JP2017015095A - Gas turbine engine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To increase efficiency of a gas turbine engine.SOLUTION: A gas turbine engine comprises: a core nacelle defined about an engine centerline axis; a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow; a fan variable area nozzle axially movable relatively to fan nacelle to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation; and a gear system driven by a core engine within the core nacelle to drive a fan within the fan nacelle, the gear system defining a gear reduction ratio greater than or equal to about 2.3.SELECTED DRAWING: Figure 1A

Description

  This application claims priority to US patent application Ser. No. 13 / 346,100 filed Jan. 9, 2012, which is a continuation-in-part of U.S. patent application Ser. No. 11 / 8,213 filed Jul. 27, 2007. Is.

  The present invention relates to a gas turbine engine, and more particularly, to a turbofan engine including a variable form fan outlet guide vane (FEGV) device that changes a fan bypass passage area.

  Conventional gas turbine engines generally include a fan portion and a core portion, the fan portion having a larger diameter than the core portion. The fan portion and the core portion are provided around the longitudinal axis and are surrounded by the engine nacelle assembly. The combustion gas is discharged from the core portion through the core exhaust nozzle, and the annular fan bypass flow radially outside the main core exhaust passage is discharged through the annular fan exhaust nozzle along the fan bypass passage. Most of the thrust is generated by the bypass flow and the rest is provided by the combustion gas.

  The fan bypass channel is a compromise suitable not only for take-off and landing conditions but also for cruise conditions. The minimum area along the fan bypass channel determines the maximum air mass flow rate. Under engine-out conditions, the flow area along the bypass flow path is not sufficient, which can result in excessive flow outflow and associated drag. The fan nacelle diameter is the dimension that minimizes drag in such engine-out conditions, so the fan nacelle diameter is larger than required for normal cruise conditions and is part of the aircraft mission. Drag is lower than optimum.

  A gas turbine engine according to an exemplary embodiment herein includes a core nacelle defined about an engine central axis and at least partially of a core nacelle to define a fan bypass flow path for fan bypass airflow. A fan nacelle mounted around and a fan variable area nozzle movable around the axis relative to the fan nacelle to adjust the fan bypass air flow pressure ratio by changing the fan nozzle outlet area during engine operation And a gear device driven by a core engine in the core nacelle to drive a fan in the fan nacelle, wherein the gear device defines a gear reduction ratio of about 2.3 or more.

  In other non-limiting embodiments with respect to any of the gas turbine engine embodiments described above, the engine can further include a plurality of fan outlet guide vanes in communication with the fan bypass flow path, the plurality of fans The outlet guide vane is rotatable about the rotation axis so as to change the fan bypass flow path. In addition or alternatively, the plurality of fan outlet guide vanes can rotate simultaneously. Additionally or alternatively, a plurality of fan outlet guide vanes can be mounted within the intermediate engine case structure.

  In another non-limiting embodiment relating to any of the gas turbine engine embodiments described above, each of the plurality of fan outlet guide vanes has a pivot portion rotatable relative to the stationary portion about the rotational axis. May be included. In addition or alternatively, the pivot portion may include a leading edge flap.

  In other non-limiting embodiments relating to any of the gas turbine engine embodiments described above, the controller may change the fan nozzle outlet area to adjust the pressure ratio of the fan bypass air flow to change the fan variable area nozzle. Is operable to control In addition or alternatively, the controller is operable to reduce the fan nozzle exit area in cruise flight conditions. In addition or alternatively, the controller is operable to control the fan nozzle exit area to reduce fan instability.

  In other non-limiting embodiments relating to any of the gas turbine engine embodiments described above, the fan variable area nozzle may define the trailing edge of the fan nacelle.

  In other non-limiting embodiments relating to any of the gas turbine engine embodiments described above, the gearing can define a gear reduction ratio of about 2.5 or greater. In addition or alternatively, the gear device may define a gear reduction ratio of 2.5 or greater.

  In other non-limiting embodiments relating to any of the gas turbine engine embodiments described above, the core engine may include a low pressure turbine that defines a pressure ratio greater than about five.

  In other non-limiting embodiments relating to any of the gas turbine engine embodiments described above, the core engine may include a low pressure turbine that defines a pressure ratio greater than five.

  In other non-limiting embodiments relating to any of the gas turbine engine embodiments described above, the bypass flow may define a bypass ratio greater than about 6. In addition or alternatively, the bypass flow may define a bypass ratio greater than about 10. In addition or alternatively, the bypass flow may define a bypass ratio greater than 10.

  In other non-limiting embodiments with respect to any of the gas turbine engine embodiments described above, the engine can further include a plurality of fan outlet guide vanes in communication with the fan bypass flow path, the plurality of fan outlet guides. The vane is rotatable around the rotation axis so as to change the fan bypass flow path.

  Various features and advantages of the present invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows.

1 is an overall schematic diagram illustrating a partial cross-section of an exemplary gas turbine engine embodiment used with the present invention. It is a side view which shows the partial cross section of the FEGV apparatus which provides a fan variable area nozzle. 1 is a cross-sectional view of a single FEGV airfoil. FIG. 2B is a cross-sectional view of the FEGV of FIG. 2A in a first position. It is sectional drawing of FEGV of FIG. 2A in a rotation position. FIG. 6 is a cross-sectional view of another embodiment of a single FEGV airfoil. FIG. 3B is a cross-sectional view of the FEGV of FIG. 3A in a first position. 3B is a cross-sectional view of the FEGV of FIG. 3A in a rotational position. FIG. 6 is a cross-sectional view of another embodiment of a single FEGV airfoil with slats. FIG. 4B is a cross-sectional view of the FEGV of FIG. 4A in a first position. 4B is a cross-sectional view of the FEGV of FIG. 4A in a rotational position.

  FIG. 1 shows a partial cross-section of an entire gas turbofan engine 10 provided in an engine nacelle assembly N and suspended from an engine pylon P, as is typical for aircraft designed for subsonic flight. FIG.

  The turbofan engine 10 includes a core portion in a core nacelle 12, and the core nacelle 12 accommodates a low speed spool 14 and a high speed spool 24. The low speed spool 14 includes a low pressure compressor 16 and a low pressure turbine 18. The low speed spool 14 drives the fan unit 20 directly or through the gear train 22. The high speed spool 24 includes a high pressure compressor 26 and a high pressure turbine 28. A combustor 30 is provided between the high pressure compressor 26 and the high pressure turbine 28. The low speed and high speed spools 14 and 24 rotate about the engine rotation axis A.

  The engine 10 is an aircraft engine having a structure with a high bypass gear. In one non-limiting embodiment, the bypass ratio of the engine 10 is greater than about 6, in the exemplary embodiment greater than about 10, and the gear train 22 is a planetary gear unit having a gear reduction ratio greater than about 2.3. Such as an epicyclic gear train or other gear arrangement, wherein the low pressure turbine 18 has a pressure ratio greater than about 5. The engine 10 of the disclosed embodiment is a turbofan aircraft engine with a high bypass gear, where the bypass ratio of the engine 10 is greater than 10, the diameter of the turbofan is significantly greater than the diameter of the low pressure compressor 16, and the low pressure turbine 18 has a pressure ratio greater than 5. The pressure ratio of the low pressure turbine 18 is the pressure measured before the inlet of the low pressure turbine 18 in relation to the pressure at the outlet of the low pressure turbine 18 before the exhaust nozzle. The gear train 22 can be an epicyclic gear train, such as a planetary gear train having a gear reduction ratio greater than about 2.5, or other gear train. However, the parameters described above are exemplary for a geared turbofan engine, and the present invention is equally applicable to other gas turbine engines including direct drive turbofans.

  The air flow enters a fan nacelle 34 that at least partially surrounds the core nacelle 12. The fan unit 20 causes the air flow to communicate with the core nacelle 12, where the air flow is compressed by the low pressure compressor 16 and the high pressure compressor 26. The core air stream compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with fuel in the combustor 30 and expanded across the high pressure turbine 28 and the low pressure turbine 18. The turbines 28 and 18 are coupled to corresponding spools so as to rotate together with the spools 24 and 14 in response to expansion, and rotationally drive the fan unit via the compressors 26 and 16 and the gear train 22. The core engine exhaust E is discharged from the core nacelle 12 through a core nozzle 43 defined between the core nacelle 12 and the tail cone 32.

  A bypass channel 40 is defined between the core nacelle 12 and the fan nacelle 34. The engine 10 forms a high bypass flow structure in which approximately 80% of the air flow entering the fan nacelle 34 is the bypass flow B. The bypass flow B flows through the generally annular bypass flow path 40 and defines a variable fan nozzle outlet area 44 between the fan nacelle 34 and the core nacelle 12 in the rear segment 34S of the fan nacelle 34 downstream of the fan section 20. It is discharged from the engine 10 through a fan variable area nozzle (FVAN) 42.

  Referring to FIG. 1B, the core nacelle 12 is supported by the core engine case structure 46 as a whole. A fan case structure 48 is defined around the core engine case structure 46 to support the fan nacelle 34. The core engine case structure 46 is fixed to the fan case 48 by a plurality of fan outlet guide vanes (FEGV) 50 that are spaced apart in the circumferential direction and extend in the radial direction. The fan case structure 48, the core engine case structure 46, and a plurality of circumferentially spaced and radially extending fan outlet guide vanes 50 extending therebetween typically form a complete unit. Often referred to as an intermediate case. The fan outlet guide vane 50 can take various forms. The intermediate case structure in the disclosed embodiment includes a variable form fan outlet guide vane (FEGV) device 36.

  Thrust is a function of density, speed and area. One or more of these parameters can be manipulated to change the amount and direction of thrust obtained by the bypass flow B. Due to the high bypass ratio, a considerable amount of thrust is obtained by the bypass flow B. The fan section 20 of the engine 10 is nominally designed for specific flight conditions, typically cruises at about 0.8 Mach and about 35,000 feet. The engine's maximum fuel economy at 0.8 Mach and 35,000 ft flight conditions, also called “fuel consumption per bucket cruise thrust ('TSFC')”, is the engine at the lowest point of lbm of the fuel to be burned Is an industry standard parameter divided by lbf of the thrust generated. The “low fan pressure ratio” is the pressure ratio over only the fan blades that do not include the fan outlet guide vane (FEGV) device 36. The low fan pressure ratio in one non-limiting embodiment disclosed herein is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed (ft / sec) as an industry standard temperature correction value [(Tamient deg R) / (518.7) ^ 0.5]. Divided. In one non-limiting example disclosed herein, the “low fan tip correction speed” is less than about 1150 ft / sec.

  Because the fan section 20 is efficiently designed with a specific fixed stagger angle for efficient cruise conditions, the FEGV device 36 and / or FVAN 42 is an efficient engine in other flight conditions such as takeoff and landing. The fan bypass air flow is manipulated to maintain the fan blade angle of attack close to the design range of operation. FEGV device 36 and / or FVAN 42 can be adjusted to selectively adjust the pressure ratio of bypass flow B in response to controller C. For example, it is adjusted to increase the mass flow rate when the windmill or engine is out, or to cancel the thrust when landing. In addition, the FEGV device 36 helps the FVAN 42 and in some cases replaces the FVAN. For example, the variable flow area is utilized to manage and optimize the fan operating line, which provides operating margin and lower fan pressure ratio and lower fan tip by operating near the peak efficiency. Speed design is possible. The variable area also improves fan blade aerodynamics and reduces noise by changing the angle of attack of the blade. Thus, the FEGV device 36 provides optimal engine operation over various flight conditions with respect to other operating parameters such as performance and noise.

  Referring to FIG. 2A, each fan outlet guide vane 50 includes an airfoil portion 52 defined by an outer airfoil wall surface 54 that extends between a leading edge 56 and a trailing edge 58. The outer airfoil wall 54 typically has a generally concave portion forming a pressure surface and a generally convex portion forming a suction surface. Each airfoil portion 52 defined by the outer airfoil wall surface 54 may be generally identical and may be individually adjusted to optimize flow characteristics.

  Each fan outlet guide vane 50 is mounted about a vane longitudinal axis of rotation 60. The vane rotating shaft 60 is typically provided at a right angle or a predetermined angle with respect to the engine axis A. Various support struts 61 and other similar members may be provided through the airfoil portion 52 to provide a fixed support structure between the core engine case structure 46 and the fan case structure 48. The rotation axis 60 can be provided around the geometric center of gravity (CG) of the airfoil cross section. Actuators 62 (shown schematically in FIG. 1A), shown as unison rings for illustrative purposes only, rotate each fan outlet guide vane 50 to selectively change the fan nozzle throat area (see FIG. 2B). To work. The unison ring can be provided in the intermediate case structure, for example, inside one or both of the core engine case structure 46 and the fan case 48 (see FIG. 1A).

  In operation, the FEGV device 36 communicates with the controller C and rotates the fan outlet guide vane 50 to effectively change the fan nozzle outlet area 44. Other controllers including engine controllers and aircraft control systems can also be used with the present invention. The rotation of the fan outlet guide vane 50 between the nominal position and the rotational position selectively changes the fan bypass flow path 40. That is, both the throat area (see FIG. 2B) and the projected area (see FIG. 2C) are changed by adjusting the fan outlet guide vane 50. By adjusting the fan outlet guide vane 50 (see FIG. 2C), the bypass flow B increases in certain flight conditions such as engine-out conditions. Since less bypass flow flows out of the fan nacelle 34, the maximum fan nacelle diameter required to prevent flow separation can be reduced. This reduces fan nacell drag and nacelle assembly weight under normal cruise conditions. Conversely, by closing the FEGV device 36 to reduce the flow area for a given bypass flow, the engine thrust is substantially canceled and the need for a reverse thrust device is minimized or eliminated, Furthermore, the required weight and packaging are reduced. Other configurations and essentially infinite intermediate positions can be used with the present invention as well.

  Adjustment of the FEGV device 36, in which all fan outlet guide vanes 50 move simultaneously, maximizes engine thrust and fuel economy in each flight condition. By selectively adjusting only a particular fan outlet guide vane 50 to provide an asymmetric fan bypass flow path 40, the direction of engine bypass flow can be selectively changed, for example, by way of example only, trim balance, thrust control Improved maneuvering, ground handling and short range performance can be provided.

  Referring to FIG. 3A, another embodiment of the FEGV device 36 ′ includes a plurality of fan outlet guide vanes 50 ′ that include a fixed airfoil portion 66F and a fixed airfoil portion 66F. Pivot airfoil portions 66P pivoting relative to each other. Pivot airfoil portion 66P may include a leading edge flap operable by actuator 62 'to change both the throat area (see FIG. 3B) and the projected area (see FIG. 3C) as described above.

  Referring to FIG. 4A, yet another embodiment of the FEGV device 36 "includes a plurality of slatted fan outlet guide vanes 50" that include a fixed airfoil portion 68F, a slat. Pivot / sliding airfoil portion 68P that pivots and slides relative to the stationary airfoil portion 68F to form a 70, respectively, so that, as outlined above, the throat area ( Both FIG. 4B) and the projected area (FIG. 4C) are changed. This slatted vane method not only increases the flow area, but also from the high pressure convex surface of the fan outlet guide vane 50 "when there is a negative incidence on the fan outlet guide vane 50". The flow of air to the low pressure concave surface of the fan outlet guide vane 50 "is possible, and the flow separation is slow.

  The above description is illustrative and not restrictive. Various modifications and variations of the present invention are possible in light of the above teachings. While preferred embodiments of the invention have been disclosed, those skilled in the art will recognize that certain modifications are within the scope of the invention. Thus, within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. Thus, in order to determine the true scope and content of the present invention, the following claims should be considered.

Claims (14)

A core nacelle defined around the center axis of the engine,
A core engine located in the coana cell;
A fan nacelle mounted at least partially around the core nacelle and defining a fan bypass flow path for the fan bypass air flow;
A fan unit driven by a core engine and rotatable in the fan nacelle;
A fan variable area nozzle movable about the axis relative to the fan nacelle to change the fan nozzle outlet area during engine operation to adjust the pressure ratio of the fan bypass air flow;
A plurality of fan outlet guide vanes in communication with the fan bypass flow path;
In order to drive a fan in the fan nacelle, a gear device driven by a core engine in the core nacelle is provided, and the gear device defines a gear reduction ratio of 2.3 or more,
A gas turbine engine characterized in that the fan pressure ratio is smaller than 1.45 and the fan portion rotates at a fan tip speed slower than 1150 ft / sec.   The gas turbine engine according to claim 1, wherein the plurality of fan outlet guide vanes are rotatable about a rotation axis so as to change a fan bypass flow path.   The gas turbine engine according to claim 2, wherein the plurality of fan outlet guide vanes are rotatable at the same time.   The gas turbine engine according to claim 2, wherein the plurality of fan outlet guide vanes are attached to an intermediate engine case structure including a core engine case structure and a fan case structure.   The gas turbine engine according to claim 2, wherein the plurality of fan outlet guide vanes include a pivot portion that is rotatable with respect to the fixed portion about the rotation shaft.   The gas turbine engine according to claim 5, wherein the pivot portion includes a leading edge flap.   The gas turbine of claim 1, further comprising a controller operable to control the fan variable area nozzle to change the fan nozzle outlet area to adjust the pressure ratio of the fan bypass air flow. engine.   The gas turbine engine of claim 7, wherein the controller is operable to reduce a fan nozzle exit area under cruise flight conditions.   The gas turbine engine of claim 7, wherein the controller is operable to control a fan nozzle exit area to reduce fan instability.   The gas turbine engine according to claim 1, wherein the fan variable area nozzle defines a trailing edge of the fan nacelle.   The gas turbine engine according to claim 1, wherein the gear device defines a gear reduction ratio of 2.5 or more.   The gas turbine engine of claim 1, wherein the core engine includes a low pressure turbine that defines a pressure ratio greater than five.   The gas turbine engine of claim 1, wherein the bypass flow defines a bypass ratio greater than six.   The gas turbine engine of claim 1, wherein the bypass flow defines a bypass ratio that is greater than ten.

Images