EP3464833A2 - Procédé et système pour turbine à gaz à deux bâtis - Google Patents

Procédé et système pour turbine à gaz à deux bâtis

Info

Publication number
EP3464833A2
EP3464833A2 EP17817936.2A EP17817936A EP3464833A2 EP 3464833 A2 EP3464833 A2 EP 3464833A2 EP 17817936 A EP17817936 A EP 17817936A EP 3464833 A2 EP3464833 A2 EP 3464833A2
Authority
EP
European Patent Office
Prior art keywords
turbine
low pressure
frame member
gas turbine
engine
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Withdrawn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
EP17817936.2A
Other languages
German (de)
English (en)
Inventor
Brandon Wayne Miller
Thomas Ory MONIZ
Jeffrey Donald Clements
Joseph George ROSE
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP3464833A2 publication Critical patent/EP3464833A2/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • 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
    • F02K3/06—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 with front fan
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
    • F01D25/162—Bearing supports
    • 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
    • F02C7/06—Arrangements of bearings; Lubricating
    • 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
    • F05D2260/403—Transmission of power through the shape of the drive components
    • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
    • F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the field of the disclosure relates generally to gas turbine engines and, more particularly, to a method and system for a reduced frame gas turbine engine assembly.
  • Gas turbine engine assemblies using integral drive with high speed booster compressors directly coupled to a low pressure (LP) turbine require a frame between the booster compressor and fan assembly.
  • An additional frame is typically required aft of a LP turbine.
  • These frames tend to increase the length of the gas turbine engine assembly and thereby also tend to increase weight and cost of the gas turbine engine assembly.
  • the booster compressor rotating at a high speed for example, approximately the LP turbine speed, highly loads the booster compressor causing it to operate at a non-optimal pressure ratio than might otherwise be attained.
  • a gas turbine engine assembly includes a core engine, a low pressure turbine, a low pressure compressor, a fan assembly, and an engine frame assembly.
  • the core engine includes a high pressure compressor, a combustor, and a high pressure (HP) turbine in a serial flow arrangement.
  • the low pressure turbine is positioned axially aft. of the core engine and includes a plurality of stages of stator vanes and rotor blades. A last stage of rotor blades of the plurality of stages of stator vanes and rotor blades of the low pressure turbine includes a low swirl outlet rotor blade stage.
  • the low- pressure compressor is positioned axially forward of the core engine and rotatably coupled to the low pressure turbine through a gearbox.
  • the low pressure compressor is aligned axially with die gearbox and positioned radially outward from the gearbox.
  • the fan assembly is directly coupled to the low pressure compressor such that the fan assembly and the low pressure compressor rotate at the same speed.
  • the engine frame assembly includes a forward fan frame member positioned axially between the low pressure compressor and the high pressure compressor, and positioned axially aft of the gearbox.
  • the engine frame assembly also includes a turbine center frame member positioned axially between the high pressure turbine and the low pressure turbine.
  • a method of assembling a two-frame gas turbine engine includes providing a core gas turbine engine including a high pressure compressor, a combustor, and a high pressure turbine coupled together in serial flow communication.
  • the method also includes coupling the core gas turbine engine to a forward fan frame member positioned axially forward of the core engine.
  • the method further includes coupling the core gas turbine engine to a turbine center frame member.
  • the turbine center frame member is coupled to the core gas turbine engine axially aft of the high pressure turbine.
  • the method also includes coupling a low pressure turbine to a first shaft axially aft of the turbine cente frame member.
  • the method also includes coupling an input of a gearbox to the first shaft axially forward of the forward fan frame member.
  • the method further includes coupling a fan assembly and a lo pressure compressor to an output of the gearbox axially forward of the forward fan frame member.
  • gas turbine engine assembly configured to drive a bladed rotatable member of a fan assembly.
  • the gas turbine engine includes a core engine, a low pressure turbine, a low pressure compressor, and an engine frame assembly.
  • the core engine includes a high pressure compressor, a combustor, and a high pressure turbine in a serial flow arrangement.
  • the low pressure turbine is positioned axially aft of the core engine and includes a plurality of stages of stator vanes and rotor blades.
  • a last stage of rotor blades of the plurality of stages of stator vanes and rotor blades of the low pressure turbine includes a low swirl outlet rotor blade stage.
  • the low pressure compressor is positioned axially forward of the core engine and rotatably coupled to the low pressure turbine through a gearbox.
  • the low pressure compressor is aligned axially with the gearbox.
  • the low pressure compressor is positioned radially outward from the gearbox.
  • the engine frame assembly includes a forward fan frame member positioned axially between the low pressure compressor and the high pressure compressor.
  • the gearbox is positioned axially forward of the forward frame.
  • the gearbox is positioned radially inward.
  • FIG. 1 is a perspective view of an aircraft.
  • FIG. 2 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure that may be used with the aircraft shown in FIG. 1.
  • FIG. 3 is a side elevation view of the turbofan engine shown in FIGS. 1 and 2.
  • FIG. 4 is a side elevation view of an aft. portion of the turbofan engine shown in FIGS. 1, 2, and 3.
  • FIG. 5 is a flow diagram of a method of constructing the turbofan engine shown in FIGS. 1 , 2, and 3.
  • Approximating language may be applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about,” “approximately,” and “substantially,” are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value.
  • range limitations may be combined and/or interchanged; such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.
  • Embodiments of the gas turbine engine and method of assembly described herein provide a cost effective method for providing a gas turbine engine assembly that is shorter and lighter than known engines of similar capability.
  • Gas turbine engine assemblies using integral drive with high speed booster compressors typically require a frame between the booster compressor and the fan assembly.
  • An additional frame is typically required aft of a LP turbine. These frames tend to increase the length of the gas turbme engine assembly and thereby also tend to increase weight and cost of the gas turbine engine assembly.
  • LP low swirl low pressure
  • the gas turbine engine assembly includes a core engine including a high pressure compressor, a combustor, and a high pressure (HP) turbine in a serial flow arrangement.
  • a low swirl LP turbine is positioned axially aft of the core engine and a low pressure LP compressor is positioned axially forward of the core engine.
  • the LP compressor is rotatably coupled to the LP turbine through a gearbox, which may be a speed changing gearbox or a reduction gearbox, and is aligned axially with the gearbox.
  • the LP compressor is positioned radially outward from the gearbox.
  • the gas turbine engine assembly also includes an engine frame assembly including only two frames, a forward fan frame member and a turbine center frame member.
  • Tire forward fan frame member is positioned axially between the low pressure compressor and the high pressure compressor and axially aft of the gearbox.
  • the forward fan frame member is configured to support the low pressure compressor and the high pressure compressor.
  • the turbine center frame member is positioned axially between the HP turbine and the LP turbine.
  • the gas turbine engine assembly includes a longitudinal centerline and the forward fan frame member and the turbine center frame member are coaxially aligned with the centerline.
  • the core engine includes a high pressure rotor shaft and the gas turbine engine assembly includes a low pressure rotor shaft.
  • the turbine rear frame member is configured to rotatably support an aft end portion of the high pressure rotor shaft and an aft end portion of the low pressure rotor shaft.
  • the engine fan assembly is directly coupled to the low pressure compressor and consequently the fan assembly and the low pressure compressor rotate at the same speed. Because the fan assembly and the low pressure compressor are coupled to the LP turbine through the gearbox, the fan assembly and the low- pressure compressor may rotate at a speed that is the same or that is different than a speed of rotation of the LP turbine depending on the configuration of the gearbox.
  • the fan assembly and the low pressure compressor rotate at a first speed and the LP turbine rotates at a second speed.
  • the first and second speeds can be the same, the first speed can be greater than or less than the second speed depending, in some embodiments, on a configuration of the gearbox.
  • a method of assembling a gas turbine engine includes providing a core engine including a high pressure compressor, a cornbusior, and a turbine coupled together in axial flow communication, coupling a low s irl LP turbine to a first shaft axially aft of the core engine, coupling an input of a gearbox to the first shaft axially forward of the core engine, and coupling a fan assembly and a booster compressor to an output of the gearbox axially forward of the core engine.
  • Embodiments described herein disclose a booster compressor to the fan assembly in an integral drive configuration .
  • the boost power is sent through the gearbox from the low swirl LP turbine to the fan and booster as a common spool.
  • embodiments described herein disclose including a low swirl low pressure turbine rotor blade stage which eliminates the need for a turbine rear frame or outlet guide vanes to reduce the swirl of exhaust gases.
  • Such a configuration eliminates the need for two frames of the engine and shortens the engine.
  • the engine configuration described herein permits increasing the fan speed such that the booster compressor speed is increased thus reducing loading on the booster compressor and improving pressure ratio possible from the booster compressor.
  • increasing fan assembly- speed is beneficial because this makes the fan more distortion tolerant or operable.
  • improvements in the fan tip speed range combined with lower fan pressure ratio result from, the described configuration.
  • FIG. 1 is a perspective view of an aircraft 100.
  • aircraft 100 includes a fuselage 102 that includes a nose 104, a tail 106, and a hollow, elongate body 108 extending therebetween.
  • Aircraft 100 also includes a wing 110 extending away from fuselage 102 in a lateral direction 112.
  • Wing 110 includes a forward leading edge 1 14 in a direction 116 of motion of aircraft 100 during normal flight and an aft. trailing edge 1 18 on an opposing edge of wing 110.
  • Aircraft 100 further includes at least one engine assembly 120, which may be embodied in a gas turbine engine and/or a gas turbine engine of the high bypass turbofan type or the like, configured to drive a bladed rotatable member 122 or fan to generate thrust.
  • Engine assembly 120 is coupled to at least one of wing 110 and fuselage 102, for example, in a pusher configuration (not shown) proximate tail 106.
  • FIG. 2 is a schematic cross-sectional view of gas turbine engine assembly 120 in accordance with an exemplary embodiment of the present disclosure.
  • gas turbine engine assembly 120 is embodied in a high bypass turbofan jet engine.
  • turbofan engine assembly 120 defines an axial direction A (extending parallel to a longitudinal axis 202 provided for reference) and a radial direction R.
  • turbofan 120 includes a fan assembly 204 and a core engine 206 disposed downstream from fan assembly 204.
  • core engine 206 includes an approximately tubular outer casing 208 that defines an annular inlet 220.
  • Outer casing 208 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor 222 and a high pressure (HP) compressor 224; a combustion section 226; a turbine section including a high pressure (HP) turbine 228 and a lo pressure (LP) turbine 230: and a jet exhaust nozzle section 232.
  • a high pressure (HP) shaft or spool 234 drivingly connects HP turbine 228 to HP compressor 224.
  • a low pressure (LP) shaft, or spool 236 drivingly connects LP turbine 230 to LP compressor 222.
  • the compressor section, combustion section 226, turbine section, and nozzle section 232 together define a core air flowpath 237.
  • fan assembly 204 includes a variable pitch fan 238 having a plurality of fan blades 240 coupled to a disk 242 in a spaced apart relationship.
  • a variable pitch fan is shown in FIG. 2, other fan configurations are anticipated including a configuration as shown in FIG. 3 without a variable pitch fan .
  • Fan blades 240 extend radially outwardly from disk 242. Each fan blade 240 is rotatable relative to disk 242 about a pitch axis P by virtue of fan blades 240 being operatively coupled to a suitable pitch change mechanism (PCM) 244 configured to vary the pitch of fan blades 240.
  • PCM pitch change mechanism
  • pitch change mechanism (PCM) 244 is configured to collectively van' the pitch of fan blades 240 in unison.
  • Fan blades 240, disk 242, and pitch change mechanism 244 are together rotatable about longitudinal axis 202 by LP shaft 236 across a power gearbox 246,
  • Power gearbox 246 includes a plurality of gears for adjusting the rotational speed of fan 238 relative to LP shaft 236 to a more efficient rotational fan speed.
  • Disk 242 is covered by rotatable front hub 248 aerodynamically contoured to promote an airflow through the plurality of fan blades 240.
  • fan assembly 204 includes an annular fan casing or outer nacelle 250 that circumferentially surrounds fan 238 and/or at least a portion of core engine 206.
  • nacelle 250 is configured to be supported relative to core engine 206 by a plurality of circurnferentialiy-spaced outlet guide vanes 252 coupled to a forward fan frame member 259.
  • a downstream section 254 of nacelle 250 may extend over an outer portion of core engine 206 so as to define a bypass airflow passage 256 therebetween.
  • Gas turbine engine assembly 120 includes an engine frame assembly 257 including, in one embodiment, only two frames, forward fan frame member 259 and a turbine center frame member 261.
  • a frame member supports a bearing and may incorporate an aerodynamic fairing to swirl or de-swirl the air through gas turbine engine assembly 120 during operation.
  • turbine rear frame member 255 is positioned aft of the LP turbine.
  • Forward fan frame member 259 is positioned axially between low pressure compressor 222 and high pressure compressor 224 and axially aft of gearbox 246.
  • Forward fan frame member 259 is configured to support LP compressor 222 and HP compressor 224.
  • Turbine center frame member 261 is positioned axially between HP turbine 228 and LP turbine 230.
  • gas turbine engine assembly 120 includes a longitudinal axis 202 and forward fan frame member 259 and turbine center frame member 261 are coaxially aligned with the centerline.
  • turbine rear frame member 255 is added to provide additional support to LP turbine 230.
  • gas turbine engine assembly 120 includes a three frame engine frame assembly in some embodiments.
  • Core engine 206 includes a high pressure rotor shaft 234 and gas turbine engine assembly 120 includes a low pressure rotor shaft 236.
  • Turbine center frame member 26 ! is configured to rotatably support an aft end portion 239 of HP turbine 228 and a forward end portion 241 of LP turbine 230.
  • a volume of air 258 enters turbofan 120 through an associated inlet 260 of nacelle 250 and/or fan assembly 204.
  • a first portion 262 of volume of air 258 is directed or routed into bypass airflow passage 256 and a second portion 264 of volume of air 258 is directed or routed into core air flowpath 237, or more specifically into LP compressor 222.
  • a ratio between first portion 262 and second portion 264 is commonly referred to as a bypass ratio.
  • the pressure of second portion 264 is then increased as it is routed through high pressure (HP) compressor 224 and into combustion section 226, where it is mixed with fuel and burned to provide combustion gases 266.
  • HP high pressure
  • Combustion gases 266 are routed through HP turbine 228 where a portion of thermal and/or kinetic energy from combustion gases 266 is extracted via sequential stages of HP turbine stator vanes 268 that are coupled to outer casing 208 and HP turbine rotor blades 270 that are coupled to HP shaft or spool 234, thus causing HP shaft or spool 234 to rotate, which then drives a rotation of HP compressor 224.
  • Combustion gases 266 are then routed through LP turbine 230 where a second portion of thermal and kinetic energy is extracted from, combustion gases 266 via sequential stages of LP turbine stator vanes 272 that are coupled to outer casing 208 and LP turbine rotor blades 274 that are coupled to LP shaft or spool 236, which drives a rotation of LP shaft or spool 236 and LP compressor 222 and/or rotation of fan 238.
  • Combustion gases 266 are subsequently routed through jet exhaust nozzle section 232 of core engine 206 to provide propulsive thrust. Simultaneously, the pressure of first portion 262 is substantially increased as first portion 262 is routed through bypass airflow passage 256 before it is exhausted from a fan nozzle exhaust section 276 of turbofan 120, also providing propulsive thrust.
  • HP turbine 228, LP turbine 230, and jet exhaust nozzle section 232 at least partially define a hot gas path 278 for routing combustion gases 266 through core engine 206.
  • Turbofan engine assembly 120 is depicted in FIG. 1 by way of example only, and that in other exemplary embodiments, turbofan engine assembly 120 may have any other suitable configuration including for example, a turboprop engine.
  • FIG. 3 is another schematic cross-sectional view of the turbofan engine assembly 120 (shown in FIGS. 1 and 2).
  • gearbox 246 is positioned axiaily aligned and radially inward of LP compressor 222.
  • Forward fan frame member 259 is positioned axiaily between LP compressor 222 and HP compressor 224. Such relative positions permits eliminating a frame member that is typically found in other gas turbine engines of similar size and configuration.
  • epicyclic gear train 246 is embodied in, for example, an epicyclic gear and a compound gear.
  • Forward fan frame member 259 provides support for fan assembly 204, LP compressor 222, gearbox 246, and a forward end portion 247 of HP compressor 224.
  • turbofan engine assembly 120 includes three frames wherein turbine rear frame member includes an airfoil portion configured to deswirl exhaust gases exiting LP turbine 230. In other embodiments, turbofan engine assembly 120 includes only two frames, forward fan frame member 259 and turbine center frame member 261.
  • turbofan engme assembly 120 does not include aft frame aft 255
  • the deswirling function typically provided by aft frame aft 255 is provided for elsewhere, for example, by the addition of a stage to LP turbine 230.
  • This final stage is configured to deswirl the exhaust gases channeled from the preceding stages of LP turbine 230.
  • LP turbine 230 includes a low swirl LP turbme last stage configured to deswirl the exhaust gases.
  • FIG. 4 is a schematic cross-sectional view of an aft portion of gas turbine engme 120 in accordance with an exemplary embodiment of the present disclosure.
  • LP turbine 230 includes four stages of LP turbine rotor blades 402, 404, 406, and 408 coupled to LP shaft 236 and four stages of LP turbine stator vanes 409, 410, 412, and 414.
  • LP turbine 230 may include more or fewer stages of LP turbme rotor blades, such as one, two, three, or five LP turbine rotor blades, or any other suitable number of LP turbine rotor blades that enables LP turbine 230 to function as described herein.
  • LP turbine 230 may include more or fewer stages of LP turbine stator vanes, such as one, two, three, or five LP turbine stator vanes, or any other suitable number of LP turbine stator vanes that enables LP turbine 230 to function as described herein.
  • combustion gases 266 are routed sequentially to a first LP turbine stator vane 409, a first LP turbine rotor blade stage 402, a second LP turbme stator vane 410, a second LP turbine rotor blade stage 404, a third LP turbine stator vane 412, a third LP turbine rotor blade stage 406, a fourth LP turbine stator vane 414, and a fourth LP turbine rotor blade stage 408, [0034]
  • Combustion gases 266 include LP turbine stator vane velocities 415, 418, 422, and 426 and LP turbine rotor blade velocities 416, 420, 424, and 428.
  • LP turbine stator vane velocities 418, 422, and 426 and LP turbine rotor blade velocities 416, 420, 424, and 428 each include an axial component and a circumferential component, LP turbine stator vane velocities
  • Fourth LP turbine rotor blade stage 408 is a low swirl LP turbine stage which does not require an outlet guide vane or turning vanes to reduce the swirl of exhaust gases.
  • FIG. 5 is a flow diagram, of a method 500 of constracting a gas turbine engine, such as, gas turbine engine 120 (shown in FIG. 1).
  • Method 500 includes providing 502 core turbine engine 206 including HP compressor 224, combustion section 226, and HP turbine 228 coupled together in serial flo communication.
  • Method 500 also includes coupling 504 core turbine engine 206 to forward fan frame member 259 positioned axially forward of core engine 206.
  • Method 500 further includes coupling 506 core turbine engine 206 to turbine center frame member 261.
  • Turbine center frame member 261 is coupled to core turbine engine 206 axially aft of HP turbine 228.
  • Method 500 also includes coupling 508 LP turbine 230 to LP shaft 236 axially aft of turbine center frame member 261 .
  • Method 500 further includes coupling 510 an input of power gearbox 246 to LP shaft 236 axially forward of forward fan frame member 259.
  • Method 500 also includes coupling 512 fan assembly 206 and LP compressor 222 to an output of power gearbox 246 axially forward of forward fan frame member 259.
  • the above described embodiments of a method and system of a reduced frame gas turbine engine assembly provides a cost effective and reliable means for reducing the length, weight, and cost of the gas turbine engine assembly. More specifically, the methods and systems described herein facilitate optimizing the fan and booster speed independent of the LP turbine speed to allow an optimized pressure ratio and performance from the fan and booster. Also increasing fan assembly speed is beneficial to make the fan more distortion tolerant or operable. Moreover, improvements in the fan tip speed range combined with lower fan pressure ratio result from the described configuration. As a result, the methods and systems described herein facilitate improving the fan tip speed range and permitting a lower fan pressure ratio in a shorter, lighter engine in a cost effective and reliable manner.

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

Abstract

Selon l'invention, une turbine à gaz comprend un bloc réacteur comprenant un compresseur haute pression (HP), une chambre de combustion, et une turbine HP dans un agencement de flux en série. Une turbine basse pression (LP) est disposée axialement à l'arrière du bloc moteur et comprend une pluralité d'étages de pales de rotor. Un dernier étage de pales de rotor comprend un étage de pale de rotor à faible tourbillon. Un compresseur LP est positionné axialement à l'avant du bloc moteur et est accouplé à la turbine LP par une boîte de vitesses. Le compresseur LP est positionné radialement à l'extérieur en partant de la boîte de vitesses. Un ensemble ventilateur est directement accouplé au compresseur LP de telle manière que l'ensemble ventilateur et le compresseur LP tournent à la même vitesse. Un ensemble bâti-moteur comprend un élément de bâti de ventilateur avant disposé axialement entre le compresseur LP et le compresseur HP. L'ensemble bâti-moteur comprend également un élément de bâti central de turbine placé entre la turbine HP et la turbine LP.
EP17817936.2A 2016-05-25 2017-05-12 Procédé et système pour turbine à gaz à deux bâtis Withdrawn EP3464833A2 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US201615164726A 2016-05-25 2016-05-25
PCT/US2017/032319 WO2018026408A2 (fr) 2016-05-25 2017-05-12 Procédé et système pour turbine à gaz à deux bâtis

Publications (1)

Publication Number Publication Date
EP3464833A2 true EP3464833A2 (fr) 2019-04-10

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP17817936.2A Withdrawn EP3464833A2 (fr) 2016-05-25 2017-05-12 Procédé et système pour turbine à gaz à deux bâtis

Country Status (3)

Country Link
EP (1) EP3464833A2 (fr)
CN (1) CN109196187B (fr)
WO (1) WO2018026408A2 (fr)

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US11174916B2 (en) 2019-03-21 2021-11-16 Pratt & Whitney Canada Corp. Aircraft engine reduction gearbox
US11268453B1 (en) 2021-03-17 2022-03-08 Pratt & Whitney Canada Corp. Lubrication system for aircraft engine reduction gearbox

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DE102017211649A1 (de) * 2017-07-07 2019-01-10 MTU Aero Engines AG Gasturbine mit einer schnelllaufenden Niederdruckturbine und einem Turbinengehäuse
US11156097B2 (en) * 2019-02-20 2021-10-26 General Electric Company Turbomachine having an airflow management assembly
GB201903257D0 (en) * 2019-03-11 2019-04-24 Rolls Royce Plc Efficient gas turbine engine installation and operation
FR3097012B1 (fr) * 2019-06-06 2022-01-21 Safran Aircraft Engines Procédé de régulation d’une accélération d’une turbomachine
US11560840B2 (en) * 2020-10-16 2023-01-24 General Electric Company Damper engine mount links

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US7513102B2 (en) * 2005-06-06 2009-04-07 General Electric Company Integrated counterrotating turbofan
US7926259B2 (en) * 2006-10-31 2011-04-19 General Electric Company Turbofan engine assembly and method of assembling same
US20130186058A1 (en) * 2012-01-24 2013-07-25 William G. Sheridan Geared turbomachine fan and compressor rotation
US8678743B1 (en) * 2013-02-04 2014-03-25 United Technologies Corporation Method for setting a gear ratio of a fan drive gear system of a gas turbine engine
US8869504B1 (en) * 2013-11-22 2014-10-28 United Technologies Corporation Geared turbofan engine gearbox arrangement
US9932902B2 (en) * 2014-07-15 2018-04-03 United Technologies Corporation Turbine section support for a gas turbine engine
US10221771B2 (en) * 2014-09-24 2019-03-05 United Technologies Corporation Fan drive gear system

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11174916B2 (en) 2019-03-21 2021-11-16 Pratt & Whitney Canada Corp. Aircraft engine reduction gearbox
US11268453B1 (en) 2021-03-17 2022-03-08 Pratt & Whitney Canada Corp. Lubrication system for aircraft engine reduction gearbox

Also Published As

Publication number Publication date
WO2018026408A3 (fr) 2018-04-26
CN109196187A (zh) 2019-01-11
WO2018026408A2 (fr) 2018-02-08
CN109196187B (zh) 2021-12-07

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