air inlet and an air outlet are provided on each of the front and aft fans, whereby at takeoff, air is introduced to each of the air inlets of the front and aft fans to drive the front and aft fans in parallel, and the air is then ejected from each of the outlets. On the other hand, during a supersonic cruise, an inlet of an aft fan duct is coupled to an outlet of a front fan duct, where external air is introduced only to an air inlet of the front fan duct and the air compressed through the front fan is fed to the aft fan duct and further compressed through the aft fan, and ejected from a nozzle, so that the front and aft fans are driven in series.

Pivoting door thrust reverser for a turbofan gas turbine engine

Abstract: A pivot arrangement for a thrust reverser door of a gas turbine engine, the pivot fitting having a base and a shaft projecting from a main side of the base. The shaft receives a preferably curved pivot arm of the door to provide a low profile arrangement which improves performance when the doors are stowed.

Operation of an aircraft engine after emergency shutdown

Abstract: A method of emergency operation of an aircraft turbofan engine during an aircraft flight is provided. The engine includes a fan shaft with a fan, and an electric motor/generator mounted for rotation therewith. The method includes shutting down the engine while allowing the engine to windmill, operating the electric motor/generator to rotate the shaft at a determined windmilling speed which is desired for the fan shaft, and operating the engine at the desired windmilling speed for substantially a remainder of the aircraft flight.

Integrated propulsive system comprising a bypass turbojet engine

Abstract: An integrated propulsive turbofan engine system includes a downstream cylindrical nacelle structure in two parts. The first of the two parts is fixed and fastens engine attachment elements and an intermediate casing to the aircraft. The second of the two parts may move between a closed position used during operation and an open position during maintenance when access to engine components is needed.

Bleed Air versus Electric Power Off-takes from a Turbofan Gas Turbine over the Flight Cycle

Abstract: [] In this paper the performance of the gas turbine engines of a commercial passenger aircraft is evaluated for both bleed air off-take and electric power off-take. As these types of engine power off-takes are not directly comparable, an exergy analysis is used to establish the most efficient type of off-take. From this analysis appears that it is indeed more efficient to bleed air from the engine instead of generating the equivalent amount of exergy in terms of electric power. However, when also taking into account the performance of the largest pneumatic power consumer, the Environmental Control System (ECS) it appears that about 2% thrust specific fuel consumption can be saved, by using a MoreElectric ECS instead of a conventional bleed air powered ECS.

Turbofan engine assembly

Abstract: PROBLEM TO BE SOLVED: To provide a turbofan engine with a counter-rotating fan without posing such a problem as weight increase caused by a counter-rotating low-pressure turbine. SOLUTION: The turbine engine assembly includes a core gas turbine engine 12 including a high-pressure compressor 18, a combustor 20, and a high-pressure turbine 22. The turbine engine assembly further includes a booster compressor 30 coupled upstream from the core gas turbine engine. The turbine engine assembly further includes an intermediate-pressure turbine 32 coupled to the booster compressor, the intermediate-pressure turbine disposed downstream from the core gas turbine engine. The turbine engine assembly further includes a counter-rotating fan assembly disposed upstream from the booster compressor, the counter-rotating fan assembly includes a first fan assembly configured to rotate in a first direction and a second fan assembly configured to rotate in an opposite second direction. The turbine engine assembly further includes a low-pressure turbine 14 disposed downstream from the intermediate-pressure turbine, the low-pressure turbine configured to drive the counter-rotating fan assembly. COPYRIGHT: (C)2008,JPO&INPIT

Aircraft engine pre-dressing unit for testing facility

Abstract: A turbofan text cell pre-dressing unit that enables the installation of the engine in a text cell in minimal time is provided.

A Parametric Study on the Effects of Inlet Swirl on Compression System Performance and Operability Using Numerical Simulations

Abstract: The integration of the airframe and propulsion system is a key design issue in the development and deployment of military aircraft. Many disciplines comprise this issue, one being the aerodynamic interaction between the inlet system and the engine. The external airframe and inlet system must capture flow from the free stream and deliver it to the installed engine in a manner that maintains engine stability. Using the existing SAE S-16 developed methodologies, inlet distortion has traditionally been characterized by consideration of total-pressure distortion, total-temperature distortion, or planar waves, either singularly or in combination. However, many gas turbine installations can generate significant flow angularity as well as total pressure distortion at the Aerodynamic Interface Plane (AIP). The flow angularities may have both radial- and circumferential-velocity components. Swirl is generally considered as being that portion of the flow vector which is directed circumferentially, since it is this velocity component which directly affects the work of a downstream fan or compressor. The objective of this paper is to demonstrate that simple but effective numerical simulations can provide qualitative insights into the effect of swirl on compression system performance and operability. The analysis presented in this paper uses both a single rotor and an advanced 2-stage fan exhibiting many modern military turbofan design features.

Nacelle for turbofan

Abstract: The nacelle of the invention includes an air-inlet front section, a middle section for encasing the fan of the turbofan (7) and a rear section including means for connection to a pylon to be connected to the rigid structure or an airplane. The rear section includes a structural framework (18) having at least one aerodynamic smoothing and acoustic panel (21). The acoustic panel (21) is secured on the structural framework (18) by floating or elastic fixation means so that said acoustic panel (21) can be deformed in a direction (31) essentially centrifugal and radial relative to he turbofan (7) in the presence of overpressure air (3) in the engine compartment.

Testing and Performance Verification of a High Bypass Ratio Turbofan Rotor in an Internal Flow Component Test Facility

Abstract: A 1/5 scale model rotor representative of a current technology, high bypass ratio, turbofan engine was installed and tested in the W8 single-stage, high-speed, compressor test facility at NASA Glenn Research Center (GRC). The same fan rotor was tested previously in the GRC 9x15 Low Speed Wind Tunnel as a fan module consisting of the rotor and outlet guide vanes mounted in a flight-like nacelle. The W8 test verified that the aerodynamic performance and detailed flow field of the rotor as installed in W8 were representative of the wind tunnel fan module installation. Modifications to W8 were necessary to ensure that this internal flow facility would have a flow field at the test package that is representative of flow conditions in the wind tunnel installation. Inlet flow conditioning was designed and installed in W8 to lower the fan face turbulence intensity to less than 1.0 percent in order to better match the wind tunnel operating environment. Also, inlet bleed was added to thin the casing boundary layer to be more representative of a flight nacelle boundary layer. On the 100 percent speed operating line the fan pressure rise and mass flow rate agreed with the wind tunnel data to within 1 percent. Detailed hot film surveys of the inlet flow, inlet boundary layer and fan exit flow were compared to results from the wind tunnel. The effect of inlet casing boundary layer thickness on fan performance was quantified. Challenges and lessons learned from testing this high flow, low static pressure rise fan in an internal flow facility are discussed.

Optimization of Annular and Cylindrical Liners for Mixed Exhaust Aeroengines

Abstract: The increase in bypass ratios for modern day turbofan engines has led to fan rearward broadband noise emerging as one of the principal aircraft noise sources. The most widespread method for mitigating this noise source is the installation of acoustic liners in the ducting downstream of the fan. Mixed exhaust designs have the potential to be quieter than a three quarter cowl aeroengine due to the additional surface available for liner application. As a result, the Silent Aircraft concept design contains an embedded distributed propulsion system to exploit this opportunity. This paper outlines a suitable method for optimizing the annular and cylindrical liners that make up the exhaust system. To gain maximum benefit, liners should be designed to absorb the modal disturbances that are the most significant for an observer on the ground. The optimization uses an advanced noise footprint cost function to achieve this. Consequently, the liners and the length of the installation have been chosen to meet the aggressive noise target set for the aircraft. In addition, we explore the effect that modal scattering between axial liner junctions can have on the acoustic performance of the system. We find that energy interchange between radial modes will often have a detrimental effect on the overall attenuation, although this can be addressed by selecting the order of the liners in a more considered way. The assumption that scattering between liners can be neglected during optimization to allow design of the individual components therefore appears to be valid in this case. However, subsequent attention has to be directed to the order in which they are placed. Current and next generation mixed exhaust aeroengines will be restricted to smaller liner segment lengths than the Silent Aircraft. This analysis suggests that scattering and hence liner order is likely to more important for these designs.

Adaptive On-Wing Gas Turbine Engine Performance Estimation

Abstract: A key technological concept for producing reliable engine diagnostics and prognostics exploits the benefits of fusing sensor data, information, and/or processing algorithms. In this paper, we consider a real-time physics based model of a commercial turbofan engine called STORM: self tuning on-board, real-time engine model. The STORM system provides a means for tracking engine module performance changes in real-time. However, modeling error can have a corruptive effect on STORM's estimation of performance changes. Fusing an empirical neural network based model with STORM forms a unique hybrid model of the engine called enhanced STORM (eSTORM). This approach can eliminate the STORM engine diagnostic errors. A practical consideration for implementing the hybrid engine model, involves the application of some form of sequential model building to construct and specify the empirical elements. A methodology for constructing the empirical model (EM) in a sequential manner without the requirement for storing all of the original data has been developed. This paper describes the development of the adaptive hybrid model scheme for a commercial turbofan engine. This adaptive hybrid-modeling scheme has been implemented in real-time on an intelligent automation corporation (IAC) computational platform. Model performance achieved with the automated update algorithm using real on-wing commercial aircraft engine data will be presented.

Effects of Compressor Tip Injection on Aircraft Engine Performance and Stability

Abstract: This analytical study discusses the system aspects of active stability enhancement using mass flow injection in front of the rotor blade tip of a high pressure compressor. Tip injection is modeled as a recirculating bleed in a performance simulation of a commercial turbofan engine. A map correction procedure accounts for the changes in compressor characteristics caused by injection. The correction factors are derived from stage stacking calculations, which include a simple correlation for stability enhancement. The operational characteristic of the actively controlled engine is simulated in steady and transient states. The basic steady-state effect consists of a local change in mass flow and a local increase in gas temperature. This alters the component matching in the engine. The mechanism can be described by the compressor-to-turbine flow ratio and the injection temperature ratio. Both effects reduce the cycle efficiency resulting in an increased turbine temperature and fuel consumption at constant thrust. The negative performance impact becomes negligible if compressor recirculation is only employed at the transient part power and if valves remain closed at the steady-state operation. Detailed calculations show that engine handling requirements and temperature limits will still be met. Tip injection increases the high pressure compressor stability margin substantially during critical maneuvers. The proposed concept in combination with an adequate control logic offers promising benefits at transient operation, leading to an improvement potential for the overall engine performance.

Moving blade of turbofan engine

Abstract: [OBJECT] An object of the invention is to provide a fan rotating blade for a turbofan engine capable of increasing a bypass ratio by increasing an intake air flow rate without enlarging a diameter of a fan and an inner diameter of a casing and of realizing a decrease in weight of an engine as well as a decrease in fuel consumption and noise. [MEANS FOR SOLVING THE PROBLEMS] A leading edge part 11 of a fan rotating blade 10 for taking air thereinto is provided with a vertical hub portion 12 positioned on the hub side so as to be substantially perpendicular to a fan rotary shaft, a backward mid-span portion 13 inclined toward the downstream side from the hub side to the mid-span portion, and a forward inclined tip portion 14 inclined toward the upstream side from the mid-span side to the tip portion.