Showing papers on "Turbofan published in 2012"

Gas turbine engine in-board cooled cooling air system

Abstract: A system for supplying turbine cooling air flow includes a turbofan engine, a heat exchanger, and a door. The turbofan engine includes an engine case that has an inner volume within which at least a gas turbine engine is mounted, and a bypass flow passage that is defined by an outer fan duct and an inner fan duct and that is configured to direct fan air flow therethrough. The heat exchanger is disposed within the turbofan engine, is coupled to receive fluid and cooling air from the bypass flow passage, and is configured to transfer heat between fluid and the cooling air. The door is movably mounted in the turbofan engine and is movable between a closed position, in which the cooling air will not flow through the heat exchanger, and an open position, in which the cooling air may flow through the heat exchanger.

Feasible optimum design of a turbocompound diesel brayton cycle for diesel-turbo-fan aircraft propulsion

Abstract: This paper is conceived to optimize the design of the thermodynamic cycle of a turbine (Brayton cycle) that uses a modern common rail diesel engine as an “active” combustion chamber. In this case the “active” combustion chamber produces the mechanical energy that drives the fan. The incoming air is compressed by the compressor, then is cooled (aftercooler) and inputted in the diesel engine. A high pressure common rail system optimizes the combustion in the diesel combustion chamber and the expansion begins inside the diesel engine. At the exhaust of the combustion chamber a turbine completes the expansion of the hot gases. A nozzle accelerates the exhaust from the turbine to increase the overall thrust. The mechanical energy from the diesel and from the turbine engizes the compressor and the fan. The system can be seen as a turbocharged diesel engine with the turbocharger that outputs energy to the turbofan, increasing the output power and or the efficiency. A diesel-turbine compound can be realized in this way. The coupling of the two system may be obtained in several different ways. The simplest is to put on the same shaft the compressor, the diesel crankshaft and the turbine. In front of the compressor a speed reducer drives the fan. A second example is to connect the turbine and the diesel on to electric generators. Electric engines are connected to the compressor and to the fan. The traditional turbo-diesel has the compressor coupled to the turbine, and the diesel engine that moves the fan. In this latter case, however, the turbine does not energize the fan. Many other hybrid and non hybrid solution are possible. The problem is to optimize temperatures, pressures and rpm to the different machines that form the compound. The availability of many experimental data for diesel and turbines makes it possible to obtain a design of a “true” feasible optimum Diesel-Brayton cycle. The high efficiency justifies the huge manufacturing and development costs of these turbocompound engines.

Turbine section of high bypass turbofan

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Engine Bearing Support

Abstract: A turbofan engine comprises an engine case. A gaspath extends through the engine case. A fan has a circumferential array of fan blades. A fan case encircles the fan blades radially outboard of the engine case. A plurality of fan case vanes extend outward from the engine case to the fan case. A front frame assembly includes a plurality of vanes extending radially across the gaspath. A transmission couples a shaft to a fan shaft to drive the fan. A bearing assembly couples the shaft to the front frame assembly. A bearing support extends aftward and radially inward from the front frame assembly to the bearing assembly.

Flight Mechanics Modeling of the PrandtlPlane for Conceptual and Preliminary Design

Abstract: The conceptual and preliminary design of a 300-passenger box-wing aircraft configuration, designated the PrandtlPlane, is investigated. Currently there are still a number of technical issues which must be investigated thoroughly to demonstrate the feasibility of this configuration. This research study is focused on two aspects of the PrandtlPlane design, (1) the propulsion system and (2) the flight control system. A nonlinear aircraft model is created with an in-house developed flight mechanics toolbox, which is designed for its application in the conceptual and preliminary design phase. The resulting propulsion system design has two conventional turbofan engines at the tail of the aircraft. For a large version of the PrandtlPlane, it might be beneficial to consider large open-rotor systems underneath the rear wing. The volume of the wing system, which is smaller than that of conventional aircraft, poses constraints on the fuel system design. Flight control of the PrandtlPlane is quite different from the control of conventional aircraft. If control surfaces are placed on the front and rear wings, then a pure moment can be created by differential deflection of these controls. Furthermore, a combined deflection of the front and rear wing control surfaces allows the use of direct lift control. The aircraft exhibits good inherent handling qualities in the longitudinal axis. The Dutch roll mode is slightly unstable. Improvements are expected if the vertical tails of the aircraft are redesigned. Finally, a model-based inversion flight control law is presented which provides a rate command response type in all axes. An additional outer control loop is designed which provides direct lift control. The control law is tested on the nonlinear aircraft model and demonstrates the potential of the PrandtlPlane control characteristics.

Gas turbine engine with low fan pressure ratio

Abstract: A turbofan engine includes a fan variable area nozzle axially movable relative to the fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation

Fan Rotor Support

Abstract: A turbofan engine comprises an engine case. A gaspath extends through the engine case. A fan has a circumferential array of fan blades. A fan case encircles the fan blades radially outboard of the engine case. A plurality of fan case vanes extend aftward and outward from the engine case to the fan case. A forward frame comprises a plurality of vanes radially across the gaspath. A torque box couples the fan case vanes to the forward frame. A transmission couples a shaft to a fan shaft to drive the fan. A fan bearing assembly couples a stationary forward hub structure to the fan shaft.

Experimental and Numerical Investigation of a Contra Rotating Open-Rotor Flowfield

Abstract: Contrarotating open rotor propulsion systems have seen renewed interest as a possible economic and environmentally friendly powerplant for future transport aircraft. While the potential efficiency benefits are well accepted, concerns persist regarding the probable rotor-to-rotor interaction-driven noise penalty this type of engine would have in comparison to modern ducted turbofan engines. This paper presents results of a collaborative experimental and numerical study to quantify and study in-depth the complex flowfield of a generic contrarotating open rotor model at wind-tunnel scale. The model has 10 front blades and 8 aft blades, with blade design similar to modern propellers for high-disk loadings. The comparison of flow visualization results obtained through the use of modern stereoscopic particle image velocimetry and unsteady Reynolds-averaged Navier–Stokes simulations helps to improve understanding of the interactions of front-rotor-blade wakes and tip vortices with the aft rotor, which is an important aspect to guide the design of future efficient and quiet contrarotating open rotor engines. The generally good match between the experimental and numerical slipstream results gives confidence in the utility for their analysis capabilities in this field.

Design of an Economical Counter Rotating Fan: Comparison of the Calculated and Measured Steady and Unsteady Results

Abstract: Within the framework of the EU funded Project VITAL, SNECMA (Group Safran), as the work package leader, developed a counter rotating low-speed fan-concept for a high bypass ratio engine. The detailed aerodynamic and mechanical optimization of one blading version (CRTF2.b) was carried out at the German Aerospace Center (DLR), by applying one of the newest design methods featuring a multi-objective automatic optimization method based on an Evolutionary Algorithm [1].The final design goals were high efficiency, a sufficient stall margin and adequate acoustic performances for the given cycle parameters. The fan stage developed was tested in an anechoic test facility at CIAM in Moscow. The test routine included the measurement of the performance map based on total pressure and total temperature measurements at the inlet and the outlet of the test rig and acoustic measurement as well.The unsteady flow field of the low speed Contra-Rotating Turbo Fan has been measured with four hot-wire probes at different axial positions.In the evaluation the measured data are compared with high resolution CFD results. Special emphasis was given to the comparison of the radial distribution of total pressure and total temperature in the bypass channel, the comparison of the measured and the calculated fan maps and to the comparison of the hot-wire measurements with high resolution, unsteady CFD results. The tests and the URANS-results confirmed the design goals.Copyright © 2012 by ASME

Turboelectric Distributed Propulsion System Modelling for Hybrid-Wing- Body Aircraft

Abstract: () In 2005, NASA released plans of next generation commercial airplane for 2030, with a crossdisciplinary effort on: reduced fuel consumption, aviation reliability, fundamental noise reduction and shorter take-off length. Meeting these requirements will need a fundamental shift in aircraft and engine design. Turboelectric distributed propulsion system was chosen to achieve these targets. Different from traditional turbofan, distributed propulsion system employs a large number of fans embedded on upper surface of the airframe and two turbogenerators at wing tip. This novel configuration benefits from boundary layer ingestion and distributed fans to achieve higher bypass ratio but lower fuel burn. The N3-X hybrid-wing-body is used as a baseline aircraft for the study. This paper gives basic simulation methods, as well as computational models for turboelectric distributed propulsion system. Initially, a boundary layer ingesting model has been built from computational results for embedded propulsor at different inlet conditions. In a further step, a weight estimation model of propulsors was concluded to estimate propulsors’ weight and size. Then, thermal cycle model was built to calculate engine’s performance at both design point and off design conditions. Finally, effects of boundary layer ingestion on the propulsion system were examined. The boundary layer ingesting model showed mass-average inlet pressure and Mach number are function of flight Mach number and fan inlet mass flow, on the N3-X airframe. The weight estimation model shows the overall system weight decreased with increased number of propulsors, which also caused total inlet width of propulsors increasing. So for a given total inlet width, the propulsor should be used as many as possible to reduce weight. Thermal cycle results show that fan shaft speed should be chosen as high as possible before reaching the fan tip speed limitation, and fan pressure ratio (FPR) between 1.3 and 1.35 yields minimum thrust specific fuel consumption (TSFC) at the aerodynamic design point. A fan pressure ratio of 1.3 is chosen for its potential effects on noise control. In the end, a turboelectric distributed engine was simulated to satisfy NASA N+3 subsonic commercial airplane goals.

User's Guide for the Commercial Modular Aero-Propulsion System Simulation (C-MAPSS): Version 2

Abstract: This report is a Users Guide for version 2 of 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 v.2 has some enhancements over the original, including three actuators rather than one, the addition of actuator and sensor dynamics, and an improved controller, while retaining or improving on the convenience and user-friendliness of the original. C-MAPSS v.2 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.

Modeling Contra-Rotating Turbomachinery Components for Engine Performance Simulations: The Geared Turbofan With Contra-Rotating Core Case

Abstract: This paper presents a method of modelling contra-rotating turbomachinery components for engine performance simulations. The first step is to generate the performance characteristics of such components. In this study, suitably modified one-dimensional mean line codes are used. The characteristics are then converted to three-dimensional tables (maps). Compared to conventional turbomachinery component maps, the speed ratio between the two shafts is included as an additional map parameter and the torque ratio as an additional table. Dedicated component models are then developed that use these maps to simulate design and off-design operation at component and engine level.Using this approach, a performance model of a geared turbofan with a Contra-Rotating Core (CRC) is created. This configuration was investigated in the context of the European program NEWAC (NEW Aero-engine core Concepts). The core consists of a seven-stage compressor and a two-stage turbine without inter-stage stators and with successive rotors running in opposite direction through the introduction of a rotating outer spool. Such a configuration results in reduced parts count, length, weight and cost of the entire HP system. Additionally, the core efficiency is improved due to reduced cooling air flow requirements.The model is then coupled to an aircraft performance model and a typical mission is carried out. The results are compared against those of a similar configuration employing a conventional core and identical design point performance. For the given aircraft-mission combination and assuming a 10% engine weight saving when using the CRC arrangement over the conventional one, a total fuel burn reduction of 1.1% is predicted.Copyright © 2012 by ASME

Technologies for jet noise reduction in turbofan engines

Abstract: Turbofan engines are commonly used for commercial transport due to their advantages of higher performance and lower noise. Jet noise is one of the principal noise sources of turbofan aeroplane engines and remains an acute environmental problem that requires advanced solutions. The ever-increasing demand for quieter engines requires the exploration of alternative techniques that could be used by themselves or in conjunction with existing methods. Significant progress continues to be made with noise reduction for turbofan engines. Analytical and semiempirical models have been developed to investigate the influence of some design tools when they are employed in a multidisciplinary optimisation framework. This paper discusses the major components of jet noise in turbofan engines and presents a review of jet noise reduction technologies.

The Fifth compression strategy for hypersonic aircar: proposed by three experiments of shock-tube, computation, and thought

Abstract: In our previous reports and patents, a single lightweight engine capable of operating over a wide range of Mach numbers from startup to the hypersonic regime was proposed for aircars, aircrafts, and spaceships. A new compression system of colliding super multijets with pulsation was proposed for this purpose. The new compression system essentially differs from those for the traditional four types of engines with piston, turbofan, ran-scram, and pulse-detonation. This is the fifth compression principle. Shocktube experiments and computational fluid dynamics with a chemical reaction model clarifies a large potential and stability of this system. This ultimate engine system can be extended with a special piston and scram jet systems to achieve an improved fuel consumption rate at various situations between the ground and the space, while maintaining a very low noise level with silent detonation. The present engine system will also solve the problem of the buzz at highersonic conditions.

Air-cooled oil cooler for turbofan engine

Abstract: A turbofan gas turbine engine comprises a nacelle cowl and a core engine. A bypass duct is between an outer surface of a casing of the core engine, and an inner surface of the nacelle cowl. An air channel is in the nacelle cowl, an inlet and an outlet of the air channel being in an outer surface of the nacelle cowl. An oil cooler has at least one oil passage for oil circulation, the air cooler having a first heat exchange surface in the air channel exposed to air circulating in the air channel, the air channel having a second heat exchange surface in the bypass duct exposed to air circulating in the bypass duct. A method for cooling oil in a turbofan gas turbine engine is also provided.

Far-Field Sound Radiation Due to an Installed Open Rotor:

Abstract: Future single rotation propeller and contra-rotating advanced open rotor concepts promise a significant fuel efficiency advantage over current generation turbofan engines. The development of rotors which produce a minimum level of noise is a critical technical issue which needs to be resolved in order for these concepts to become viable aircraft propulsors. Noise and emissions are subject to stringent legislative requirements, thus accurate models are required in order to predict the noise radiated from aircraft engines. In this article, the development of a theoretical model to predict noise levels of an installed open rotor is reported. First a canonical problem is examined: how to predict the pressure field produced by a rotating ring of point sources adjacent to a rigid cylinder. Analytic expressions for the far-field pressure from a rotating ring of single-frequency monopole and dipole point sources, located near an infinitely long rigid cylinder, immersed in a constant axial mean flow, are explicitl...

Geared turbofan with distributed accessory gearboxes

Abstract: A disclosed gas turbine engine includes a core engine defined about an engine centerline axis, the core engine including a compressor section driven through a shaft by a turbine section. A core nacelle surrounds the core engine and a fan nacelle is mounted at least partially around the core nacelle to define a fan bypass flow path for a bypass airflow. A fan section disposed within the fan nacelle is driven by the turbine section of the core engine through a geared architecture. An engine pylon supports the core nacelle and the fan nacelle. A towershaft is driven by the shaft of the core engine and drives a generator mounted within the core nacelle. The generator powers an electric motor mounted within the engine pylon. The electric motor drives a plurality of accessory components that are also mounted within the engine pylon.

Fan Response to Total Pressure Distortions Produced by Boundary Layer Ingesting Serpentine Inlets

Abstract: Boundary layer ingesting serpentine inlets can lead to substantial flow distortion reaching embedded engines. Total pressure distortions as a result of boundary layer ingestion couple with secondary flows produced by duct o↵set to impose a challenging operating environment for fan BLI propulsion systems. This paper presents results of an experimental investigation of the impacts of total pressure distortions on the performance of a turbofan engine. Measurement of the dynamic response of a Pratt & Whitney JT15D-1 fan to screen-produced total pressure distortions patterned after BLI flow patterns were made in a test facility designed and constructed for the purpose at the Virginia Tech Turbomachinery and Propulsion Research Laboratory. The impact of total pressure distortions on fan pressure ratio and wake production was determined using high-frequency and steady 5-hole probe measurements of the fan blade wakes with the fan subjected to the NASA “Inlet A” total pressure profile. Measurements behind the fan indicate that response is governed by the blade time response to the local variation in flow conditions. Changes in the relative blade incidence angle result from reduced inlet velocity and swirl produced in the distorted region. The change in incidence angle leads to a complex dynamic flow response within the blade passages. Variation in turning, wake thickness, and turbulence production alter the performance of the fan and reduce eciency. The complex cycle through which a blade passes during each rotation is identified and described.

Turbofan engine bearing support

Abstract: A turbofan engine (20) comprises an engine case (22). A gaspath extends through the engine case (22). A fan (42) has a circumferential array of fan blades (100). A fan case (40) encircles the fan blades (100) radially outboard of the engine case (22). A plurality of fan case vanes (44) extend outward from the engine case (22) to the fan case (40). A front frame assembly (172) includes a plurality of vanes (173) extending radially across the gaspath. A transmission (46) couples a shaft (25) to a fan shaft (132) to drive the fan (42). A bearing assembly (176) couples the shaft (25) to the front frame assembly (172). A bearing support (454) extends aftward and radially inward from the front frame assembly (172) to the bearing assembly (176).

The UltraFan Engine and Aircraft Based Thrust Reversing

Abstract: Advances in propulsion technology have been the heart of aircraft performance improvement for the past half century. Structural weight reductions and subsystem efficiencies have only provided modest performance gains. Near term NextGen engine efficiency improvements of about 12-15% are evident by the development of the Pratt and Whitney PW1000G geared turbofan, and the CFM LeapX, both very high bypass (VHBR) turbofans. These engines, in the 25-30K thrust class, and their larger counterparts, the GE90, Trent 1000 etc. can achieve thrust specific fuel consumption TSFC in the mid 0.5s. Analysis shows that further increases in fan diameter, bypass ratio and propulsive efficiency are limited by nacelle weight and drag. Propeller systems offer the highest propulsive efficiency but have transonic speed limitations and pose acoustic challenges, especially counterrotating systems. Turboprop systems have been limited to short haul regional markets and the military. This paper explores a turbofan engine concept that combines the performance benefits of a propeller with the transonic speed and acoustic attenuation benefits of a turbofan. The UltraFan Engine (UFE) is able to maximize transonic propulsive efficiency while capitalizing on the acoustic attenuation benefits of a nacelle. The nacelle based cascade thrust reverser has been eliminated from the engine, removing a significant constraint on bypass ratio, allowing the achievement of ultrahigh bypass ratios. Bypass ratios exceeding 30 may be achievable, allowing for significant improvements in transonic propulsive efficiency. Net fuel burn reduction remains impressive even when considering the weight penalty for incorporating the thrust reverser into the aircraft.

System and method for mounting an aircraft engine

Abstract: A mounting system and method capable of reducing backbone deflection in a high-bypass turbofan engine. The system includes a rigid structure and a linkage mechanism having at least first and second links that are each pivotally connected to the rigid structure and adapted to be pivotally connected to an engine support structure of the aircraft. The first and second links are configured to define a focal point thereof at a location that is a distance from a centerline of the engine of not more than 15% of an inlet diameter at an inlet of the engine, and is located aft of a vector of an inlet load to which the engine is subjected when the aircraft is in a climb maneuver. The location of the focal point is such that a moment of a thrust load of the engine and a moment of the inlet load oppose each other, thereby reducing backbone bending of the engine during the climb maneuver.

Estimation of exhaust gas temperature using artificial neural network in turbofan engines

Abstract: This paper deals with the estimation of exhaust gas temperature (EGT) of a CFM56-7B turbofan engine using artificial neural network (ANN) at two different power settings, maximum continuous and take-off. The study was carried out using the operational parameters of the engine such as net thrust, fuel flow, low rotational speed, core rotational speed, pressure ratio, fan air inlet temperature, take-off margin temperature, and thrust specific fuel consumption. All these data are taken from test cell measurements during ground operating of the engines. In this study, the accuracy of ANN results is compared with the measurements and the results of a regression analysis earlier based multiple linear method. The comparison of the predictions of the models indicates that ANN is capable of accurately predicting EGT in used turbofan engines. The correlation between the exhaust gas temperature and the operational parameters of the engine was found to be 0.99 and 0.99 for training data and to be 0.90 and 0.97 for test data using ANN at two different power settings, maximum continuous and take-off, respectively. For both investigated power settings, maximum continuous and take-off, the mean absolute errors were found to be 2.1 per cent and 5.08 per cent, while the coefficients of variance of root mean square error were found to be 0.5705 and 0.3539, respectively. The results obtained from ANN models show good agreement with ground measurements and the regression models. Finally, we believe that ANN can be used for prediction of EGT as a predictive tool in this sort of application.

Low hub-to-tip ratio fan for a turbofan gas turbine engine

Abstract: A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (R HUB ) is the radius of the leading edge at the hub relative to a centerline of the fan. A tip radius (R TIP ) is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (R HUB /R TIP ) is at least less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than or equal to 0.25.

Gas turbine engine having fan rotor driven by turbine exhaust and with a bypass

Abstract: A gas turbine engine has a core engine incorporating a core engine turbine. A fan rotor is driven by a fan rotor turbine. The fan rotor turbine is in the path of gases downstream from the core engine turbine. A bypass door is moveable from a closed position at which the gases from the core engine turbine pass over the fan rotor turbine, and moveable to a bypass position at which the gases are directed away from the fan rotor turbine. An aircraft is also disclosed.

Active Fault Tolerant Model Predictive Control of a Turbofan Engine using C-MAPSS40k

Abstract: Aircraft engine control is a crucial component for the safe and stable operation of gas turbine engines which are complex nonlinear systems. As engines have evolved to higher capabilities it is crucial to update the control strategy to ensure maximum functionality of the engine. Current industrial baseline controllers are based in the Proportional-Integral-Derivative (PID) control scheme along with individual limit controllers having critically damped responses housed in the min-max architecture. In light of the distributed engine control architecture that exploits digital electronics and hence higher on-board computational capabilities, the baseline controller is replaced by a Model Predictive Control (MPC) law with on-line optimization. MPC is a model based control technique that can handle complex constrained dynamics thus allowing the incorporation of component faults in the design process of the controller. Component faults occur during an engine's operation mainly due to fan blade-shroud rubbing, structural wear and tear and foreign object ingestion thus affecting the engine performance. Simulations on the Linear Time Invariant (LTI) as well as the nonlinear turbofan engine of the Commercial Modular Aero-Propulsion System Simulation (C-MAPSS40k) tool are carried out. In the presence of a component fault, active fault tolerant control using the multi-model MPC approach is applied by switching between the MPC blocks, each using its respective LTI reference model.The control of both the fan speed as well as the thrust for a demand profile in the Power Level Angle (PLA) is investigated and the MPC performance is compared with that of the PID controller demonstrating the successful replacement of the baseline controller with an on-line fault tolerant MPC. The thrust control approach using MPC consumes lesser fuel when compared with the fan speed control approach.

Civil turbofan engines thrust generic model

Abstract: Engine model is one of the most important items in aircraft simulation, because all aircraft stability parameters, as well as its performance in climbing, and accelerated flights, depend on the propulsive force developed by the power plant.

Full-Scale Turbofan-Engine Turbine-Transfer Function Determination Using Three Internal Sensors

Abstract: Existing NASA/Honeywell EVNERT full-scale static engine test data is analyzed by using source-separation techniques in order to determine the turbine transfer of the currently sub-dominant combustor noise. The results are used to assess the combustor-noise prediction capability of the Aircraft Noise Prediction Program (ANOPP). Time-series data from three sensors internal to the Honeywell TECH977 research engine is used in the analysis. The true combustor-noise turbine-transfer function is educed by utilizing a new three-signal approach. The resulting narrowband gain factors are compared with the corresponding constant values obtained from two empirical acoustic-turbine-loss formulas. It is found that a simplified Pratt & Whitney formula agrees better with the experimental results for frequencies of practical importance. The 130 deg downstream-direction far-field 1/3-octave sound-pressure levels (SPL) results of Hultgren & Miles are reexamined using a post-correction of their ANOPP predictions for both the total noise signature and the combustion-noise component. It is found that replacing the standard ANOPP turbine-attenuation function for combustion noise with the simplified Pratt & Whitney formula clearly improves the predictions. It is recommended that the GECOR combustion-noise module in ANOPP be updated to allow for a user-selectable switch between the current transmission-loss model and the simplified Pratt & Whitney formula. The NASA Fundamental Aeronautics Program has the principal objective of overcoming today's national challenges in air transportation. The Subsonic Fixed Wing Project's Reduce-Perceived-Noise Technical Challenge aims to develop concepts and technologies to dramatically reduce the perceived aircraft noise outside of airport boundaries. The reduction of aircraft noise is critical to enabling the anticipated large increase in future air traffic.

Model-Based Control of an Aircraft Engine using an Optimal Tuner Approach

Abstract: This paper covers the development of a model-based engine control (MBEC) method- ology applied to an aircraft turbofan engine. Here, a linear model extracted from the Commercial Modular Aero-Propulsion System Simulation 40,000 (CMAPSS40k) at a cruise operating point serves as the engine and the on-board model. The on-board model is up- dated using an optimal tuner Kalman Filter (OTKF) estimation routine, which enables the on-board model to self-tune to account for engine performance variations. The focus here is on developing a methodology for MBEC with direct control of estimated parameters of interest such as thrust and stall margins. MBEC provides the ability for a tighter control bound of thrust over the entire life cycle of the engine that is not achievable using traditional control feedback, which uses engine pressure ratio or fan speed. CMAPSS40k is capable of modeling realistic engine performance, allowing for a verification of the MBEC tighter thrust control. In addition, investigations of using the MBEC to provide a surge limit for the controller limit logic are presented that could provide benefits over a simple acceleration schedule that is currently used in engine control architectures.