Showing papers on "Turbofan published in 2009"

Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

Abstract: Meeting NASA's N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This paper discusses a number of these new possibilities. The Boeing N2 hybrid-wing-body (HWB) is used as a baseline aircraft for this study. The two pylon mounted conventional turbofans are replaced by two wing-tip mounted turboshaft engines, each driving a superconducting generator. Both generators feed a common electrical bus which distributes power to an array of superconducting motor-driven fans in a continuous nacelle centered along the trailing edge of the upper surface of the wing-body. A key finding was that traditional inlet performance methodology has to be modified when most of the air entering the inlet is boundary layer air. A very thorough and detailed propulsion/airframe integration (PAI) analysis is required at the very beginning of the design process since embedded engine inlet performance must be based on conditions at the inlet lip rather than freestream conditions. Examination of a range of fan pressure ratios yielded a minimum Thrust-specific-fuel-consumption (TSFC) at the aerodynamic design point of the vehicle (31,000 ft /Mach 0.8) between 1.3 and 1.35 FPR. We deduced that this was due to the higher pressure losses prior to the fan inlet as well as higher losses in the 2-D inlets and nozzles. This FPR is likely to be higher than the FPR that yields a minimum TSFC in a pylon mounted engine. 1

Aircraft engine controls : design, system analysis, and health monitoring

Abstract: Covers the design of engine control & monitoring systems for both turbofan & turboshaft engines, focusing on four key topics: modeling of engine dynamics; application of specific control design methods to gas turbine engines; advanced control concepts; &, engine condition monitoring.

Turboelectric Distributed Propulsion Engine Cycle Analysis for Hybrid-Wing-Body Aircraft

Abstract: Meeting NASA's N+3 goals requires a fundamental shift in approach to aircraft and engine design. Material and design improvements allow higher pressure and higher temperature core engines which improve the thermal efficiency. Propulsive efficiency, the other half of the overall efficiency equation, however, is largely determined by the fan pressure ratio (FPR). Lower FPR increases propulsive efficiency, but also dramatically reduces fan shaft speed through the combination of larger diameter fans and reduced fan tip speed limits. The result is that below an FPR of 1.5 the maximum fan shaft speed makes direct drive turbines problematic. However, it is the low pressure ratio fans that allow the improvement in propulsive efficiency which, along with improvements in thermal efficiency in the core, contributes strongly to meeting the N+3 goals for fuel burn reduction. The lower fan exhaust velocities resulting from lower FPRs are also key to meeting the aircraft noise goals. Adding a gear box to the standard turbofan engine allows acceptable turbine speeds to be maintained. However, development of a 50,000+ hp gearbox required by fans in a large twin engine transport aircraft presents an extreme technical challenge, therefore another approach is needed. This paper presents a propulsion system which transmits power from the turbine to the fan electrically rather than mechanically. Recent and anticipated advances in high temperature superconducting generators, motors, and power lines offer the possibility that such devices can be used to transmit turbine power in aircraft without an excessive weight penalty. Moving to such a power transmission system does more than provide better matching between fan and turbine shaft speeds. The relative ease with which electrical power can be distributed throughout the aircraft opens up numerous other possibilities for new aircraft and propulsion configurations and modes of operation. This paper discusses a number of these new possibilities. The Boeing N2 hybrid-wing-body (HWB) is used as a baseline aircraft for this study. The two pylon mounted conventional turbofans are replaced by two wing-tip mounted turboshaft engines, each driving a superconducting generator. Both generators feed a common electrical bus which distributes power to an array of superconducting motor-driven fans in a continuous nacelle centered along the trailing edge of the upper surface of the wing-body. A key finding was that traditional inlet performance methodology has to be modified when most of the air entering the inlet is boundary layer air. A very thorough and detailed propulsion/airframe integration (PAI) analysis is required at the very beginning of the design process since embedded engine inlet performance must be based on conditions at the inlet lip rather than freestream conditions. Examination of a range of fan pressure ratios yielded a minimum Thrust-specific-fuel-consumption (TSFC) at the aerodynamic design point of the vehicle (31,000 ft /Mach 0.8) between 1.3 and 1.35 FPR. We deduced that this was due to the higher pressure losses prior to the fan inlet as well as higher losses in the 2-D inlets and nozzles. This FPR is likely to be higher than the FPR that yields a minimum TSFC in a pylon mounted engine. 1

Fundamental Differences Between Conventional and Geared Turbofans

Abstract: The potential for improving the thermodynamic efficiency of aircraft engines is limited because the aerodynamic quality of the turbomachines has already achieved a very high level. While in the past increasing burner exit temperature did contribute to better cycle efficiency, this is no longer the case with today’s temperatures in the range of 1900...2000K. Increasing the cycle pressure ratio above 40 will yield only a small fuel consumption benefit. Therefore the only way to improve the fuel efficiency of aircraft engines significantly is to increase bypass ratio — which yields higher propulsive efficiency. A purely thermodynamic cycle study shows that specific fuel consumption decreases continuously with increasing bypass ratio. However, thermodynamics alone is a too simplistic view of the problem. A conventional direct drive turbofan of bypass ratio 6 looks very different to an engine with bypass ratio 10. Increasing bypass ratio above 10 makes it attractive to design an engine with a gearbox to separate the fan speed from the other low pressure components. Different rules apply for optimizing turbofans of conventional designs and those with a gearbox. This paper describes various criteria to be considered for optimizing the respective engines and their components. For illustrating the main differences between conventional and geared turbofans it is assumed that an existing core of medium pressure ratio with a two stage high pressure turbine is to be used. The design of the engines is done for takeoff rating because this is the mechanically most challenging condition. For each engine the flow annulus is examined and stress calculations for the disks are performed. The result of the integrated aero-thermodynamic and mechanical study allows a comparison of the fundamental differences between conventional and geared turbofans. At the same bypass ratio there will be no significant difference in specific fuel consumption between the alternative designs. The main difference is in the parts count which is much lower for the geared turbofan than for the conventional engine. However, these parts will be mechanically much more challenging than those of a conventional turbofan. If the bypass ratio is increased significantly above 10, then the geared turbofan becomes more and more attractive and the conventional turbofan design is no longer a real option. The maximum practical bypass ratio for ducted fans depends on the nacelle drag and how the installation problems can be solved.Copyright © 2009 by ASME

Numerical methods for noise propagation in moving flows, with application to turbofan engines

Abstract: This article reviews the development and application of Computational AeroAcoustics (CAA) to acoustic propagation on subsonic mean flows, with a particular focus on methods used to predict acoustic radiation from turbofan aeroengines. The governing equations are presented and particular issues such as the formulation of impedance and far field boundary conditions, the treament of Kelvin-Helmholtz instabilities, resolution requirements and methods for controlling dispersion error, are discussed. The status of current CAA methods is reviewed. Finally, the matter of validation against benchmark problems and measured data is explored.

Design and control of a variable geometry turbofan with and independently modulated third stream

Abstract: Emerging 21st century military missions task engines to deliver the fuel efficiency of a high bypass turbofan while retaining the ability to produce the high specific thrust of a low bypass turbofan. This study explores the possibility of satisfying such competing demands by adding a second independently modulated bypass stream to the basic turbofan architecture. This third stream can be used for a variety of purposes including: providing a cool heat sink for dissipating aircraft heat loads, cooling turbine cooling air, and providing a readily available stream of constant pressure ratio air for lift augmentation. Furthermore, by modulating airflow to the second and third streams, it is possible to continuously match the engine’s airflow demand to the inlet’s airflow supply thereby reducing spillage and increasing propulsive efficiency. This research begins with a historical perspective of variable cycle engines and shows a logical progression to proposed architectures. Then a novel method for investigating optimal performance is presented which determines most favorable on design variable geometry settings, most beneficial moment to terminate flow holding, and an optimal scheduling of variable features for fuel efficient off design operation. Mission analysis conducted across the three candidate missions verifies that these three stream variable cycles can deliver fuel savings in excess of 30% relative to a year 2000 reference turbofan. This research concludes by evaluating the relative impact of each variable technology on the performance of adaptive engine architectures. The most promising technologies include modulated turbine cooling air, variable high pressure turbine inlet area and variable third stream nozzle throat area. With just these few features it is possible to obtain nearly optimal performance, including 90% or more of the potential fuel savings, with far fewer variable features than are available in the study engine. It is abundantly clear that three stream variable architectures can significantly outperform existing two stream turbofans in both fuel efficiency and at the vehicle system level with only a modest increase in complexity and weight. Such engine architectures should be strongly considered for future military applications.

Fan Root Aerodynamics for Large Bypass Gas Turbine Engines: Influence on the Engine Performance and 3D Design

Abstract: The exit flow field of the fan root of large turbofan engines defines the inlet conditions to the core compressor. This in turn could have significant impact to the performance of the core compressor. This study is aimed to resolve two related issues concerning the impact of the fan root flow on the core compressor performance: to establish the effect of an increased loss at the inlet on the engine specific fuel consumption (SFC) and to assess the effect of the radial distribution of the fan root flow on the engine performance. With understanding of these issues, the geometric parameters and design details which can produce a more uniform core flow at the exit of the fan stage module can be identified. The fan root flow is analysed with methods of different complexity and fidelity. A simple cycle analysis is used to assess the impact on engine SFC of a stagnation pressure deficit at the fan root; a throughflow code is used for the preliminary study of the curvature effect of the root flow path and 3D RANS CFD calculations are then used to simulate the flow path from the inlet of the fan to the first stage of the core compressor. The adequacy of the application of the numerical code in this case has been assessed and confirmed by the comparison with the experimental data for two rig configurations. The results of this study show that the flow at the fan hub region is very complex and dominated by 3D effects. The interaction of the secondary flow with real geometries, such as leakage flows, is found to have a strong detrimental effect on the core performance. The curvature of the hub end-wall is a key parameter controlling the fan root flow topology; it influences the strength of the secondary flow, the spanwise distribution of the flow and its sensitivity to leakage flow. With this understanding it is possible to redesign of the fan hub flow path to reduce the loss generation by a significant amount.© 2009 ASME

Dynamic Loads in the Fan Containment Structure of a Turbofan Engine

Abstract: In accordance with the FAA certification requirements, all modern commercial turbofan engines must successfully demonstrate its ability to withstand a fan blade-out (FBO) event through actual test. Possibility of losing a rotating fan blade from a running engine is a flight safety consideration, which must be addressed during the design phase of the engine. A typical fan blade-out event involves very complex nonlinear transient dynamics with large deflection of the release blade and rigid body rotation of the trailing blade as well as progressive failure and fragmentation of various components. Due to the nature of the impact type loading, the solution to the problem should also address dependence of the material behavior such as yield strength as a function of strain rates. In short, the transient dynamic analysis of a fan blade-out event highlights the complexity of the numerical technique, which includes all the nonlinearities of structural dynamics: plastic behavior of the materials, large displacemen...

Exergy analysis of a turbofan aircraft engine

Abstract: An exergy analysis is reported for a General Electric turbofan engine (the CF6-80) using sea-level data. The effects on exergy efficiencies and exergy destructions are investigated of modifying the isentropic efficiencies of turbomachinery components. The most irreversible units in the system are found to be the fan and the core engine exhaust, with exergy loss rates of 47.3 MW and 35.9 MW, respectively, and the combustion chamber, with an exergy destruction rate of 31.5 MW. The exergy efficiencies of the fan and the core engine exhausts are found to be 12.9 and 12.7%, respectively.

Exergoeconomic analysis of an aircraft turbofan engine

Abstract: This study deals with an exergoeconomic analysis of an aircraft turbofan engine utilising the kerosene as fuel. A new parameter is developed to define the thrust cost rate. The cost of exergy destruction, the relative cost difference and the exergoeconomic factor are investigated. The variation of the relative cost difference and exergoeconomic factor according to the operating and maintenance costs and the annual operating hour are also studied. For a high by-pass and high thrust rated engine, the cost rate of thrust is obtained to be 304.35 $(hkN) −1 for the hot thrust and 138.96 $ (hkN)−1 for the cold thrust, respectively.

Ultra-efficient propulsor with an augmentor fan circumscribing a turbofan

Abstract: An ultra-efficient “green” aircraft propulsor utilizing an augmentor fan is disclosed. A balanced design is provided combining a fuel efficient and low-noise high bypass ratio augmentor fan and a low-noise shrouded high bypass ratio turbofan. Three mass flow streams are utilized to reduce propulsor specific fuel consumption and increase performance relative to conventional turbofans. Methods are provided for optimization of fuel efficiency, power, and noise by varying mass flow ratios of the three mass flow streams. Methods are also provided for integration of external propellers into turbofan machinery.

Bypass air scoop for gas turbine engine

Abstract: A gas turbine engine has a plurality of radial struts in a bypass duct. At least one strut has a scoop incorporated with the fairing of the strut and in communication with an air passage of an engine secondary air system. The scoop faces a bypass air flow to scoop a portion of the bypass air flow using available dynamic pressure in the bypass duct. Scooped air may be provided, for example, to an active tip clearance control apparatus in a long duct turbofan engine.

Noise-Source Separation Using Internal and Far-Field Sensors for a Full-Scale Turbofan Engine

Abstract: Noise-source separation techniques for the extraction of the sub-dominant combustion noise from the total noise signatures obtained in static-engine tests are described. Three methods are applied to data from a static, full-scale engine test. Both 1/3-octave and narrow-band results are discussed. The results are used to assess the combustion-noise prediction capability of the Aircraft Noise Prediction Program (ANOPP). A new additional phase-angle-based discriminator for the three-signal method is also introduced.

Attenuation of FJ44 Turbofan Engine Noise With a Foam-Metal Liner Installed Over-the-Rotor

Abstract: A Williams International FJ44-3A 3000-lb thrust class turbofan engine was used as a demonstrator for a Foam-Metal Liner (FML) installed in close proximity to the fan. Two FML designs were tested and compared to the hardwall baseline. Traditional single degree-of-freedom liner designs were also evaluated to provide a comparison. Farfield acoustic levels and limited engine performance results are presented in this paper. The results show that the FML achieved up to 5 dB Acoustic Power Level (PWL) overall attenuation in the forward quadrant, equivalent to the traditional liner design. An earlier report presented the test set-up and conditions.

Analysis of Turbofan Design Options for an Advanced Single-Aisle Transport Aircraft

Abstract: The desire for higher engine efficiency has resulted in the evolution of aircraft gas turbine engines from turbojets, to low bypass ratio, first generation turbofans, to today's high bypass ratio turbofans. It is possible that future designs will continue this trend, leading to very-high or ultra-high bypass ratio (UHB) engines. Although increased bypass ratio has clear benefits in terms of propulsion system metrics such as specific fuel consumption, these benefits may not translate into aircraft system level benefits due to integration penalties. In this study, the design trade space for advanced turbofan engines applied to a single-aisle transport (737/A320 class aircraft) is explored. The benefits of increased bypass ratio and associated enabling technologies such as geared fan drive are found to depend on the primary metrics of interest. For example, bypass ratios at which fuel consumption is minimized may not require geared fan technology. However, geared fan drive does enable higher bypass ratio designs which result in lower noise. Regardless of the engine architecture chosen, the results of this study indicate the potential for the advanced aircraft to realize substantial improvements in fuel efficiency, emissions, and noise compared to the current vehicles in this size class.

High bypass-ratio turbofan jet engine

Abstract: To provide a turbofan jet engine which is capable of increasing the bypass ratio without increasing the fan diameter, and of reducing air resistance acting on the engine, a front fan duct that discharges air compressed by a front fan to the atmosphere and an aft fan duct that introduces air into an aft fan are disposed such as to change cross-sectional shapes thereof by rotating around a core engine in a counterclockwise direction, so that the cross-sectional geometric relationship between the front fan duct and the aft fan duct at a position immediately posterior to the front fan and a cross-sectional geometric relationship between the front fan duct and the aft fan duct at a position immediately anterior to the aft fan are inverted.

Effects of incidence angle in bird strike on integrity of aero-engine fan blade

Abstract: The possibilities of blade deformation and even fracture, the need for subsequent containment and reduction in thrust and power supply make bird strikes to aero-engine fan blades very serious events. Due to the different bird–plane flight paths and the different types of turbofan engines, the incidence angle between the bird and the engine fan blade can vary within a wide range. The present explicit non-linear three-dimensional finite-element analyses examine in detail the effect of the incidence angle when a canonical 4-lb bird strikes a typical engine fan blade, using the commercial code LS-DYNA. Both the bird and the blade are simulated in a Lagrangian framework. The homogenised fluidic constitutive equation of the bird follows the Brockman hydrodynamic model, while the blade is modelled as a viscoplastic material of the Perzyna type. It was found that normal incidence results in maximum impact forces and plastic strains leading to severe deformation. For the case in which the incidence angle is equal ...

Evaluation of the relationship between exhaust gas temperature and operational parameters in CFM56-7B engines

Abstract: This study presents the relationship between exhaust gas temperature (EGT) and engine operational parameters at two different power settings, including maximum continuous and take-off, in the CFM56-7B turbofan engine. The ground measurements of engine operational parameters including net thrust, fuel flow, low rotational speed, core rotational speed, pressure ratio, air temperature at engine fan inlet, take-off margin temperature, and thrust-specific fuel consumption of 51 different CFM56-7B engines are used to find the relationship mentioned in the study. This engine type is selected due to its common use by the civil aviation sector. In accordance with the results of multiple linear regression analysis, it was shown that EGT is affected by the engine operational parameters in different rate. The relationship between EGT and the operational parameters used in the maximum continuous power setting is slightly stronger than that of take-off power setting, R2=0.73 and 0.69, respectively. The fuel flo...

Intercooled turbofan engine design and technology research in the EU Framework 6 NEWAC programme

Abstract: This paper describes design studies on the high OPR intercooled aero engines in sub-programme 1 of the EU Framework 6 NEWAC programme, and rig tests and experiments in sub-programme 3 that relate to the design of efficient, compact and lightweight intercooled compression systems with enhanced operability. It summarizes the results obtained up to the third year of the project including a preliminary assessment of the new intercooled turbofan technologies in engines of 140 kN and 320 kN takeoff thrust. Glossary

Real-Time Transient Three Spool Turbofan Engine Simulation: A Hybrid Approach

Abstract: This paper presents a transient three-spool turbofan engine simulation model that uses a combination of intercomponent volume and iterative techniques. The engine model runs in real time and has been implemented in MATLAB /SIMULINK environment. The main advantage of this hybrid approach is that it preserves the accuracy of the iterative method while maintaining the simplicity of the intercomponent volume method. The iterative approach is applied at each engine subsystem to solve algebraic thermodynamic equations for exit enthalpy, entropy, and temperature, whereas the intercomponent volume method is used to calculate pressures derivatives and hence pressures at corresponding engine stations. This allows the engine state vector to be updated at each pass through the engine calculations. This technique was applied as a test case on the Rolls Royce Trent 500 three-spool turbofan engine, and the results were compared with an iterative method. As the engine state vector is updated during each cycle, the model lends itself for easy integration into nonlinear aircraft simulations, real-time engine diagnostics/prognostics, and jet engine control applications.

Gas turbine engine nacelle

Abstract: A nacelle for a turbofan gas turbine engine is provided. An inner wall of the nacelle defines an air intake which directs air into the fan section of the engine. The intake has, in flow series, an intake lip, a throat and a diffuser. The diffuser has, in flow series, first and second flow conditioning sections over both of which the flow cross-sectional area of the diffuser increases with increasing downstream distance from the throat. In addition, over the second section the nacelle inner wall lies substantially on a surface of an oblique circular cone having an apex which is offset from the centreline of the engine.

Gas turbine engine variable area exhaust nozzle

Abstract: A turbofan gas turbine engine ( 10 ) comprises a variable area exhaust nozzle ( 12 ) arranged at the downstream end of a casing ( 17 ). A control unit ( 66 ) analyses the power produced by the gas turbine engine ( 10 ), the flight speed of the gas turbine engine ( 1 ) and/or the altitude of the gas turbine engine ( 10 ). The control unit ( 66 ) configures the variable area nozzle ( 12 ) at a first cross-sectional area ( 70 A) when the flight speed of the gas turbine engine ( 10 ) is less than a first predetermined value. The control unit ( 66 ) configures the variable area nozzle ( 12 ) at a second, smaller, cross-sectional area ( 70 B) when the flight speed of the gas turbine engine ( 10 ) is greater than the first predetermined value and the power produced by the gas turbine engine ( 10 ) is greater than a second predetermined value. The control unit ( 66 ) configures the variable area nozzle ( 12 ) at a third, intermediate, cross-sectional area ( 70 C) when the flight speed of the gas turbine engine ( 10 ) is greater than the first predetermined value and the power produced by the gas turbine engine ( 10 ) is less than the second predetermined value.

Attenuation of FJ44 Turbofan Engine Noise with a Foam-Metal Liner Installed Over-the-Rotor

Abstract: A Williams International FJ44-3A 3000-lb thrust class turbofan engine was used as a demonstrator for a Foam-Metal Liner (FML) installed in close proximity to the fan. Two FML designs were tested and compared to the hardwall baseline. Traditional single degree-of-freedom liner designs were also evaluated to provide a comparison. Farfield acoustic levels and limited engine performance results are presented in this paper. The results show that the FML achieved up to 5 dB Acoustic Power Level (PWL) overall attenuation in the forward quadrant, equivalent to the traditional liner design. An earlier report presented the test set-up and conditions.

Heat regeneration for a turbofan, a Velarus Propulsion

Abstract: The invention adds details and alternate design supplementing the concept established with my patent application Ser. No. 12/013,431, Aircraft Propulsion System (APS). The APS ultimate fuel economy objectives requires long term design development, and this invention compromises some fuel economy for the expediency of short term implementation of a heat regeneration for turbofans via the re-arrangement of existing components and a few unique items readily designed. While this Velarus Propulsion (VPx) attains only 42% fuel economy, it retains the original APS fundamental architecture implementing heat regeneration for a turbofan engine, as well as the additional benefits of noise and emission abatement. This invention consists of the three APS technologies as follows: a) A novel arrangement of the power generation core features the turbine exhaust entering directly into the thrust chamber, thus providing heat regeneration with an appropriate configuration of the thrust chamber. b) The design of a modified combustor introduces the concept of a supersonic nozzle driving the turbines which allows for greater combustor's chamber temperature while injecting the gases at temperatures acceptable to current turbine blade metallurgy. This feature increases the engine thermal efficiency. c) The hub design allows for two options, a simpler fixed fanblade design or an advanced controllable pitch fan blade, increasing the mission and performance flexibility of a given turbofan size. In addition, this invention adds functionality to the aft cone, such as debris purge, accessories installation, and thrust reversers free from hot gases.

Modelling of the Performance of a F100 -PW229 Equivalent Engine under Sea -level Static Condition s

Abstract: The behaviour of the Pratt and Whitney F100 -PW229 low -bypass turbofan engine has been employed as the benchmark in this investigation of the attributes of a novel engine concept and in particular the potential of a two -combustor engine for military fighter application s. Because of the lack of complete commercially -classified information concerning the F100 -PW229 engine , an engine ( designated the F100 -EQ) with performance closely approximating to that of F100 -PW229 engine is modelled. This paper will discuss the modelling and simulation of the F100 -EQ engine. Generic components ’ characteristics are described and the assumptions made in order to model the F100 -EQ engine are stated . All the analysis has been undertaken using published data taken from the open literature. Nomenclature Bi metal = Metal’s Biot -number Bi tbc = Thermal -barrier coating’s Biot -number H = Flight altitude (km) Kcomb = Combustion -pattern factor Kcool = Cooling -flow factor M0 = Flight Mach -number SFC = Specific fuel -consumption rate (kg s -1 N -1 ) ST = Specific thrust (N s kg -1

Low Pressure System Component Advancements and its Impact on Future Turbofan Engine Emissions

Abstract: Within the European research project EnVIronmenTALly Friendly Aero Engines, VITAL, a number of low pressure system component technologies are being investigated. The emerging progress will allow the design of new power plants capable of providing a step change in engine fuel burn and noise. As part of the VITAL project a Technoeconomic, Environmental and Risk Assessment tool, the TERA2020, is being developed. Within this tool, means to assess the impact of component technology progress on the engine/aircraft system level has been implemented. Sensitivities relating parameters traditionally used to describe component performance, such as allowable shaft torque, low pressure turbine stage loading, fan blade weight and system level parameters have previously been published. The current paper makes an assessment of the impact of failing to deliver specific technology advancements, as researched under the VITAL project. The impact has been quantified, in terms of power plant noise and CO2 emissions.

Design and Construction of a Two Shaft Test Turbine for Investigation of Mid Turbine Frame Flows

Abstract: In order to reduce the specific fuel consumption and consequently to minimise the CO2 emission of aircraft turbine engines, development and integration of a contra-rotating open rotor architecture and the geared turbofan concept are promising alternatives to the classical turbofan engines. The large dimensions of future contra-rotating fans lead to lower rotational speed and furthermore a bigger size of the powering low pressure stages. Therefore the flow leaving the high pressure turbine (HPT) has to be guided to the low pressure turbine (LPT) inlet at larger diameter, if possible without interference or separation. To minimise weight and costs of the aircraft engine the flow diverting mid turbine frame (MTF) has to be designed preferable short. Except for flow redirection this intermediate duct has to transfer the forces from the turbine bearings of both shafts to the turbine casing and further to the engine mount. Therefore the flow channel has to be equipped with thick rigid struts. These struts may also be used to accelerate the flow and take the function of the inlet vanes. In that case this so called integrated concept helps to reduce engine weight and length. One goal of the EU project DREAM is to analyse the flow through such a MTF with turning struts and a downstream arranged counter rotating LPT. The investigation of these complex interrelationships needs a test facility with engine representative conditions. To perform these investigations the continuously operating transonic test turbine facility (TTTF) at Graz University of Technology has been adapted. This test setup consists of a redesigned HPT, a new developed single-stage LPT and a turning mid turbine frame (TMTF). The aerodynamic design of the first setup was performed by MTU Aero Engines. The shafts of both turbines are mechanically independent, so the test rig allows a realistic two shaft turbine operation. To examine the highly 3-dimensional flow through the TMTF and the downstream LPT detailed measurements will be performed with conventional measurement techniques (temperature and pressure rakes, static pressure taps), oil flow visualisation as well as with 5-hole-probes for steady and fast response aerodynamic probe (FRAP) for unsteady measurements. Optical and acoustical measurements are planned to be used in following projects.

Distributed Control of Turbofan Engines

Abstract: : The purpose of this paper is to develop control theoretic concepts for distributed control of gas turbine engines, and develop a dynamic engine model incorporating distributed components in compressor dynamics, engine cycles, and engine control. The latest results in distributed control combined with adaptive control theory are extended for turbofan engine distributed control. Concepts and architectures for distributed control are developed that create tangible benefits from the distribution of closed-loop feedback around the engine.

Inventing the F-35 Joint Strike Fighter

Abstract: The F-35 Joint Strike Fighter is a single aircraft developed to meet the multirole fighter requirements of the US Air Force, Navy, Marine Corps, and our allies. The Air Force variant is a supersonic, single engine stealth fighter. The Navy variant has a larger wing and more robust structure in order to operate from an aircraft carrier, while the Marine Corps variant incorporates an innovative propulsion system that can be switched from a turbofan cycle to a turboshaft cycle for vertical take off and landing. This novel propulsion system enabled the X-35 demonstrator to become the first aircraft in history to fly at supersonic speeds, hover, and land vertically. The F-35 program grew out of a design study of a supersonic replacement for the AV-8 Harrier, through the absorption of several other tactical aircraft initiatives. It became an international program with engineers from half a dozen countries developing a replacement for multiple aircraft Types.