Showing papers on "Turbofan published in 2004"

Turbofan case and method of making

Abstract: A casing for a gas turbine includes a fan case, an intermediate case and a gas generator case integrated with one another. In another aspect, the casing provides a construction including several aspects which improve structural efficiency, such as a semi-monocoque construction, improved strut design, etc. Improved load paths and means for transmitting loads in the engine case are also disclosed.

New Method for Jet Noise Reduction in Turbofan Engines

Abstract: A new method for reducing large-scale mixing noise from dual-stream jets is presented. The principle is reduction of the convective Mach number of turbulent eddies that produce intense downward sound radiation. In a jet representing the coaxial exhaust of a turbofan engine, this is achieved by tilting downward, by a few degrees, the bypass (secondary) plume relative to the core (primary) plume. The misalignment of the two flows creates a thick low-speed secondary core on the underside of the high-speed primary flow. The secondary core reduces the convective Mach number of primary eddies, thus hindering their ability to generate sound that travels to the downward acoustic far field. Tilting of the bypass stream is possible by means of fixed or variable vanes installed near the exit of the bypass duct. Subscale aeroacoustic experiments simulated the exhaust flow of a turbofan engine with bypass ratio 6.0. Deflection of the bypass stream resulted in suppression of the peak overall sound pressure level by 4.5 dB and the effective perceived noise level by 2.8 dB. For the nozzle configuration used, the thrust loss is estimated at around 0.5% with the vanes activated and 0.15% with the vanes deactivated.

High performance VTOL convertiplanes

Abstract: A single-tilt-rotor VTOL airplanes have a tiltable rotor attached to an elongated power pod containing the collective and cyclical pitch mechanism, and transmission. The power pod is pivotably attached to a base that is slidably mounted on a pair of slotted guide beams attached on top of the roof of the fuselage. The guide beams run longitudinally from the front of the aircraft to past the center of gravity (CG) of the aircraft in order to transport the power pod from the front section to the center section when converting from the horizontal cruising mode to the VTOL mode. In the horizontal cruising mode, the power pod perched horizontally on top of the fuselage front section with sufficient clearance for the rotor to rotate in front of the aircraft. Upon transitioning to the VTOL mode, a telescopic actuator is used to pivot the power pod vertically while a cable-winch system is used to move the entire power pod and base assembly rearwardly to stop at the center of gravity of the aircraft, and vice versa, thus allowing the power pod to travel significantly rearward and forward as required for proper balancing of vertical lift as the power pod pivots 90 degrees during transition from VTOL mode to the cruising mode. A single piston engine, or a single or pair of turbofan engines, mounted slightly to the rear of the CG, have drive shafts that can be clutched and mated onto respective receiving shaft from the transmission within the power pod in order to power the tiltable rotor. The engine is also attached to a propeller for horizontal propulsion, or if turbofan engines are used, jet thrust is generated for horizontal cruise. A small anti-torque rotor or ducted fan toward the tail of the aircraft is mechanically coupled to the engine via a drive shaft to provide the necessary side-way thrust to overcome the main rotor's torque. In the horizontal cruising mode, the tiltable rotor is allowed to windmill slowly at a minimum rotational speed necessary to maintain the integrity of the rotor blades. The same propulsion principle can be applied to VTOL airplanes having more than one tiltable rotor, thereby can potentially increase the speed, range and reliability of current twin-wing-mounted-tilt-rotor aircraft. A pair of high-aspect-ratio wings on both sides of the fuselage provide highly efficient lift during cruising flight with very little induced drag. Conventional horizontal and vertical tail planes are used for directional stability in the cruising mode.

Method and apparatus for cleaning a turbofan gas turbine engine

Abstract: Device for cleaning a gas turbine engine (2), and in particular an engine of turbofan type. The present invention further relates to a method for cleaning such a engine. The device comprises a plurality of nozzles (31, 33, 35) arranged to atomize cleaning liquid in the air stream in an air inlet (20) of the engine (2) up-stream of a fan (25) of the engine (2). According to the invention a first nozzle (31) is arranged at a position such that the cleaning liquid emanating from the first nozzle (31) impinges the surfaces of the blades (40) substantially on the pressure side (53); a second nozzle (35) is arranged at a position such that the cleaning liquid emanating from the second nozzle (35) impinges the surfaces of the blades (40) substantially on the suction side (54); and a third nozzle (33) is arranged at a position such that the cleaning liquid emanating from the third nozzle (33) passes substantially between the blades (40) and enters an inlet (23) of the core engine (203). Thereby, the different types of fouling found on the fan and in the core engine compressor of turbofan engine can be removed in an efficient manner.

Triple circuit turbine cooling

Abstract: A turbofan engine includes in serial flow communication a first fan, second fan, multistage compressor, combustor, first turbine, second turbine, and third turbine. The first turbine is joined to the compressor by a first shaft. The second turbine is joined to the second fan by a second shaft. And, the third turbine is joined to the first fan by a third shaft. First, second, and third cooling circuits are joined to different stages of the compressor for cooling the forward and aft sides and center bore of the first turbine with different pressure air.

Model Adaptation and Nonlinear Model Predictive Control of an Aircraft Engine

Abstract: The performance improvement of constrained nonlinear model predictive control (NMPC) with state and parameter estimation over traditional control architectures is investigated and applied to a model turbofan aircraft engine. Strong nonlinearities are present in turbofan aircraft engines due to the large range of operating conditions and power levels experienced during a typical mission. Also, turbine operation is restricted due to mechanical, aerodynamic, thermal, and flow limitations. Current control methodologies rely strictly on a priori information; therefore they fail to utilize current engine state or health information for reducing conservatism and improving engine performance. NMPC is selected because it depends on a model that can be adapted to the current engine conditions, it can explicitly handle the nonlinearities, both input and output constraints of many variables, and determine the optimal control that will meet the requirements for any engine condition all in a single control formulation. A physics based component level model is developed as the heart of the architecture. The state or health of the engine is determined using a joint state and parameter estimator utilizing extended Kalman filter (EKF) techniques. With the necessary engine information in hand, a constrained NMPC is used to determine the optimal actuator commands. Results regarding steady state performance improvements are presented.Copyright © 2004 by ASME

Aircraft and Energy Use

Abstract: bypass ratio The ratio of air passed through the fan system to that passed through the engine core. contrail The condensation trail that forms when moist, high-temperature air in a jet exhaust, as it mixes with ambient cold air, condenses into particles in the atmosphere and saturation occurs. drag The aerodynamic force on an aircraft body; acts against the direction of aircraft motion. energy intensity (EI) A measure of aircraft fuel economy on a passenger-kilometer basis; denoted by energy used per unit of mobility provided (e.g., fuel consumption per passenger-kilometer) energy use (EU) A measure of aircraft fuel economy on a seat-kilometer basis (e.g., fuel consumption per seatkilometer). great circle distance The minimum distance between two points on the surface of a sphere. hub-and-spoke system Feeding smaller capacity flights into a central hub where passengers connect with flights on larger aircraft that then fly to the final destination. lift-to-drag ratio (L/D) A measure of aerodynamic efficiency; the ratio of lift force generated to drag experienced by the aircraft. load factor The fraction of passengers per available seats. radiative forcing A measure of the change in Earth’s radiative balance associated with atmospheric changes; positive forcing indicates a net warming tendency relative to preindustrial times. structural efficiency(OEW/MTOW) The ratio of aircraft operating empty weight (OEW) to maximum takeoff weight (MTOW); a measure of the weight of the aircraft structure relative to the weight it can carry (combined weights of structure plus payload plus fuel). thrust A force that is produced by engines and propels the aircraft. thrust specific fuel consumption (SFC) A measure of engine efficiency as denoted by the rate of fuel consumption per unit thrust (e.g., kilograms/second/Newton). turbofan engine The dominant mode of propulsion for commercial aircraft today; a turbofan engine derives its thrust primarily by passing air through a large fan system driven by the engine core.

Propagation and Decay of Shock Waves in Turbofan Engine Inlets

Abstract: Numerical experiments are carried out to investigate the tone noise radiates from a turbofan engine inlet under conditions at which the relative flow past the rotor tip is supersonic. Under these conditions, the inlet tone noise is generated by the upstream-propagating rotor-locked shock wave field. The spatial evolution of this shock system is studied numerically for flows through two basic hard-walled configurations: a slender nacelle with large throat area and a thick nacelle with reduced throat area. With the flight Mach number set to 0.25, the spatial evolution of the acoustic power through the two inlets reveals that the reduced throat area inlet provides superior attenuation. This is attributed to the greater mean flow acceleration through its throat and is qualitatively in accord with one-dimensional theory, which shows that shock dissipation is enhanced at high Mach numbers. The insertion of a uniform extension upstream of the fan is shown to yield greater attenuation for the inlet with large throat area, while the acoustic performance of the reduced throat area inlet is degraded. This occurs because the interaction of the nacelle and spinner potential fields is weakened, resulting in a lower throat Mach number. The effect of forward flight on the acoustic power radiated from the two inlets is also investigated by examining a simulated static condition. It is shown that the slender nacelle radiated significantly less power at the static condition than in flight, whereas the power levels at the two conditions are comparable for the thick nacelle. The reason for this behavior is revealed to be a drastic overspeed near the leading edge of the slender nacelle, which occurs to a lesser degree in the case of the thick inlet. This has implications for ground acoustic testing of aircraft engines, which are discussed.Copyright © 2004 by ASME

High-Fidelity Three-Dimensional Simulation of the GE90

Abstract: A full-engine simulation of the three-dimensional flow in the GE90 94B high-bypass ratio turbofan engine has been achieved. It would take less than 11 hr of wall clock time if starting from scratch through the exploitation of parallel processing. The simulation of the compressor components, the cooled high-pressure turbine, and the low-pressure turbine was performed using the APNASA turbomachinery flow code. The combustor flow and chemistry were simulated using the National Combustor Code (NCC). The engine simulation matches the engine thermodynamic cycle for a sea-level takeoff condition. The simulation is started at the inlet of the fan and progresses downstream. Comparisons with the cycle point are presented. A detailed look at the blockage in the turbomachinery is presented as one measure to assess and view the solution and the multistage interaction effects.

High-Speed Jet Noise Reduction Using Microjets on a Jet Engine

Abstract: by a jet engine. Experiments were conducted at the NASA Ames Research Center using a General Electric YJ97-GE-3 turbofan jet engine that was equipped with a 317.5 mm converging nozzle. The engine was operated at conditions that resulted in jets with fully expanded Mach numbers of 0.9 and 1.3. The microjets were generated using up to 48 evenly spaced micro-nozzles that had exit diameters of 1.2 and 2.4 mm. The operating pressure of the microjets was varied from 7.9 to 42.4 bar. Various microjet configurations were used resulting in a total mass flux of the microjets that ranged from 0.5 to 2.3 % of the primary mass flux for the subsonic jet and from 0.3 to 1.0 % of the primary mass flux for the supersonic jet. Through the various configurations it was found that reductions of up to 2 dB in the OASPL could be obtained for both the subsonic and supersonic jets. The reductions for the subsonic jet were seen at all frequencies while they were seen primarily at the higher frequencies for the supersonic jet. A reduction of about 2 dB in the shock noise of the supersonic jet was also observed.

Probabilistic neural networks for validation of on-board jet engine data:

Abstract: A method is presented for identification of faults in the readings of sensors used to monitor the performance and the condition of jet engines. Probabilistic neural networks are used to detect the presence and identify the location and magnitude of faults (biases) in sensor readings. The faults can be detected on sets comprising a limited number of instruments, typical of those available for on-board monitoring of jet engines. An engine performance model is used to support the constitution of a network. Training information is built using the model to produce data for a comprehensive set of healthy and faulty situations. The network performance in detecting and quantifying sensor faults is validated on a large number of fault cases, also generated by a model, which are used for testing the network and cover a wide range of conditions that can be encountered in practice. An engine, representative of current large civil engine designs (large bypass, partially mixed turbofan), serves as the test vehi...

Analysis of the Magnetohydrodynamic Energy Bypass Engine for High-Speed Airbreathing Propulsion

Abstract: The performance of the MHD energy bypass air-breathing engine for high-speed propulsion is analyzed in this investigation. This engine is a specific type of the general class of inverse cycle engines. In this paper, the general relationship between engine performance (specific impulse and specific thrust) and the overall total pressure ratio through an engine (from inlet plane to exit plane) is first developed and illustrated. Engines with large total pressure decreases, regardless of cause or source, are seen to have exponentially decreasing performance. The ideal inverse cycle engine (of which the MHD engine is a sub-set) is then demonstrated to have a significant total pressure decrease across the engine; this total pressure decrease is cycle-driven, degrades rapidly with energy bypass ratio, and is independent of any irreversibility. The ideal MHD engine (inverse cycle engine with no irreversibility other than that inherent in the MHD work interaction processes) is next examined and is seen to have an additional large total pressure decrease due to MHD-generated irreversibility in the decelerator and the accelerator. This irreversibility mainly occurs in the deceleration process. Both inherent total pressure losses (inverse cycle and MHD irreversibility) result in a significant narrowing of the performance capability of the MHD bypass engine. The fundamental characteristics of MHD flow acceleration and flow deceleration from the standpoint of irreversibility and second-law constraints are next examined in order to clarify issues regarding flow losses and parameter selection in the MM modules. Severe constraints are seen to exist in the decelerator in terms of allowable deceleration Mach numbers and volumetric (length) required for meaningful energy bypass (work interaction). Considerable difficulties are also encountered and discussed due to thermal/work choking phenomena associated with the deceleration process. Lastly, full engine simulations utilizing inlet shock systems, finite-rate chemistry, wall cooling with thermally balanced engine (fuel heat sink), fuel injection and mixing, friction, etc. are shown and discussed for both the MHD engine and the conventional scramjet. The MHD bypass engine has significantly lower performance in all categories across the Mach number range (8 to 12.2). The lower performance is attributed to the combined effects of 1) additional irreversibility and cooling requirements associated with the MHD components and 2) the total pressure decrease associated with the inverse cycle itself.

Engine cycle and exhaust configuration for quiet supersonic propulsion

Abstract: The thermodynamics and acoustics of a fixed-cycle, bypass ratio 3 supersonic engine with an innovative noise suppression scheme is explored. The silencing method entails installation of variable turning vanes in the bypass exhaust of a separate-flow turbofan engine. During noise-sensitive segments of flight, the vanes impart a slight downward tilt to the bypass plume relative to the core plume, thus thickening the bypass stream on the underside of the jet. This results in a reduction of the convective Mach number of instability waves that produce intense downward sound radiation. Subscale experiments show that, relative to the mixed-flow exhaust, the coaxial separate-flow exhaust with vanes reduces the peak overall sound pressure level by 8 dB and the effective perceived noise level by 7 dB. The noise-equivalent specific thrust on takeoff is reduced from 490 to 390 m/s. Compared to a current-generation low-bypass turbofan engine, the bypass ratio 3 engine is estimated to be 13 dB quieter with the mixed-flow exhaust and 20 dB quieter with the aforementioned suppression scheme. The vane configuration of this study is estimated to cause a thrust loss of 1% at takeoff and 0.25% at supersonic cruise.

Test equipment for the development of an aircraft gas turbine engine

Abstract: The equipment has a nozzle drogue (12) with a variable and adjustable section (20) and installed axially at an exhaust passage of a turbofan engine to dynamically modify a flow of an exhaust gas during a series of structure tests. The drogue is fixed below a frame (14) using an aerodynamically sectioned arm (16) and fixed directly on a driving motor. The section has flaps (22) articulated along a circular edge of a tapered body (26).

Performance Cycle Analysis of A Two-spool, Separate- exhaust Turbofan With Interstage Turbine Burner

Abstract: This paper presents the performance cycle analysis of a dual-spool, separate-exhaust turbofan engine, with an Interstage Turbine Burner serving as a secondary combustor. The ITB, which is located at the transition duct between the high- and the low-pressure turbines, is a relatively new concept for increasing specific thrust and lowering pollutant emissions in modern jet engine propulsion. A detailed performance analysis of this engine has been conducted for steady-state engine performance prediction. A code is written and is capable of predicting engine performances (i.e., thrust and thrust specific fuel consumption) at varying flight conditions and throttle settings. Two design-point engines were studied to reveal trends in performance at both full and partial throttle operations. A mission analysis is also presented to assure the advantage of saving fuel by adding ITB.

ROSAS Wind Tunnel Test Campaign Dedicated to Unconventional Aircraft Concepts Study

Abstract: An extensive acoustic wind tunnel test campaign was conducted between March and September 2003 in the frame of the European research project ROSAS to assess experimentally the noise shielding effectiveness of classic airframe components for unconventional aircraft configurations for the first time ever in Europe. A complete aircraft model (1/11th-scale) was installed in the ONERA CEPrA19 anechoic wind tunnel, successively with a fan and a jet noise simulator representing the noise sources of an advanced, high bypass ratio turbofan. Various positions of the engine with respect to the airframe were tested with noise measurements being performed in the far field. The ROSAS test campaign has allowed gathering a comprehensive database on noise installation effects for novel aircraft concepts, yet with the shortcomings of the first of its kind. Hence the effects of the noise source characteristics, of a number of geometrical parameters and of the external flow were analyzed to some extent. Significant noise attenuation was evidenced as expected, and other secondary installation effects were also studied. This paper presents the ROSAS experiment, results and preliminary analyses of the acoustic shielding phenomena.

A Comparative Study of Genetic Algorithms and Gradient Methods for RM12 Turbofan Engine Diagnostics and Performance Estimation

Abstract: Recent work on gas turbine diagnostics based on optimisation techniques advocates two different approaches: 1) Stochastic optimisation, including Genetic Algorithm techniques, for its robustness when optimising objective functions with many local optima and 2) Gradient based methods mainly for their computational efficiency. For smooth and single optimum functions, gradient methods are known to provide superior numerical performance. This paper addresses the key issue for method selection, i.e. whether multiple local optima may occur when the optimisation approach is applied to real engine testing. Two performance test data sets for the RM12 low bypass ratio turbofan engine, powering the Swedish Fighter Gripen, have been analysed. One set of data was recorded during performance testing of a highly degraded engine. This engine has been subjected to Accelerated Mission Testing (AMT) cycles corresponding to more than 4000 hours of run time. The other data set was recorded for a development engine with less than 200 hours of operation. The search for multiple optima was performed starting from more than 100 extreme points. Not a single case of multi-modality was encountered, i.e. one unique solution for each of the two data sets was consistently obtained. The RM12 engine cycle is typical for a modern fighter engine, implying that the obtained results can be transferred to, at least, most low bypass ratio turbofan engines. The paper goes on to describe the numerical difficulties that had to be resolved to obtain efficient and robust performance by the gradient solvers. Ill conditioning and noise may, as illustrated on a model problem, introduce local optima without a correspondence in the gas turbine physics. Numerical methods exploiting the special problem structure represented by a non-linear least squares formulation is given special attention. Finally, a mixed norm allowing for both robustness and numerical efficiency is suggested.Copyright © 2004 by ASME

Fluidic Chevrons for Jet Noise Reduction

Abstract: Chevron mixing devices are used to reduce noise from commercial separate-flow turbofan engines. Mechanical chevron serrations at the nozzle trailing edge generate axial vorticity that enhances jet plume mixing and consequently reduces far-field noise. Fluidic chevrons generated with air injected near the nozzle trailing edge create a vorticity field similar to that of the mechanical chevrons and allow more flexibility in controlling acoustic and thrust performance than a passive mechanical design. In addition, the design of such a system has the future potential for actively controlling jet noise by pulsing or otherwise optimally distributing the injected air. Scale model jet noise experiments have been performed in the NASA Langley Low Speed Aeroacoustic Wind Tunnel to investigate the fluidic chevron concept. Acoustic data from different fluidic chevron designs are shown. Varying degrees of noise reduction are achieved depending on the injection pattern and injection flow conditions. CFD results were used to select design concepts that displayed axial vorticity growth similar to that associated with mechanical chevrons and qualitatively describe the air injection flow and the impact on acoustic performance.

A Propulsion System Analysis of Oil Free Turbomachinery for Aviation Turbofan Engines

Abstract: A strategy for the elimination of lubricating oil systems from aviation turbofans incorporates the use of foil bearings in place of traditional rolling element bearings. A foil bearing is a hydrodynamic device that utilizes air as the working fluid to separate a rigid shaft from a compliant static (non-rotating) structure. Modern developments in foil bearing design have increased the state of the art to the level where it is now appropriate to consider such devices for application in aircraft engines. This analysis is the first published work that quantifies the benefits of oil free turbomachinery. Conceptual designs are carried out for oil free gas turbine propulsions systems, which facilitate the calculation of the improved thrust to weight ratios that these engines could achieve. These engines are then applied to 50-passenger regional jet and 10-passenger supersonic business jet models to resize the vehicles and compute system benefits. Oil Free technology was applied to both the modern state of the art engines as well as advanced technology engines, which incorporate additional feature sponsored by the NASA Ultra Efficient Engine Technology Program. The figures of merit computed in this analysis include take-off gross weight, mission fuel, mission NOx, and landing and take-off operation NOx. The current state of the art 50 passenger vehicle achieved a 3.4% reduction in TOGW while the notional supersonic business jet achieved a 2.7% reduction.

Feedback Control of a Morphing Chevron for Takeoff and Cruise Noise Reduction

Abstract: Noise from commercial high-bypass ratio turbofan engines is generated by turbulent mixing of the hot jet exhaust, fan stream, and ambient air. Serrated aerodynamic devices, known as chevrons, along the trailing edges of a jet engine primary and secondary exhaust nozzle have been shown to reduce jet noise at takeoff and shock-cell noise at cruise conditions. Their optimum shape is a finely tuned compromise between noise-benefit and thrust-loss. The design of a full scale Variable Geometry Chevron (VGC) fan-nozzle incorporating Shape Memory Alloy (SMA) actuators is described in a companion paper. This paper describes the development and testing of a proportional-integral control system that regulates the heating of the SMA actuators to control the VGC s tip immersion. The VGC and control system were tested under representative flow conditions in Boeing s Nozzle Test Facility (NTF). Results from the NTF test which demonstrate controllable immersion of the VGC are described. The paper also describes the correlation between strains and temperatures on the chevron with a photogrammetric measurement of the chevron's tip immersion.

Gas-turbine performance improvements through the use of multiple turbine interstage burners

Abstract: The focus of this investigation is to increase the power and thermodynamic efficiency of a gas-turbine engine through the use of multiple interstage burners (MIBs). Liu and Sirignano [(Liu, F., and Sirignano, W. A., "Performance Increases for Gas Turbine Engines Through Combustion Inside the Turbine," Journal of Propulsion and Power, Vol. 15, No. 1, 1999, pp. 111-118), (Liu, F., and Sirignano, W. A., "Turbojet and Turbofan Engine Performance Increase Through Turbine Burners," Journal of Propulsion and Power, Vol. 17, No. 3, 2001, pp. 695-705)] suggest burning fuel in the turbine to achieve high-pressure combustion that can significantly improve engine performance. An extension to the Liu and Sirignano research, including turbine cooling for MIB engines, was performed, and results showed large performance improvements for land-based power production engines. Unfortunately, the investigation showed minimal improvements for aircraft engines due to the large amount of cooling air required to keep turbine surface temperatures below critical limits. Thus, this investigation focused on improvements for land-based energy production engines with MIBs. Furthermore, land-based engine performance may be significantly increased by the use of open-loop or closed-loop steam cooling in conjunction with MIBs. Comparisons of the performance of these engines with open-loop air cooling, open-loop steam cooling, and closed-loop steam cooling were made. Results show that open-loop steam cooling yields the highest net power output and thermal efficiency. In addition, it was discovered that one main burner and one interstage burner, that is, two MIBs, is a good compromise for all three cooling schemes.

Control of compressor face total pressure distortion on a high bypass turbofan intake using air-jet vortex generators

Abstract: The control of turbofan intake duct flow separation and the subsequent control of compressor face distortion have the potential to achieve a relaxation of the maximum crosswind constraint at take off. Wind tunnel tests have been conducted using a 10% axissymmetric scale model. Air jets, simulating micro-electrical mechanical actuators, have been implemented at the intake lip to reenergize the boundary layer locally. Improvements in distortion of up to 40% could be achieved at a 10% engine bleed.

An investigation of innovative technologies for reduction of jet noise in medium and high bypass ratio turbofan engines

Abstract: AN INVESTIGATION OF INNOVATIVE TECHNOLOGIES FOR REDUCTION OF JET NOISE IN MEDIUM AND HIGH BYPASS TURBOFAN ENGINES By William Bryan Callender Doctor of Philosophy (Ph.D.) University of Cincinnati 2004 This research project has developed a new, large-scale, nozzle acoustic test rig capable of simulating the exhaust flows of separate flow exhaust systems in medium and high bypass turbofan engines. This rig has subsequently been used to advance the understanding of two stateof-the-art jet noise reduction technologies. The first technology investigated is an emerging jet noise reduction technology known as chevron nozzles. The fundamental goal of this investigation was to advance the understanding of the fundamental physical mechanisms responsible for the acoustic benefits provided by these nozzles. Additionally, this study sought to establish the relationship between these physical mechanisms and the chevron geometric parameters. A comprehensive set of data was collected, including far-field and near-field acoustic data as well as flow field measurements. In addition to illustrating the ability of the chevron nozzles to provide acoustic benefits in important aircraft certification metrics such as effective perceived noise level (EPNL), this investigation successfully identified two of the fundamental physical mechanisms responsible for this reduction. The flow field measurements showed the chevron to redistribute energy between the core and fan streams to effectively reduce low frequency noise by reducing the length of the jet potential core. However, this redistribution of energy produced increases in turbulent kinetic energy of up to 45% leading to a degradation of the chevron benefit at higher frequencies. Trends observed with respect to the chevron geometry showed that the chevron penetration could be matched to the exhaust flow conditions to optimally balance the trade between low frequency reduction and high frequency increase to maximize reductions in EPNL. Secondly, a completely new technology, known as fluidic injection, was investigated. This technology consists of applying continuous air injection, from a number of small injection jets, at the nozzle exit plane to reduce jet noise. The principal advantage of such an approach is that it is an active technology that can be activated as needed and, as such, may be more acceptable in aircraft engines from a performance standpoint than passive technologies. This study successfully demonstrated the feasibility of this technology by showing that effective jet noise reduction can be provided in a broad range of flow conditions using less than 1% of the mean jet mass flow. An investigation of injection geometric parameters identified the injection pitch angle as the most influential parameter with respect to jet noise reduction. Furthermore, an investigation of scaling effects showed a momentum ratio of approximately 1.5% to provide reductions in sound pressure level between 1 and 2 dB across a wide range of frequencies for a wide range of flow conditions and scales including both single stream and dual stream flows. PIV flow field measurements identified the fundamental physical mechanism of the noise reduction to be a near uniform reduction in shear layer turbulence.

Mitigation of unsteady peak fan blade and disc stresses in turbofan engines through the use of flow control devices to stabilize boundary layer characteristics

Abstract: Method and apparatus for providing a turbofan blade (40) adapted to initiate and control a boundary layer transition at a side surface of the blade (40) during operation as a component in a turbofan assembly (35). The turbofan blade (40) includes a leading edge (55), a trailing edge (58), and two side surfaces including a high-pressure side surface (49) and a low-pressure side surface (52). At least one of the two side surfaces has an essentially smooth surface portion (61) located between the leading and trailing edges, and the essentially smooth surface portion is interrupted by a surface deviation (64). The surface deviation is configured to fix a positionally unstable laminar to turbulent boundary layer transition (24) at a location toward the trailing edge from the surface deviation during operation of the turbofan blade in the turbofan assembly. In this manner, fatigue inducing and/or structurally damaging unsteady aerodynamic forces experienced upon the blade and/or fan disc during operation are controlled, and the resultant fluctuating fan blade and disc peak stresses are mitigated.

Parametric fault modeling and diagnostics of a turbofan engine

Abstract: A parametric fault modeling and diagnostics approach of a turbofan engine is presented. The healthy engine model is obtained as a steady state map between input parameters representing engine operating conditions and output parameters governing engine performance characteristics. The fault modeling is conceptualized as either degraded engine performance levels or as sensor failures and both incipient and abrupt fault scenarios are considered. The fault parameters are successfully estimated using generalized least squares (GLS) techniques with the data obtained from several recorded flight data. Further, a nonlinear estimation method is introduced to improve the diagnostic performance in certain cases of large and abrupt failures.

Generic Geometry Definition of the Aircraft Engine Combustion Chamber

Abstract: The type and layout of a particular gas turbine combustion chamber are largely determined by engine specifications and, peculiarly for the aircraft application, by the effort to use the available space as effectively as possible. Therefore, large commercial turbofan engine combustors exhibit a great degree of commonality. This commonality is a result of the similarity in working environment, size constraints and also safety, performance, and weight requirements. The objective of the present work is to propose generic relations between combustor overall performance and geometry, in order to develop accurate models for combustion quality and emission studies. Therefore, an algorithm has been developed to produce a generic combustion chamber layout. The algorithm is based on a set of empirical relations, semi-analytical methods, statistical figures and design philosophy. Results have been validated in a case study, showing accurate correspondence with modern turbofan engine combustors. An alternate application of the models may be preliminary sizing or design of aero-engine combustion chambers.Copyright © 2004 by ASME

Method for making a hollow fan blade with machined internal cavities

Abstract: Hollow fan blades for turbo fan gas turbine engines are formed of two separate detail halves Each detail half has a plurality of cavities and ribs machined out to reduce weight The floor and opposite interior walls of each cavity are machined simultaneously The configuration minimizes the number of cutter plunge cuts for the internal cavities, and maximizes cutter size, in order to minimize the time required to machine them These detail halves are subsequently bonded and given an airfoil shape in the forming operation