Showing papers on "Turbofan published in 1993"

Aircraft bypass turbofan engine thrust reverser

Abstract: An aircraft bypass turbofan engine subassembly for ground deceleration (eg, thrust reversal) A fan nacelle has a through passageway connecting its inner exterior surface and outer exterior surface A passageway end is located on the inner exterior surface aft of the aft-most row of fan rotor blades A row of pivotable fan outlet guide vanes is located radially between the fan and core nacelles and longitudinally aft of the flow splitter and the passageway end Apparatus is provided to pivot the vanes to block passage of bypass air and to open the passageway for ground deceleration

Active control of fan noise from a turbofan engine

Abstract: A three-channel active control system is applied to an operational turbofan engine to reduce tonal noise produced by both the fan and the high-pressure compressor. The control approach is the feedforward filtered-x least-mean-square algorithm implemented on a digital signal processing board. Reference transducers mounted on the engine case provide blade passing and harmonics frequency information to the controller. Error information is provided by large area microphones placed in the acoustic far field. To minimize the error signal, the controller actuates loudspeakers mounted on the inlet to produce destructive interference. The sound pressure level of the fundamental tone of the fan was reduced using the three-channel controller by up to 16 dB over a +/- 30-deg angle about the engine axis. A single-channel controller could produce reduction over a +/- 15-deg angle. The experimental results show the control to be robust. Outside of the areas contolled, the levels of the tone actually increased due to the generation of radial modes by the control sources. Simultaneous control of two tones is achieved with parallel controllers. The fundamental and the first harmonic tones of the fan were controlled simultaneously with reductions of 12 and 5 dBA, respectively, measured on the engine axis. Simultaneous control was also demonstrated for the fan fundamental and the high-pressure compressor fundamental tones.

Fan blade for blade-out protection

Abstract: In a turbofan engine a fan designed for reduced damage during a blade-out condition wherein the fan blade airfoil is disposed asymmetrically on the platform and includes a forwardly displaced leading edge. The platform is configured to make early contact with a following airfoil upon blade-out.

Turbofan engine bypass and exhaust system

Abstract: A turbofan engine bypass and exhaust system has a double-hoop configuration in which a fan exhaust passageway comprises shells of a constant circular contour to resolve air pressure induced stresses and increase in bypass passageway area asymmetrically in the horizontal direction without substantially increasing the overall engine height.

Pylon and engine installation for ultra-high by-pass turbo-fan engines

Abstract: A pylon (14) for supporting a turbo-fan jet engine, in particular, an ultra-high by-pass turbo-fan jet engine (12) is disclosed. The pylon (14) attaches to the forward and aft parts of the engine core (22) and supports the engine (12) from an airplane wing (10) while resisting static and dynamic forces applied by the engine and the nacelle of the engine to the pylon. The pylon (14) includes a pylon bifurcation portion (16), having an upper end (23) and a lower end. The upper end (23) is positioned exterior to the fan duct (8) of the engine and is attached to the wing (10). The lower end is positioned inside of the fan duct (8). The lower end further includes an aft section that is attached by lugs (38) to the aft part of the engine core (22), and a forward section that is attached to the forward part of the engine core. Left and right cowl beams (18, 20) extend forward from opposite sides of the forward section of the lower portion of the pylon bifurcation portion (16) and attach to different locations on the forward part of the engine core (22). The cowl beams (18, 20) resist the majority portion of torque forces applied by the engine (12) to the pylon (14).

Approach of modeling continuous turbine engine operation from startup to shutdown

Abstract: A generalized turbine engine start simulation (mathematical model) has been developed and demonstrated. The model, designated as ATEST-V3, is capable of simulating engine operation continuously from near static (zero speed) conditions to maximum engine power including windmill starting, spooldown starting, and starterassisted starting. The enhanced capability to simulate the engine starting process provides the means to characterize and understand engine system operational behavior during critical startup and shutdown operations. ATEST-V3 is based on an aerothermodynamic matching of the major components. The component-matching technique is widely used for steady-state and transient turbine engine simulations that typically exclude subidle and starting operations. The same approach is shown to be applicable to engine starting operations by modeling component behavior continuously from zero to maximum power. The combination of an existing transient engine simulation and a numerically stable component-matching algorithm provided a foundation for extending the simulation capability to subidle engine operation and engine starting. ATEST-V3 was applied to a modern flighttype turbofan engine which demonstrated the capability to simulate windmill, spooldown, and starter-assisted starts at various flight conditions. Finally, a comparison is made between model results and engine test data.

Gas turbine engine control system having integral flight Mach number synthesis method

Abstract: A gas turbine engine digital electronic control is disclosed which is capable of maintaining safe, efficient engine thrust levels in the event of loss of communication with an aircraft flight computer. Specifically, a limited number of engine sensor inputs are used in combination with an engine performance schedule to derive total inlet pressure. A compressible flow calculation yields flight Mach number which is used in a power management module to control engine power settings, such as corrected fan speed in a turbofan engine. Various correction factors and scalars are provided to configure the control for a variety of installation configurations, as well as improve system accuracy.

Ramjet bypass duct and preburner configuration

Abstract: A combined turbofan and ramjet aircraft engine includes a forward bypass duct which allows the engine to operate more efficiently during the turbofan mode of operation. By mounting a ramjet preburner in the forward duct and isolating this duct from the turbofan bypass air, a transition from turbofan operation to ramjet operation can take place at lower flight Mach numbers without incurring pressure losses or blockage in the turbofan bypass air.

Acoustic mode measurements in the inlet of a model turbofan using a continuously rotating rake

Abstract: Comprehensive measurements of the spinning acoustic mode structure in the inlet of the Advanced Ducted Propeller (ADP) have been completed. These measurements were taken using a unique and previously untried method which was first proposed by T.G. Sofrin. A continuously rotating microphone system was employed. The ADP model was designed and built by Pratt & Whitney and tested in the NASA Lewis 9- by 15-foot Anechoic Wind Tunnel. Three inlet configurations were tested with cut-on and cutoff stator vane sets. The cutoff stator was designed to suppress all modes at the blade passing frequency. Rotating rake measurements indicate that several extraneous circumferential modes were active. The mode orders suggest that their source was an interaction between the rotor and small interruptions in the casing tip treatment. The cut-on stator produced the expected circumferential modes plus higher levels of the unexpected modes seen with the cutoff stator.

GETRAN: A Generic, Modularly Structured Computer Code for Simulation of Dynamic Behavior of Aero- and Power Generation Gas Turbine Engines

Abstract: The design concept, the theoretical background essential for the development of the modularly structured simulation code GETRAN, and several critical simulation cases are presented in this paper. The code being developed under contract with NASA Lewis Research Center is capable of simulating the nonlinear dynamic behavior of single- and multi-spool core engines, turbofan engines, and power generation gas turbine engines under adverse dynamic operating conditions. The modules implemented into GETRAN correspond to components of existing and new generation aero- and stationary gas turbine engines with arbitrary configuration and arrangement. For precise simulation of turbine and compressor components, row-by-row diabatic and adiabatic calculation procedures are implemented that account for the specific turbine and compressor cascade, blade geometry, and characteristics. The nonlinear, dynamic behavior of the subject engine is calculated solving a number of systems of partial differential equations, which describe the unsteady behavior of each component individually. To unambiguously identify each differential equation system, special attention is paid to the addressing of each component. The code is capable of executing the simulation procedure at four levels which increase with the degree of complexity of the system and dynamic event. As representative simulations, four different transient cases with single- and multi-spool thrust and power generation engines were simulated. These transient cases vary from throttling the exit nozzle area, operation with fuel schedule, rotor speed control, to rotating stall and surge.© 1993 ASME

Theoretical implications of Active Noise Control for turbofan engines

Abstract: The objective of this stud)' is to develop the analytical tools that will allow the assessment of the feasibility of using Active Noise Control (ANC) in aircraft engine ducts to suppress radiated tones. It is assumed that the active noise control system will operate using the "noise cancellation'' principle, such that control sound source transducers will inject sound into the duct that directly interferes with the sound generated by the turbomachinery sources. The approach will be to develop the theory needed bused on duct modal analysis. In addition, it is hoped that an analytical approach f o r assessing the feasibility of aircraft engine ANC systems may also provide some guidance for the design optimization of such systems. Analyses are developed for the circumferential and radial mode coupling of wall-mounted sources to a cylindrical duct. A sample case is included that examines energy requirements for ANC source generation.

A Comparison of the Measured and Predicted Flow Field in a Modern Fan-Bypass Configuration

Abstract: A three-dimensional viscous Navier-Stokes flow solver was used to predict core and bypass rotor performance and radial flow characteristics of a 4.6:1 bypass ratio, single-stage fan. The three-dimensional flow solver can handle several blade rows simultaneously and has the capability to include a downstream splitter. Results of the analysis are compared with experimental data obtained during rig testing of a modern high bypass single-stage turbofan in which rotor performance for both bypass and core steams was measured

Commercial turbofan engine exhaust nozzle flow analyses

Abstract: The recently developed three-dimensional code is able to perform a computational investigation of complex aircraft aerodynamic components. This code was developed for solving the simplified Reynolds-averaged NavierStokes equations in a three-dimensional multiblock/mul tizone structured mesh domain. The present analysis was applied to commercial turbofan exhaust flow systems. Solution sensitivity to grid density is presented. Laminar flow solutions were developed for all grids, and two-equation k-£ solutions were developed for selected grids. Static pressure distributions, mass flow, and thrust quantities were calculated for on-design engine operating conditions. Good agreement between predicted surface static pressures and experimental data was observed at different locations. Mass flow was predicted within 0.2% of experimental data. Thrust forces were typically within 0.6% of experimental data.

Turbofan engine exhaust mixing area modification for improved engine efficiency and noise reduction

Abstract: A modified exhaust tailpipe especially for use with an aircraft turbofan engine utilizing a noise suppressor. The modified tailpipe is connected to the exhaust duct of the engine. The forward end of the tailpipe is essentially the same diameter as that of the exhaust duct of the engine. The contour of the tailpipe is such that it increases in diameter to form a bulge to increase the flow area for fan gases in the area of a mixer which is supported within the forward end of the tailpipe exhaust system. This improves the operating efficiency of the jet engine.

CFD-Based Three-Dimensional Turbofan Exhaust Nozzle Analysis System

Abstract: A three-dimensional turbofan exhaust nozzle analysis system based on computational fluid dynamics (CFD) has been developed. This system has been established to aid exhaust designers in the efficient assessment and screening of their design concepts, with the prospects of a reduction in both design cycle time and wind-tunnel test costs. A reliable CFD flow solver, user-friendly grid generator, and postprocessing software are included in the system. The system is easy to use for exhaust designers who are not particularly familiar with the inner workings of CFD. Validation and applicability studies have been performed using different exhaust nozzle configurations at on-design and off-design engine operating conditions. This work demonstrates that a CFD code integrated with automatic grid generation and postprocessing can be a useful analytical tool in the practical exhaust nozzle design process. Nomenclature M = Mach number P = static pressure PA = atmospheric static pressure R = radius X — axial station XREF = reference length Subscript 00 = freestream conditions

Analysis of gas turbine engines using water and oxygen injection to achieve high Mach numbers and high thrust

Abstract: An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.

Extended range operations of two and three turbofan engined airplanes

Abstract: The objective of the present work is a comparative analysis of the behavior of two and three turbofan engined airplanes after engine failure. A simple but fairly realistic treatment of the range equation allows study of extended range operations of airplanes after any prescribed decrease in thrust. The approach takes into account the increase in parasite drag, and considers variations of thrust and specific fuel consumption with altitude and Mach number. All peculiarities of the powerplant are translated into a few nondimensional parameters. The model provides the long-range cruise conditions after engine failure, namely altitude and Mach number, and the additional fuel needed to reach the final destination. Results for a typical 5000-km route show the relative disadvantage of twins. Nomenclature A = aspect ratio of wing AF = extra fuel needed to reach destination due to engine failure aQ = speed of sound at sea level of = normalized extra fuel, AF/W* C = specific fuel consumption CDO = parasite drag coefficient CL = lift coefficient C0 = increment factor in parasite drag due to engine

Admission mixing duct assembly

Abstract: A variable cycle jet engine is provided with a mixing duct assembly which mixes core engine exhaust gas with bypass air when the engine is operating in a turbofan mode and which blocks flow from the core engine and isolates the core engine from the bypass flow when the engine is operating as a ramjet.

Heat Pipe Cooling of Turboshaft Engines

Abstract: Reduction in turbine engine cooling flows is required to meet the IHPTET Phase II engine performance levels. Heat pipes, which are devices with very high thermal conductance, can help reduce the required cooling air. A survey was conducted to identify potential applications for heat pipes in turboshaft engines. The applications for heat pipe cooling of turbine engine components included the power turbine first stage vanes, shroud, and case, the HP turbine vanes and shroud, and the T5 temperature probe. Other potential applications for heat pipe cooling include regenerative cycle and intercooling, bearing cooling, IR signature reduction, and active clearance control. Calculated performance benefits included an increase in specific shaft horsepower, and a decrease in specific fuel consumption, as determined with an IHPTET Phase II turboshaft engine performance model. For example, using heat pipes to cool the power turbine vanes, shroud, and case would increase the specific shaft horsepower by 6 percent, while decreasing the specific fuel consumption by 2.2 percent. While this study examined turboshaft engines, most of the applications are also applicable to turbofan engines.Copyright © 1993 by ASME

Applications of active adaptive noise control to jet engines

Abstract: During phase 2 research on the application of active noise control to jet engines, the development of multiple-input/multiple-output (MIMO) active adaptive noise control algorithms and acoustic/controls models for turbofan engines were considered. Specific goals for this research phase included: (1) implementation of a MIMO adaptive minimum variance active noise controller; and (2) turbofan engine model development. A minimum variance control law for adaptive active noise control has been developed, simulated, and implemented for single-input/single-output (SISO) systems. Since acoustic systems tend to be distributed, multiple sensors, and actuators are more appropriate. As such, the SISO minimum variance controller was extended to the MIMO case. Simulation and experimental results are presented. A state-space model of a simplified gas turbine engine is developed using the bond graph technique. The model retains important system behavior, yet is of low enough order to be useful for controller design. Expansion of the model to include multiple stages and spools is also discussed.

Measurement effects on the calculation of in-flight thrust for an F404 turbofan engine

Abstract: A study has been performed that investigates parameter measurement effects on calculated in-flight thrust for the General Electric F404-GE-400 afterburning turbofan engine which powered the X-29A forward-swept wing research aircraft. Net-thrust uncertainty and influence coefficients were calculated and are presented. Six flight conditions were analyzed at five engine power settings each. Results were obtained using the mass flow-temperature and area-pressure thrust calculation methods, both based on the commonly used gas generator technique. Thrust uncertainty was determined using a common procedure based on the use of measurement uncertainty and influence coefficients. The effects of data nonlinearity on the uncertainty calculation procedure were studied and results are presented. The advantages and disadvantages of using this particular uncertainty procedure are discussed. A brief description of the thrust-calculation technique along with the uncertainty calculation procedure is included.

Advanced control for airbreathing engines, volume 1: Pratt and Whitney

Abstract: The application of advanced control concepts to air breathing engines may yield significant improvements in aircraft/engine performance and operability. Screening studies of advanced control concepts for air breathing engines were conducted by three major domestic aircraft engine manufacturers to determine the potential impact of concepts on turbine engine performance and operability. The purpose of the studies was to identify concepts which offered high potential yet may incur high research and development risk. A target suite of proposed advanced control concepts was formulated and evaluated in a two phase study to quantify each concept's impact on desired engine characteristics. To aid in the evaluation specific aircraft/engine combinations were considered: a Military High Performance Fighter mission, a High Speed Civil Transport mission, and a Civil Tiltrotor mission. Each of the advanced control concepts considered in the study are defined and described. The concept potential impact on engine performance was determined. Relevant figures of merit on which to evaluate the concepts are determined. Finally, the concepts are ranked with respect to the target aircraft/engine missions. A final report describing the screening studies was prepared by each engine manufacturer. Volume 1 of these reports describes the studies performed by Pratt & Whitney.