EP2604810A2 - Module d'actionneur intégré pour moteur de turbine à gaz - Google Patents

Module d'actionneur intégré pour moteur de turbine à gaz Download PDF

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Publication number
EP2604810A2
EP2604810A2 EP12196780.6A EP12196780A EP2604810A2 EP 2604810 A2 EP2604810 A2 EP 2604810A2 EP 12196780 A EP12196780 A EP 12196780A EP 2604810 A2 EP2604810 A2 EP 2604810A2
Authority
EP
European Patent Office
Prior art keywords
actuator
engine
variable vane
recited
vane set
Prior art date
Legal status (The legal status 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 status listed.)
Withdrawn
Application number
EP12196780.6A
Other languages
German (de)
English (en)
Other versions
EP2604810A3 (fr
Inventor
Gabriel L. Suciu
Brian D. Merry
Christopher M. Dye
James S. Elder
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
RTX Corp
Original Assignee
United Technologies Corp
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
Priority claimed from US13/327,722 external-priority patent/US9097137B2/en
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP2604810A2 publication Critical patent/EP2604810A2/fr
Publication of EP2604810A3 publication Critical patent/EP2604810A3/fr
Withdrawn legal-status Critical Current

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    • 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
    • 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
    • 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
    • F04D27/0207—Surge control by bleeding, bypassing or recycling fluids
    • F04D27/0215—Arrangements therefor, e.g. bleed or by-pass valves
    • 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

Definitions

  • the present invention relates to a gas turbine engine integrated actuator module.
  • variable compressor vane system Gas turbine engine performance is typically enhanced through a variable compressor vane system to effectively utilize engine power capacity and enhance transitional performance.
  • the variable compressor vane system typically includes a low pressure compressor variable vane set and a high pressure variable compressor vane set. Each variable vane in each set is rotated in unison through a crank arm linkage. Each crank arm in a set is linked together through a unison ring located circumferentially around the respective compressor case. Each unison ring is rotated by an individual respective actuator to operate the respective variable vane set.
  • each actuator is individually mounted in various locations about the engine case structure such that each actuator requires a separate individual mount platform and hardware. Relatively significant amounts of space within the engine core nacelle and weight redundancies may thereby be generated.
  • An actuator module for a gas turbine engine includes an actuator housing, a first actuator mounted within the actuator housing, the first actuator operable to actuate a first variable vane set, and a second actuator mounted within the actuator housing, the second actuator operable to actuate a second variable vane set.
  • the actuator module may further comprise a third actuator mounted within the actuator housing, the third actuator operable to actuate a bleed valve system.
  • the bleed valve system may comprise a 2.5 bleed valve actuator system.
  • the actuator housing may be mountable to an accessory gearbox. Additionally or alternatively, the actuator housing may be mountable to an engine static structure. Additionally or alternatively, the actuator housing may be mountable to an intermediate case (IMC).
  • IMC intermediate case
  • the first variable vane set may be a low pressure compressor variable vane set.
  • the second variable vane set may be a high pressure compressor variable vane set.
  • a gas turbine engine includes a static structure, a core engine, a gear system supported by the static structure, a fan section driven by the core engine through the gear system, an actuator housing mounted to the static structure, and a first actuator mounted within the actuator housing, the first actuator operable to actuate a first variable vane set of the core engine.
  • the first variable vane set may be a low pressure compressor variable vane set.
  • the gas turbine engine may further comprise a second actuator mounted within the actuator housing, the second actuator may be operable to actuate at least one second variable vane set.
  • the second variable vane set may be a high pressure compressor variable vane set.
  • the gear system may define a gear reduction ratio of greater than or equal to about 2.3.
  • the gear system may define a gear reduction ratio of greater than or equal to about 2.5.
  • the gear system may define a gear reduction ratio of greater than or equal to 2.5.
  • the core engine may include a low pressure turbine which defines a pressure ratio that is greater than about five (5).
  • the core engine may include a low pressure turbine which defines a pressure ratio that is greater than five (5).
  • the fan section may be mounted within a fan nacelle and the core engine may be mounted within a core nacelle, the fan nacelle may be mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow, the fan bypass airflow may define a bypass ratio greater than about six (6).
  • bypass airflow may define a bypass ratio greater than about ten (10).
  • bypass flow may define a bypass ratio greater than ten (10).
  • Figure 1A illustrates a general partial fragmentary schematic view of a gas turbine engine 10 suspended from an engine pylon 12 within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation.
  • the engine 10 includes a core engine within a core nacelle C that houses a low spool 14 and high spool 24.
  • the low spool 14 includes a low pressure compressor 16 and low pressure turbine 18.
  • the low spool 14 drives a fan section 20 connected to the low spool 14 either directly or through a gear train 25.
  • the high spool 24 includes a high pressure compressor 26 and high pressure turbine 28.
  • a combustor 30 is arranged between the high pressure compressor 26 and high pressure turbine 28.
  • the low and high spools 14, 24 rotate about an engine axis of rotation A.
  • the engine 10 in one non-limiting embodiment is a high-bypass geared architecture aircraft engine.
  • the engine 10 bypass ratio is greater than about six (6) to ten (10)
  • the gear train 22 is an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3
  • the low pressure turbine 18 has a pressure ratio that is greater than about 5.
  • the bypass ratio is greater than ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 16.
  • the gear train 25 may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio approximately 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
  • the fan section 20 communicates airflow into the core nacelle C to the low pressure compressor 16.
  • Core airflow compressed by the low pressure compressor 16 and the high pressure compressor 26 is mixed with the fuel in the combustor 30 where it is ignited, and burned.
  • the resultant high pressure combustor products are expanded through the high pressure turbine 28 and low pressure turbine 18.
  • the turbines 28, 18 are rotationally coupled to the compressors 26, 16 respectively to drive the compressors 26, 16 in response to the expansion of the combustor product.
  • the low pressure turbine 18 also drives the fan section 20 through gear train 25.
  • the engine static structure 44 generally has substructures including a case structure often referred to as the engine backbone.
  • the engine static structure 44 generally includes a fan case 46, an intermediate case (IMC) 48, a high pressure compressor case 50, a combustor case 52A, a high pressure turbine case 52B, a thrust case 52C, a low pressure turbine case 54, and a turbine exhaust case 56 ( Figure 1B ).
  • the combustor case 52A, the high pressure turbine case 52B and the thrust case 52C may be combined into a single case. It should be understood that this is an exemplary configuration and any number of cases, and case arrangements may be utilized.
  • the fan section 20 includes a fan rotor 32 with a plurality of circumferentially spaced radially outwardly extending fan blades 34.
  • the fan blades 34 are surrounded by the fan case 46.
  • the core engine case structure is secured to the fan case 46 at the IMC 48 which includes a multiple of circumferentially spaced radially extending struts 40 which radially span the core engine case structure and the fan section 20.
  • the multiple of circumferentially spaced radially extending struts 40 are often generically referred to as Fan Exit Guide Vanes (FEGVs).
  • the engine static structure 44 further supports a bearing system upon which the turbines 28, 18, compressors 26, 16 and fan rotor 32 rotate.
  • a #1 fan dual bearing 60 which rotationally supports the fan rotor 32 is axially located generally within the fan case 46.
  • the #1 fan dual bearing 60 is preloaded to react fan thrust forward and aft (in case of surge).
  • a #2 LPC bearing 62 which rotationally supports the low spool 14 is axially located generally within the intermediate case (IMC) 48.
  • the #2 LPC bearing 62 reacts thrust.
  • a #3 HPC bearing 64 which rotationally supports the high spool 24 and also reacts thrust.
  • the #3 HPC bearing 64 is also axially located generally within the IMC 48 just forward of the high pressure compressor case 50.
  • a #4 bearing 66 which rotationally supports a rear segment of the low spool 14 reacts only radial loads.
  • the #4 bearing 66 is axially located generally within the thrust case 52C in an aft section thereof.
  • a #5 bearing 68 rotationally supports the rear segment of the low spool 14 and reacts only radial loads.
  • the #5 bearing 68 is axially located generally within the thrust case 52C just aft of the #4 bearing 66. It should be understood that this is an exemplary configuration and any number of bearings may be utilized.
  • the #4 bearing 66 and the #5 bearing 68 are supported within a mid-turbine frame (MTF) 70 to straddle radially extending structural struts 72 which are preloaded in tension.
  • the MTF 70 provides aft structural support within the thrust case 52C for the #4 bearing 66 and the #5 bearing 68 which rotatably support the spools 14, 24.
  • a dual rotor engine such as that disclosed in the illustrated embodiment typically includes a forward frame and a rear frame that support the main rotor bearings.
  • the intermediate case (IMC) 48 also includes the radially extending struts 40 which are generally radially aligned with the #2 LPC bearing 62. It should be understood that various engines with various case and frame structures will benefit from the present invention.
  • the turbofan gas turbine engine 10 is mounted to aircraft structure such as an aircraft wing through a mount system 80 attachable by the pylon 12.
  • the mount system 80 includes a forward mount 82 and an aft mount 84.
  • the forward mount 82 is secured to the IMC 48 and the aft mount 84 is secured to the MTF 70 at the thrust case 52C.
  • the forward mount 82 and the aft mount 84 are arranged in a plane containing the axis A of the engine 10.
  • an accessory gearbox 90 may be mounted to the intermediate case (IMC) 48. That is the accessory gearbox 90 may be mounted to the intermediate case (IMC) 48 or be formed integral therewith. It should be understood that the accessory gearbox 90 may be mounted anywhere on the engine static structure 44. In one non-limiting embodiment the accessory gearbox 90 is located axially between the low pressure compressor 16 and the high pressure compressor 26.
  • the accessory gearbox 90 provides significant radial area within the core nacelle (C) inboard of the struts 40 to support accessory engine components such as, for example only, a starter/generator (SG), a hydraulic pump (HP), an oil pump (OP), an integrated oil tank (OT), a fuel pump (FP) and others which thereby saves weight and space within the core nacelle (C).
  • accessory engine components such as, for example only, a starter/generator (SG), a hydraulic pump (HP), an oil pump (OP), an integrated oil tank (OT), a fuel pump (FP) and others which thereby saves weight and space within the core nacelle (C).
  • SG starter/generator
  • HP hydraulic pump
  • OP oil pump
  • OT integrated oil tank
  • FP fuel pump
  • the engine static structure 44 includes a variable compressor vane system 91 which may include a multiple of low pressure compressor variable vane sets 92A, 92B, 92C and a multiple of high pressure compressor variable vane sets 96A, 96B, 96C, 96D.
  • the engine static structure 44 may also include a bleed valve system such as a 2.5 bleed valve actuator system 94.
  • the intermediate case (IMC) 48 supports the multiple of low pressure compressor variable vane sets 92A, 92B, 92C.
  • the intermediate case (IMC) 48 may also support the 2.5 bleed valve actuator system 94. It should be understood that the 2.5 bleed valve actuator system 94 is located generally between the 2 nd and 3 rd stage, but other bleed valve actuator systems may alternatively or additionally benefit herefrom.
  • the high pressure compressor case 50 supports the multiple of high pressure compressor variable vane sets 96A, 96B, 96C. It should be understood that any number of compressor variable vane sets may alternatively or additionally be provided.
  • an actuator module 98 is mounted between lobes of the accessory gearbox 90 in one non-limiting embodiment.
  • the actuator module 98 generally includes a common actuator housing 100 having a multiple of actuators 102, 104, 106 contained therein.
  • Each actuator 102, 104, 106 such as a hydraulic, pneumatic, or electric actuator, for example, drives the respective low pressure compressor variable vane sets 92A, 92B, 92C; 2.5 bleed valve actuator system 94; and the high pressure compressor variable vane sets 96A, 96B, 96C, 96D.
  • Each actuator 102, 104, 106 may be connected to various linkages and be actuated independently as required through a control.
  • the actuator 102 drives a linkage system 108 such as a bell crank mechanism to operate the low pressure compressor variable vane sets 92A, 92B, 92C.
  • the actuator 104 drives a linkage system 110 such as an actuator ring to operate the 2.5 bleed valve actuator system 94.
  • the actuator 106 drives a linkage system 112 such as a bell crank mechanism to operate the high pressure compressor variable vane sets 96A, 96B, 96C, 96D. It should be understood that any multiple of actuators may be contained within the actuator module 98 to operate various additional or alternative engine systems.
  • an actuator module 98' is mounted directly to the intermediate case (IMC) 48 in another non-limiting embodiment.
  • the actuator module 98' is located axially between the low pressure compressor 16 and the high pressure compressor 26.
  • the actuator module 98' is also radially located in an annulus defined between the multiple of circumferentially spaced radially extending struts 40 and a multiple of struts 116 within an inner frame 118 of the intermediate case (IMC) 48.
  • the inner frame 118 may provide a forward structural support for the #2 LPC bearing 62 which rotatably support the spools 14, 24 within the intermediate case (IMC) 48 which also includes the radially extending struts 40.
  • Each actuator may be serviced independently by removing the individual actuator parts from the housing or by removing and replacing the actuator module as a unit.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Life Sciences & Earth Sciences (AREA)
  • Sustainable Development (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Control Of Turbines (AREA)
EP12196780.6A 2011-12-15 2012-12-12 Module d'actionneur intégré pour moteur de turbine à gaz Withdrawn EP2604810A3 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/327,722 US9097137B2 (en) 2008-06-12 2011-12-15 Integrated actuator module for gas turbine engine

Publications (2)

Publication Number Publication Date
EP2604810A2 true EP2604810A2 (fr) 2013-06-19
EP2604810A3 EP2604810A3 (fr) 2013-09-18

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EP12196780.6A Withdrawn EP2604810A3 (fr) 2011-12-15 2012-12-12 Module d'actionneur intégré pour moteur de turbine à gaz

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104747293A (zh) * 2013-12-27 2015-07-01 中航商用航空发动机有限责任公司 一种涡轮风扇发动机所用起动机的选型方法

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2098704A2 (fr) * 2008-03-05 2009-09-09 United Technologies Corporation Système de gestion de flottement de soufflante dans un turboréacteur à double flux

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8210800B2 (en) * 2008-06-12 2012-07-03 United Technologies Corporation Integrated actuator module for gas turbine engine

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2098704A2 (fr) * 2008-03-05 2009-09-09 United Technologies Corporation Système de gestion de flottement de soufflante dans un turboréacteur à double flux

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN104747293A (zh) * 2013-12-27 2015-07-01 中航商用航空发动机有限责任公司 一种涡轮风扇发动机所用起动机的选型方法
CN104747293B (zh) * 2013-12-27 2017-04-19 中航商用航空发动机有限责任公司 一种涡轮风扇发动机所用起动机的选型方法

Also Published As

Publication number Publication date
EP2604810A3 (fr) 2013-09-18

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