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WO2006059972A1 - Compressor variable stage remote actuation for turbine engine - Google Patents

Compressor variable stage remote actuation for turbine engine Download PDF

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Publication number
WO2006059972A1
WO2006059972A1 PCT/US2004/039974 US2004039974W WO2006059972A1 WO 2006059972 A1 WO2006059972 A1 WO 2006059972A1 US 2004039974 W US2004039974 W US 2004039974W WO 2006059972 A1 WO2006059972 A1 WO 2006059972A1
Authority
WO
WIPO (PCT)
Prior art keywords
vane
compressor
axis
turbine engine
actuator
Prior art date
Application number
PCT/US2004/039974
Other languages
French (fr)
Inventor
Gabriel L. Suciu
Brian Merry
Original Assignee
United Technologies Corporation
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Technologies Corporation filed Critical United Technologies Corporation
Priority to EP04822095A priority Critical patent/EP1831530B1/en
Priority to PCT/US2004/039974 priority patent/WO2006059972A1/en
Priority to DE602004019710T priority patent/DE602004019710D1/en
Priority to US11/719,594 priority patent/US7934902B2/en
Publication of WO2006059972A1 publication Critical patent/WO2006059972A1/en

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/022—Blade-carrying members, e.g. rotors with concentric rows of axial blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
    • F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
    • F02C3/06—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
    • F02C3/073—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages the compressor and turbine stages being concentric
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
    • F02C7/057—Control or regulation
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/32—Arrangement, mounting, or driving, of auxiliaries
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/068—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type being characterised by a short axial length relative to the diameter
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D27/00—Control, e.g. regulation, of pumps, pumping installations or pumping systems specially adapted for elastic fluids
    • F04D27/02—Surge control
    • F04D27/0246—Surge control by varying geometry within the pumps, e.g. by adjusting vanes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/40—Transmission of power

Definitions

  • the present invention relates to turbine engines, and more particularly to a remote actuator for a variable stage of a compressor for a turbine engine, such as a tip turbine engine.
  • An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis.
  • a high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft.
  • the high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream.
  • the gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high spool shaft.
  • the gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low spool shaft.
  • turbofan engines operate in an axial flow relationship.
  • the axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
  • Tip turbine engines include hollow fan blades that receive core airflow therethrough such that the hollow fan blades operate as a high pressure centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490.
  • the tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.
  • Conventional turbine engines and tip turbine engines may include a low pressure compressor upstream of the high pressure compressor.
  • Conventional gas turbine engines include variable guide vanes at the inlet to the high pressure compressor. Actuators for the variable guide vanes are located on the compressor case. In the tip turbine engine, however, the compressor case has a small radial thickness because it is located radially inward of the bypass airflow path. Packaging the actuator for variable compressor vanes on the compressor case would be difficult or would require increasing the diameter of the splitter, which would thereby encroach on the bypass airflow path.
  • a turbine engine locates an actuator for the compressor variable guide vanes at a location remote from the variable guide vanes.
  • the actuator is connected to the variable guide vanes through a torque rod inside an inlet guide vane.
  • the torque rod rotatably drives an activation ring about the engine centerline.
  • the activation ring is coupled to each of the variable compressor vanes via levers, such that rotation of the activation ring about the engine centerline causes all of the variable compressor vanes to pivot.
  • the tip turbine engine can be more easily provided with variable compressor vanes for the axial compressor.
  • This configuration also allows easy access to the actuators and plumbing for servicing.
  • the present invention could also be used in conventional gas turbine engines.
  • Figure 1 is a partial sectional perspective view of a tip turbine engine
  • Figure 2 is a longitudinal sectional view of the tip turbine engine of Figure 1 along an engine centerline.
  • FIG 1 illustrates a general perspective partial sectional view of a tip turbine engine (TTE) type gas turbine engine 10.
  • the engine 10 includes an outer nacelle 12, a rotationally fixed static outer support structure 14 and a rotationally fixed static inner support structure 16.
  • a plurality of fan inlet guide vanes 18 are mounted between the static outer support structure 14 and the static inner support structure 16.
  • Each inlet guide vane preferably includes a variable trailing edge 18 A.
  • a nosecone 20 is preferably located along the engine centerline A to improve airflow into an axial compressor 22, which is mounted about the engine centerline A behind the nosecone 20.
  • a fan-turbine rotor assembly 24 is mounted for rotation about the engine centerline A aft of the axial compressor 22.
  • the fan-turbine rotor assembly 24 includes a plurality of hollow fan blades 28 to provide internal, centrifugal compression of the compressed airflow from the axial compressor 22 for distribution to an annular combustor 30 located within the rotationally fixed static outer support structure 14.
  • a turbine 32 includes a plurality of tip turbine blades 34 (two stages shown) which rotatably drive the hollow fan blades 28 relative a plurality of tip turbine stators 36 which extend radially inwardly from the rotationally fixed static outer support structure 14.
  • the annular combustor 30 is disposed axially forward of the turbine 32 and communicates with the turbine 32.
  • the rotationally fixed static inner support structure 16 includes a splitter 40, a static inner support housing 42 and a static outer support housing 44 located coaxial to said engine centerline A.
  • the axial compressor 22 includes the axial compressor rotor 46, which is mounted for rotation upon the static inner support housing 42 through an aft bearing assembly 47 and a forward bearing assembly 48.
  • a plurality of stages of compressor blades 52a-c extend radially outwardly from the axial compressor rotor 46.
  • a fixed compressor case 50 is mounted within the splitter 40.
  • a plurality of compressor vanes 54a-c extend radially inwardly from the compressor case 50 between stages of the compressor blades 52a-c.
  • the compressor blades 52a-c and compressor vanes 54a-c are arranged circumferentially about the axial compressor rotor 46 in stages (three stages of compressor blades 52a-c and compressor vanes 54a-c are shown in this example).
  • the first compressor vane 54a is variable, i.e. it is selectively pivotable about an axis P that is transverse to the engine centerline.
  • the axis P is approximately perpendicular, but tilted slightly to accommodate the narrowing core airflow path of the axial compressor 22.
  • the other compressor vanes 54b-c could optionally be variable as well, either dependently or independently of the first compressor vane 54a.
  • the rotational position of the compressor vane 54a is controlled by an actuator 55 that is mounted remotely from the compressor vane 54a and remotely from the axial compressor 22.
  • the actuator 55 may be hydraulic, electric or any other type of suitable actuator.
  • the actuator 55 is located within the nacelle 12, radially outward of the bypass airflow path.
  • the compressor vane 54a is operatively connected to the compressor vane 54a via a torque rod 56 that is routed through one of the inlet guide vanes 18.
  • the torque rod 56 is coupled to an activation ring 57 via a torque rod lever 58.
  • the activation ring 57 is rotatable about the engine centerline A.
  • the activation ring 57 is in turn coupled to a shaft 63 of the variable guide vane 54a via an activation lever 59.
  • the plurality of variable guide vanes 54a (only one shown) are disposed circumferentially about the engine centerline A, and each is connected to the activation ring 57 in the same manner.
  • the actuator 55 is coupled to the torque rod 56 by an actuator lever 60. As will be noted, the actuator 55 is spaced away from the variable guide vane 54a in a direction having a component generally parallel to the pivot axis P and by a distance that is substantially greater than a vane height of the variable guide vane 54a as measured along the pivot axis P.
  • the fan-turbine rotor assembly 24 includes a fan hub 64 that supports a plurality of the hollow fan blades 28.
  • Each fan blade 28 includes an inducer section 66, a hollow fan blade section 72 and a diffuser section 74.
  • the inducer section 66 receives airflow from the axial compressor 22 generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction.
  • the airflow is radially communicated through a core airflow passage 80 within the fan blade section 72 where the airflow is centrifugally compressed. From the core airflow passage 80, the airflow is diffused and turned once again toward an axial airflow direction toward the annular combustor 30.
  • the airflow is diffused axially forward in the engine 10, however, the airflow may alternatively be communicated in another direction.
  • the tip turbine engine 10 may optionally include a gearbox assembly 90 aft of the fan-turbine rotor assembly 24, such that the fan-turbine rotor assembly 24 rotatably drives the axial compressor 22 via the gearbox assembly 90.
  • the gearbox assembly 90 provides a speed increase at a 3.34-to- one ratio.
  • the gearbox assembly 90 is an epicyclic gearbox, such as the planetary gearbox shown, that is mounted for rotation between the static inner support housing 42 and the static outer support housing 44.
  • the gearbox assembly 90 includes a sun gear 92, which rotates the axial compressor 22, and a planet carrier 94, which rotates with the fan-turbine rotor assembly 24.
  • a plurality of planet gears 93 each engage the sun gear 92 and a rotationally fixed ring gear 95.
  • the planet gears 93 are mounted to the planet carrier 94.
  • the gearbox assembly 90 is mounted for rotation between the sun gear 92 and the static outer support housing 44 through a gearbox forward bearing 96 and a gearbox rear bearing 98.
  • the gearbox assembly 90 may alternatively, or additionally, reverse the direction of rotation and/or may provide a decrease in rotation speed.
  • the actuator 55 selectively moves the end of the actuator lever 60 to cause the torque rod 56 to pivot on its longitudinal axis within the inlet guide vane 18. This causes rotation of the activation ring 57 about the engine centerline A, which in turn causes each of the activation levers 59 to pivot and cause rotation of the corresponding compressor vane 54a about its axis P to a selected angle.
  • the selected angle of the compressor vane 54a adjusts the core airflow entering the axial compressor 22, where it is compressed by the compressor blades 52a-c.
  • the compressed air from the axial compressor 22 enters the inducer section 66 in a direction generally parallel to the engine centerline A, and is then turned by the inducer section 66 radially outwardly through the core airflow passage 80 of the hollow fan blades 28.
  • the airflow is further compressed centrifugally in the hollow fan blades 28 by rotation of the hollow fan blades 28.
  • From the core airflow passage 80 the airflow is turned and diffused axially forward in the engine 10 into the annular combustor 30.
  • the compressed core airflow from the hollow fan blades 28 is mixed with fuel in the annular combustor 30 and ignited to form a high-energy gas stream.
  • the high-energy gas stream is expanded over the plurality of tip turbine blades 34 mounted about the outer periphery of the fan-turbine rotor assembly 24 to drive the fan-turbine rotor assembly 24, which in turn rotatably drives the axial compressor 22 either directly or via the optional gearbox assembly 90.
  • the fan- turbine rotor assembly 24 discharges fan bypass air axially aft to merge with the core airflow from the turbine 32 in an exhaust case 106.
  • a plurality of exit guide vanes 108 are located between the static outer support housing 44 and the rotationally fixed static outer support structure 14 to guide the combined airflow out of the engine 10 and provide forward thrust.
  • An exhaust mixer 110 mixes the airflow from the turbine blades 34 with the bypass airflow through the fan blades 28.
  • the actuator 55 is located remotely from the variable compressor vanes 54a, the size of the splitter 40 and compressor case 50 can be minimized. Routing the linkage between the actuator 55 and the compressor vane 54a within the interiors of the nacelle 12, inlet guide vane 18 and splitter 40 does not require any additional space and provides easy access to the various components for service.
  • exemplary configurations described above are considered to represent a preferred embodiment of the invention. However, it should be noted that the invention can be practiced otherwise than as specifically illustrated and described without departing from its spirit or scope. For example, there are many configurations of linkages, rigid and/or flexible, that could be used to connect the remote actuator 55 to the compressor vane 54a. Also, although the remote actuator 55 has been shown in connection with a tip turbine engine 10, it could also be used in conventional or other turbine engines.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Geometry (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A turbine engine locates an actuator (55) for the compressor variable guide vanes (54a, b, c) at a location remote from the variable guide vanes. The actuator is connected to the variable guide vanes through a torque rod (56) inside an inlet guide vane. The torque rod rotatably drives an activation ring (57) about the engine centerline. The activation ring is coupled to each of the variable compressor vanes via linkage, such that rotation of the activation ring about the engine centerline causes all of the variable compressor vanes to pivot. With this configuration, a tip turbine engine can be more easily provided with variable compressor vanes for the axial compressor.

Description

COMPRESSOR VARIABLE STAGE REMOTE ACTUATION FOR TURBINE ENGINE
This invention was conceived in performance of U.S. Air Force contract F33657-03-C-2044. The government may have rights in this invention.
BACKGROUND OF THE INVENTION
The present invention relates to turbine engines, and more particularly to a remote actuator for a variable stage of a compressor for a turbine engine, such as a tip turbine engine.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan, a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low spool shaft.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines include hollow fan blades that receive core airflow therethrough such that the hollow fan blades operate as a high pressure centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.
Conventional turbine engines and tip turbine engines may include a low pressure compressor upstream of the high pressure compressor. Conventional gas turbine engines include variable guide vanes at the inlet to the high pressure compressor. Actuators for the variable guide vanes are located on the compressor case. In the tip turbine engine, however, the compressor case has a small radial thickness because it is located radially inward of the bypass airflow path. Packaging the actuator for variable compressor vanes on the compressor case would be difficult or would require increasing the diameter of the splitter, which would thereby encroach on the bypass airflow path.
SUMMARY OF THE INVENTION
A turbine engine according to the present invention locates an actuator for the compressor variable guide vanes at a location remote from the variable guide vanes. The actuator is connected to the variable guide vanes through a torque rod inside an inlet guide vane. The torque rod rotatably drives an activation ring about the engine centerline. The activation ring is coupled to each of the variable compressor vanes via levers, such that rotation of the activation ring about the engine centerline causes all of the variable compressor vanes to pivot.
As a result, the tip turbine engine can be more easily provided with variable compressor vanes for the axial compressor. This configuration also allows easy access to the actuators and plumbing for servicing. The present invention could also be used in conventional gas turbine engines.
BRIEF DESCRIPTION OF THE DRAWINGS Other advantages of the present invention can be understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein: Figure 1 is a partial sectional perspective view of a tip turbine engine; and Figure 2 is a longitudinal sectional view of the tip turbine engine of Figure 1 along an engine centerline.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
Figure 1 illustrates a general perspective partial sectional view of a tip turbine engine (TTE) type gas turbine engine 10. The engine 10 includes an outer nacelle 12, a rotationally fixed static outer support structure 14 and a rotationally fixed static inner support structure 16. A plurality of fan inlet guide vanes 18 are mounted between the static outer support structure 14 and the static inner support structure 16. Each inlet guide vane preferably includes a variable trailing edge 18 A. A nosecone 20 is preferably located along the engine centerline A to improve airflow into an axial compressor 22, which is mounted about the engine centerline A behind the nosecone 20. A fan-turbine rotor assembly 24 is mounted for rotation about the engine centerline A aft of the axial compressor 22. The fan-turbine rotor assembly 24 includes a plurality of hollow fan blades 28 to provide internal, centrifugal compression of the compressed airflow from the axial compressor 22 for distribution to an annular combustor 30 located within the rotationally fixed static outer support structure 14.
A turbine 32 includes a plurality of tip turbine blades 34 (two stages shown) which rotatably drive the hollow fan blades 28 relative a plurality of tip turbine stators 36 which extend radially inwardly from the rotationally fixed static outer support structure 14. The annular combustor 30 is disposed axially forward of the turbine 32 and communicates with the turbine 32.
Referring to Figure 2, the rotationally fixed static inner support structure 16 includes a splitter 40, a static inner support housing 42 and a static outer support housing 44 located coaxial to said engine centerline A.
The axial compressor 22 includes the axial compressor rotor 46, which is mounted for rotation upon the static inner support housing 42 through an aft bearing assembly 47 and a forward bearing assembly 48. A plurality of stages of compressor blades 52a-c extend radially outwardly from the axial compressor rotor 46. A fixed compressor case 50 is mounted within the splitter 40. A plurality of compressor vanes 54a-c extend radially inwardly from the compressor case 50 between stages of the compressor blades 52a-c. The compressor blades 52a-c and compressor vanes 54a-c are arranged circumferentially about the axial compressor rotor 46 in stages (three stages of compressor blades 52a-c and compressor vanes 54a-c are shown in this example). As will be further explained, the first compressor vane 54a is variable, i.e. it is selectively pivotable about an axis P that is transverse to the engine centerline. In the example shown, the axis P is approximately perpendicular, but tilted slightly to accommodate the narrowing core airflow path of the axial compressor 22. The other compressor vanes 54b-c could optionally be variable as well, either dependently or independently of the first compressor vane 54a.
The rotational position of the compressor vane 54a is controlled by an actuator 55 that is mounted remotely from the compressor vane 54a and remotely from the axial compressor 22. The actuator 55 may be hydraulic, electric or any other type of suitable actuator. In the embodiment shown, the actuator 55 is located within the nacelle 12, radially outward of the bypass airflow path. The compressor vane 54a is operatively connected to the compressor vane 54a via a torque rod 56 that is routed through one of the inlet guide vanes 18. Within the splitter 40, the torque rod 56 is coupled to an activation ring 57 via a torque rod lever 58. The activation ring 57 is rotatable about the engine centerline A. The activation ring 57 is in turn coupled to a shaft 63 of the variable guide vane 54a via an activation lever 59. The plurality of variable guide vanes 54a (only one shown) are disposed circumferentially about the engine centerline A, and each is connected to the activation ring 57 in the same manner. The actuator 55 is coupled to the torque rod 56 by an actuator lever 60. As will be noted, the actuator 55 is spaced away from the variable guide vane 54a in a direction having a component generally parallel to the pivot axis P and by a distance that is substantially greater than a vane height of the variable guide vane 54a as measured along the pivot axis P. The fan-turbine rotor assembly 24 includes a fan hub 64 that supports a plurality of the hollow fan blades 28. Each fan blade 28 includes an inducer section 66, a hollow fan blade section 72 and a diffuser section 74. The inducer section 66 receives airflow from the axial compressor 22 generally parallel to the engine centerline A and turns the airflow from an axial airflow direction toward a radial airflow direction. The airflow is radially communicated through a core airflow passage 80 within the fan blade section 72 where the airflow is centrifugally compressed. From the core airflow passage 80, the airflow is diffused and turned once again toward an axial airflow direction toward the annular combustor 30. Preferably, the airflow is diffused axially forward in the engine 10, however, the airflow may alternatively be communicated in another direction.
The tip turbine engine 10 may optionally include a gearbox assembly 90 aft of the fan-turbine rotor assembly 24, such that the fan-turbine rotor assembly 24 rotatably drives the axial compressor 22 via the gearbox assembly 90. In the embodiment shown, the gearbox assembly 90 provides a speed increase at a 3.34-to- one ratio. The gearbox assembly 90 is an epicyclic gearbox, such as the planetary gearbox shown, that is mounted for rotation between the static inner support housing 42 and the static outer support housing 44. The gearbox assembly 90 includes a sun gear 92, which rotates the axial compressor 22, and a planet carrier 94, which rotates with the fan-turbine rotor assembly 24. A plurality of planet gears 93 each engage the sun gear 92 and a rotationally fixed ring gear 95. The planet gears 93 are mounted to the planet carrier 94. The gearbox assembly 90 is mounted for rotation between the sun gear 92 and the static outer support housing 44 through a gearbox forward bearing 96 and a gearbox rear bearing 98. The gearbox assembly 90 may alternatively, or additionally, reverse the direction of rotation and/or may provide a decrease in rotation speed.
In operation, the actuator 55 selectively moves the end of the actuator lever 60 to cause the torque rod 56 to pivot on its longitudinal axis within the inlet guide vane 18. This causes rotation of the activation ring 57 about the engine centerline A, which in turn causes each of the activation levers 59 to pivot and cause rotation of the corresponding compressor vane 54a about its axis P to a selected angle. The selected angle of the compressor vane 54a adjusts the core airflow entering the axial compressor 22, where it is compressed by the compressor blades 52a-c. The compressed air from the axial compressor 22 enters the inducer section 66 in a direction generally parallel to the engine centerline A, and is then turned by the inducer section 66 radially outwardly through the core airflow passage 80 of the hollow fan blades 28. The airflow is further compressed centrifugally in the hollow fan blades 28 by rotation of the hollow fan blades 28. From the core airflow passage 80, the airflow is turned and diffused axially forward in the engine 10 into the annular combustor 30. The compressed core airflow from the hollow fan blades 28 is mixed with fuel in the annular combustor 30 and ignited to form a high-energy gas stream.
The high-energy gas stream is expanded over the plurality of tip turbine blades 34 mounted about the outer periphery of the fan-turbine rotor assembly 24 to drive the fan-turbine rotor assembly 24, which in turn rotatably drives the axial compressor 22 either directly or via the optional gearbox assembly 90. The fan- turbine rotor assembly 24 discharges fan bypass air axially aft to merge with the core airflow from the turbine 32 in an exhaust case 106.
A plurality of exit guide vanes 108 are located between the static outer support housing 44 and the rotationally fixed static outer support structure 14 to guide the combined airflow out of the engine 10 and provide forward thrust. An exhaust mixer 110 mixes the airflow from the turbine blades 34 with the bypass airflow through the fan blades 28.
In the present invention, because the actuator 55 is located remotely from the variable compressor vanes 54a, the size of the splitter 40 and compressor case 50 can be minimized. Routing the linkage between the actuator 55 and the compressor vane 54a within the interiors of the nacelle 12, inlet guide vane 18 and splitter 40 does not require any additional space and provides easy access to the various components for service. In accordance with the provisions of the patent statutes and jurisprudence, exemplary configurations described above are considered to represent a preferred embodiment of the invention. However, it should be noted that the invention can be practiced otherwise than as specifically illustrated and described without departing from its spirit or scope. For example, there are many configurations of linkages, rigid and/or flexible, that could be used to connect the remote actuator 55 to the compressor vane 54a. Also, although the remote actuator 55 has been shown in connection with a tip turbine engine 10, it could also be used in conventional or other turbine engines.

Claims

1. A turbine engine comprising: a compressor case including a plurality of vanes disposed about a longitudinal axis of the compressor case, at least one of the vanes pivotable about a pivot axis, a bypass air flow path disposed radially outwardly of the compressor case; and an actuator disposed radially outward of at least a portion of the bypass air flow path, the actuator mechanically coupled to the at least one vane to selectively pivot the at least one vane about the pivot axis.
2. The turbine engine of claim 1 further including: an inlet guide vane extending between a static outer structure and a static inner structure; and a linkage disposed within the inlet guide vane, the linkage mechanically coupling the actuator to the at least one of the vanes.
3. The turbine engine of claim 2 wherein the linkage includes a ring disposed about a longitudinal axis of the compressor case, rotation of the ring about the axis of the compressor case selectively pivoting the at least one of the vanes about the pivot axis.
4. The turbine engine of claim 3 wherein the linkage further includes a lever coupled between the ring and a shaft on the pivot axis of the at least one vane.
5. The turbine engine of claim 1 further including a fan rotatable about an axis coaxial with the longitudinal axis of the compressor case, the fan generating a bypass airflow through the bypass air flow path.
6. The turbine engine of claim 5 wherein the actuator is disposed radially outward of all of the bypass airflow.
7. The turbine engine of claim 5 wherein the fan is disposed axially aft of the compressor case.
8. The turbine engine of claim 5 further including: an inlet guide vane extending between a static outer structure and a static inner structure; and a linkage disposed within the inlet guide vane, the linkage mechanically coupling the actuator to the at least one of the vanes.
9. The turbine engine of claim 8 wherein the linkage includes a ring disposed about the axis of the compressor case, rotation of the ring about the axis selectively pivoting the at least one vane about the pivot axis.
10. The turbine engine of claim 9 wherein the linkage further includes a lever coupled between the ring and a shaft on the pivot axis of the at least one vane.
11. A compressor vane assembly for an axial compressor of a turbine engine comprising: a vane pivotable about a pivot axis; and an actuator for selectively pivoting the vane about the pivot axis, the actuator spaced away from the vane in a direction having a component generally parallel to the pivot axis.
12. The compressor vane assembly of claim 11 wherein the actuator is spaced from the vane by a distance that is substantially greater than a vane height of the vane as measured along the pivot axis.
13. The compressor vane assembly of claim 11 wherein the actuator is radially outward of a bypass airflow path and the vane is radially inward of the bypass airflow path.
14. The compressor vane assembly of claim 11 further including a linkage operatively connecting the actuator to the vane.
15. The compressor vane assembly of claim 14 wherein the linkage includes an elongated rigid connector.
16. The compressor vane assembly of claim 14 wherein the linkage includes a torque rod transmitting motion rotatably about a major axis of the torque rod.
17. The compressor vane assembly of claim 14 wherein the linkage is disposed within an inlet guide vane.
18. The compressor vane assembly of claim 14 wherein the linkage includes a ring disposed about a ring axis generally transverse to the pivot axis of the vane, rotation of the ring about the ring axis causing pivoting of the vane.
19. A method for actuating a variable guide vane on a compressor, the method including the steps of: generating an actuation force at a remote location from a variable guide vane having a pivot axis; transmitting the actuation force to the pivot axis; and selectively pivoting the guide vane about the pivot axis with the transmitted actuation force.
20. The method of claim 19 wherein the guide vane is one of a plurality of guide vanes disposed circumferentially about a compressor axis, each having a different said pivot axis generally transverse to a longitudinal axis of the compressor, the method further including the steps of: rotating a ring about the compressor axis with the transmitted actuation force; and pivoting the plurality of guide vanes about their respective pivot axes by rotation of the ring.
PCT/US2004/039974 2004-12-01 2004-12-01 Compressor variable stage remote actuation for turbine engine WO2006059972A1 (en)

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PCT/US2004/039974 WO2006059972A1 (en) 2004-12-01 2004-12-01 Compressor variable stage remote actuation for turbine engine
DE602004019710T DE602004019710D1 (en) 2004-12-01 2004-12-01 REMOTE CONTROL FOR AN ADJUSTABLE STAGE OF A COMPRESSOR FOR A TURBINE ENGINE
US11/719,594 US7934902B2 (en) 2004-12-01 2004-12-01 Compressor variable stage remote actuation for turbine engine

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US20090155057A1 (en) 2009-06-18
EP1831530B1 (en) 2009-02-25
US7934902B2 (en) 2011-05-03
DE602004019710D1 (en) 2009-04-09
EP1831530A1 (en) 2007-09-12

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