WO2019168590A1 - Moteur à turbine à gaz avec système de distribution d'air de refroidissement de turbine - Google Patents

Moteur à turbine à gaz avec système de distribution d'air de refroidissement de turbine Download PDF

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Publication number
WO2019168590A1
WO2019168590A1 PCT/US2019/012452 US2019012452W WO2019168590A1 WO 2019168590 A1 WO2019168590 A1 WO 2019168590A1 US 2019012452 W US2019012452 W US 2019012452W WO 2019168590 A1 WO2019168590 A1 WO 2019168590A1
Authority
WO
WIPO (PCT)
Prior art keywords
self
gas turbine
seal
turbine engine
passage
Prior art date
Application number
PCT/US2019/012452
Other languages
English (en)
Inventor
David May
Constant CHARRETON
Original Assignee
Siemens Aktiengesellschaft
Siemens Energy, Inc.
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 Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Publication of WO2019168590A1 publication Critical patent/WO2019168590A1/fr

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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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/18—Lubricating arrangements
    • F01D25/22—Lubricating arrangements using working-fluid or other gaseous fluid as lubricant
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • F01D11/025—Seal clearance control; Floating assembly; Adaptation means to differential thermal dilatations
    • 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
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • F01D5/082—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades on the side of the rotor disc

Definitions

  • Disclosed embodiments are generally related to turbine engines, and in particular to cooling systems of the turbine engine.
  • a high pressure turbine (HPT) cooling air delivery system typically includes a pre-swirl system.
  • the system takes air from the high pressure compressor and injects it in front of the HPT disc through angled holes or nozzles (pre-swirlers).
  • pre-swirlers angled holes or nozzles
  • the angle is set so that the air tangential velocity ejected through the pre-swirls matches that of the disc speed. This yields minimum air temperature rise when washing the surface of the disc with incoming air.
  • the air also feeds the HPT blade internal passage. For the same reason as the disc, it is desirable to feed air at the same tangential velocity as the blade/disc so the blade is cooled using the lowest possible feed air temperature.
  • the static and rotating surfaces that contain the swirled cooling air i.e. the pre-swirl cavity
  • labyrinth seals There is a certain amount of leakage into the pre-swirl cavity through the lower seal. This leakage is hotter than the pre-swirled air for at least two reasons. First, windage heating of the leakage, due to friction on the multiple teeth. Second, the leakage swirls at a lower tangential velocity (around half) of that of the pre-swirled air which creates an offset between the mixed air swirl and the disc speed.
  • Another variation of arrangement of the gas turbine engine comprises dropping the pressure upstream of the lower labyrinth seal by means of another sealing element so as to reverse the flow direction.
  • the result is that higher temperature air is no longer leaking into the pre-swirl cavity, i.e. only pre-swirled air is feeding the disc and the blade internal passages. This reversed air flow is directed to the rim cavity through a bypass in the pre-swirlers support structure.
  • aspects of the present disclosure relate to gas turbine engines and more particularly to sealing portions of the gas turbine engine.
  • An aspect of the present disclosure is directed to a gas turbine engine.
  • the gas turbine engine includes a first passage formed by an exterior wall of a combustor and an inner wall within the gas turbine engine and a second passage formed by the inner wall and a rotating arm.
  • a pre-swirl cavity is located at a downstream end of the first passage.
  • a first self-adjusting seal is located in the second passage.
  • a second self- adjusting seal located radially outward of the pre-swirl cavity.
  • FIG. 1 is diagrammatic view of a gas turbine engine.
  • Fig. 2 is a cut-away view of a self-adjusting seal.
  • FIG. 3 is a diagrammatic view of a gas turbine engine with self-adjusting seals.
  • Fig. 4 is a close up view of the pre-swirl cavity shown in Fig. 3.
  • Fig. 5 is a close up view of the pre-swirl support structure shown in Fig. 3.
  • the gas turbine engine 10 may include a compressor section 11 for compressing air.
  • the compressed air from the compressor section 11 is conveyed to a combustion section 12, which produces hot combustion gases by burning fuel in the presence of the compressed air from the compressor section 11.
  • the combustion gases are conveyed through a plurality of transition ducts to a turbine section 14 of the engine 10.
  • the turbine section 14 comprises alternating rows of stationary vanes 8 and rotating blades 9.
  • blade disc structures 15 are positioned adjacent to one another in an axial direction.
  • the blade disc structures 15 define a rotor.
  • Each of the blade disc structures 15 supports circumferentially spaced apart blades 9 and each of a plurality of lower stator support structures support a plurality of circumferentially spaced apart vanes 8.
  • the vanes 8 direct the combustion gases from the transition ducts along a hot gas flow path to the blades 9 such that the combustion gases cause rotation of the blades 9, which in turn causes corresponding rotation of the blade disc structures 15 of the rotor.
  • a supply of fluid can supply fluid within the gas turbine engine 10.
  • the fluid may have a temperature of, for example, between about 1000-1200° F.
  • the fluid flows through passages formed between an inner wall 17 and an outer combustor wall 18.
  • the fluid also flows between a rotating arm 16 and inner wall 17.
  • the rotating arm 16 may be connected to the compressor section 11 and the turbine section 14.
  • the rotating arm 16 may also be referred to as a“drive cone” or“shaft.”
  • the first passage 19 is formed between the outer combustor wall 18 and the inner wall 17.
  • the second passage 20 is formed between the rotating arm 16 and the inner wall 17.
  • a pre-swirl structure 33 is located downstream of, and in fluid communication with the first passage 19.
  • the pre-swirl structure 33 comprises pre- swirlers 34, which may be configured, for example, as angled holes or nozzles, which are configured so that the tangential velocity of fluid ejected through the pre-swirl ers
  • the fluid from the first passage 19 flows through pre- swirl ers 34 into a pre-swirl cavity 30 formed between the pre-swirl er structure 33 and the blade disc structure 15.
  • the fluid flow from the second passage 20 may leak into the pre-swirl cavity 30 through a lower seal 35.
  • the lower seal 35 In the shown example, the lower seal
  • a labyrinth seal defined by knife edges formed on rotating blade disc structures 15 which interface with stationary honeycombed structures provided on the lower extension of the pre-swirl structure 33.
  • the area of leakage is shown at position 31.
  • An example of a self-adjusting seal 40, 41 is shown in Fig. 2.
  • Self-adjusting seals 40, 41 of the type that may be employed in the instant invention may be found in U.S. Patent nos. 8,002,285; 8,172,232 and 8,919,781, the contents of which are hereby incorporated by reference..
  • each self-adjusting seal 40, 41 comprises a seal shoe 43 supported by flexible beams 42.
  • the beams 42 which may be, for example, configured as leaf springs, permit the radial position of the seal shoe 43 to adjust by being responsive to the pressures in the surrounding environment.
  • the beams 42 can be made of a material that is both durable and flexible.
  • the beams 42 may be made of stainless steel.
  • the beams 42 may also be made of inconel for high temperature application.
  • the pressure at the location of the self-adjusting seal 40, 41 determines the movement of the self-adjusting seal 40, 41. Too little leakage will result in a small pressure drop which may not be sufficient to activate the self-adjusting seal 40, 41, so the beams 42 will stay at their cold build position. Too much leakage and pressure drop across the self-adjusting seal 40, 41 may result in the seal shoe 43 being pulled into place by the suction pressure, whereby the position of the seal shoe 43 self-adjusts to reduce the leakage.
  • the pre-swirl structure 33 is located downstream of and in fluid communication with the first passage 19.
  • the pre-swirl structure 33 comprises pre-swirlers 34, which may be configured, for example, as angled holes or nozzles, which are configured so that the tangential velocity of the fluid ejected through the pre-swirlers 34 matches that of the disc speed.
  • the fluid from the first passage 19 flows through pre-swirlers 34 into a pre-swirl cavity 30 formed between the pre-swirler structure 33 and the blade disc structure 15.
  • a lower seal 35 is positioned radially inward of the pre-swirlers 34, which may be defined by knife edges formed on rotating blade disc structures 15 and honeycombed structures on the lower extension of the stationary pre-swirl structure 33.
  • Cooling fluid exits the pre-swirl cavity 30 with a velocity component in a direction tangential to the circumferential direction of the gas turbine engine 10.
  • a swirl ratio is defined as the velocity component in the direction tangential to the circumferential direction of the cooling fluid as compared to a velocity component of a rotating shaft in the direction tangential to the circumferential direction.
  • the second self-adjusting seal 41 is oriented so that the seal shoe 43 is located proximate to the transition duct and the first passage 19 within the pre-swirl cavity 30.
  • the beams 42 of the second self-adjusting seal 41 move in response to the pressures within the pre-swirl cavity 30 and surrounding environment.
  • the second self-adjusting seal 41 is located radially outward of the pre- swirl cavity 30 and above the lower seal 35.
  • the second self-adjusting seal 41 may be secured to the inner wall 17 using a bolt 3.
  • the first self-adjusting seal 40 is located in the second passage 20 between the inner wall 17 and the rotating arm 16.
  • the beams 42 of the first self-adjusting seal 40 are located proximate to the inner wall 17.
  • the shoe 43 is located proximate to the rotating arm 16.
  • the first self-adjusting seal 40 may be secured to the inner wall 17 using bolt 4.
  • the first self-adjusting seal 40 seals a leakage flow between the second passage 20 and a collection chamber 21 located downstream of the first self-adjusting seal 40.
  • the first self-adjusting seal 40 thus acts to produce a pressure drop at the collection chamber 21.
  • the reduced pressure in the collection chamber 21 causes fluid to move from the pre-swirl cavity 30 into the collection chamber 21 via a reverse flow 37 across the lower seal 35.
  • Fluid from the collection chamber 21 is bypassed, via a bypass channel 23 into a rim cavity 46, where it contributes to purging and cooling of the rim cavity.
  • the bypass channel 23 allows evacuating the mix of air from the second passage20 and the lower seal 35, without contaminating the pre-swirl cavity 30.
  • the bypass channel 23 may be composed of drillings and pockets in the pre-swirl support structure 27. Fluid from the first passage 19 enters through the pockets 28 into the pre-swirl cavity 30. The flow in the bypass channel 23 enters through the pockets 29 in the pre- swirl support structure 27.
  • Forming part of the bypass channel 23 is a diaphragm 48.
  • the diaphragm 48 isolates the bypass air flow 23 from the leakage flow 47 through the self-adjusting seal 40 that comes from pre-swirl cavity 30.
  • a blade cooling passage 50 is defined through the blade disc structure 15, located downstream of the pre-swirl cavity 30.
  • the second self-adjusting seal 41 prevents or minimizes a leakage flow from the pre-swirl cavity 30 to the rim cavity 46, causing a fluid flow 51 to move from the pre-swirl structure 33 through the blade cooling flow passage 50 into the turbine section 14 in order to provide cooling for the turbine discs 15.
  • the first self-adjusting seal 40 causes the fluid flow to reverse through the lower seal 35 that is located in the pre-swirl cavity 30, causing a reverse flow 37
  • the lower seal 35 may be a labyrinth seal, as illustrated above, or a brush seal.
  • the first self-adjusting seals 40 and the second self-adjusting seal 41 react to the pressure drop across them to maintain a tight running clearance over a wide range of conditions. Unlike labyrinth seals, the first self-adjusting seal 40 and the second self- adjusting seal 41 are effectively insensitive to thermal movements and instead respond to changes in pressure.
  • a gas turbine engine 10 having an air system architecture with a collection chamber 21, bypass channel 23 combined with the first self-adjusting seals 40 and the second self-adjusting seal 41 has several advantages.
  • the first self-adjusting seal 40 and the second self-adjusting seal 41 provide a uniform performance of the gas turbine engine 10 over a wide range of operating conditions.
  • the first self-adjusting seals 40 and the second self-adjusting seal 41 also enable a robust and consistent reverse flow 37 through the lower seal 35, which results in reduced blade cooling air temperature.
  • the resulting temperature control can provide a potential for power increase and/or gain in the life of the gas turbine components.
  • the first self-adjusting seals 40 and the second self-adjusting seal 41 have a tight clearance with respect to traditional labyrinth seals. Wear of the honeycomb or line facing the lower seal 35 can contribute to the deterioration of the system performance. The low leakages and flow through the lower seal 35 improve the aerodynamics in the pre-swirl cavity 30. Better cooling is equivalent to additional power margins to maintain the same blade metal temperature.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un moteur à turbine à gaz qui comprend un premier passage formé par une paroi externe d'une chambre de combustion et une paroi interne à l'intérieur du moteur à turbine à gaz et un second passage formé par la paroi interne et un bras rotatif. Une cavité de pré-tourbillonnement (30) est située à une extrémité aval du premier passage (19). Un premier joint d'étanchéité auto-ajustable (40) est situé dans le second passage (20). Un second joint d'étanchéité auto-ajustable (41) est situé radialement vers l'extérieur de la cavité de pré-tourbillonnement (30).
PCT/US2019/012452 2018-02-27 2019-01-07 Moteur à turbine à gaz avec système de distribution d'air de refroidissement de turbine WO2019168590A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
PCT/US2018/019943 WO2019168501A1 (fr) 2018-02-27 2018-02-27 Système de distribution d'air de refroidissement de turbine
USPCT/US2018/019943 2018-02-27

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Publication Number Publication Date
WO2019168590A1 true WO2019168590A1 (fr) 2019-09-06

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PCT/US2018/019943 WO2019168501A1 (fr) 2018-02-27 2018-02-27 Système de distribution d'air de refroidissement de turbine
PCT/US2019/012452 WO2019168590A1 (fr) 2018-02-27 2019-01-07 Moteur à turbine à gaz avec système de distribution d'air de refroidissement de turbine

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Application Number Title Priority Date Filing Date
PCT/US2018/019943 WO2019168501A1 (fr) 2018-02-27 2018-02-27 Système de distribution d'air de refroidissement de turbine

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Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111441828B (zh) * 2020-03-12 2022-09-16 中国科学院工程热物理研究所 一种带预旋喷嘴和导流盘的发动机涡轮盘腔结构
CN111828108B (zh) * 2020-07-24 2023-02-21 中国科学院工程热物理研究所 一种发动机涡轮盘预旋系统用盖板盘结构

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2292868A1 (fr) * 1974-11-27 1976-06-25 Gen Electric Systeme de joints a labyrinthe pour turbine a gaz
EP0501066A1 (fr) * 1991-02-28 1992-09-02 General Electric Company Disque de moteur de turbine avec rainures et ailettes intégrales pour le pompage d'air de refroidissement
US5402636A (en) * 1993-12-06 1995-04-04 United Technologies Corporation Anti-contamination thrust balancing system for gas turbine engines
EP0785338A1 (fr) * 1996-01-18 1997-07-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Dispositif de refroidissement d'un disque de turbine
EP1785651A1 (fr) * 2005-11-15 2007-05-16 Snecma Léchette annulaire destinée à un labyrinthe d'étanchéité, et son procédé de fabrication
US8002285B2 (en) 2003-05-01 2011-08-23 Justak John F Non-contact seal for a gas turbine engine
US8172232B2 (en) 2003-05-01 2012-05-08 Advanced Technologies Group, Inc. Non-contact seal for a gas turbine engine
US20130195627A1 (en) * 2012-01-27 2013-08-01 Jorn A. Glahn Thrust balance system for gas turbine engine
US8919781B2 (en) 2003-05-01 2014-12-30 Advanced Technologies Group, Inc. Self-adjusting non-contact seal
EP3133241A1 (fr) * 2015-08-19 2017-02-22 United Technologies Corporation Ensemble de joint d'étanchéité sans contact pour équipement rotatif
EP3159480A1 (fr) * 2015-10-19 2017-04-26 United Technologies Corporation Joint d'étanchéité de rotor et commande d'équilibrage de poussée de rotor

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20080041064A1 (en) * 2006-08-17 2008-02-21 United Technologies Corporation Preswirl pollution air handling with tangential on-board injector for turbine rotor cooling
US9810079B2 (en) * 2013-03-15 2017-11-07 General Electric Company Cyclonic dirt separating turbine accelerator
US10370991B2 (en) * 2014-11-07 2019-08-06 United Technologies Corporation Gas turbine engine and seal assembly therefore
GB2536628A (en) * 2015-03-19 2016-09-28 Rolls Royce Plc HPT Integrated interstage seal and cooling air passageways

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2292868A1 (fr) * 1974-11-27 1976-06-25 Gen Electric Systeme de joints a labyrinthe pour turbine a gaz
EP0501066A1 (fr) * 1991-02-28 1992-09-02 General Electric Company Disque de moteur de turbine avec rainures et ailettes intégrales pour le pompage d'air de refroidissement
US5402636A (en) * 1993-12-06 1995-04-04 United Technologies Corporation Anti-contamination thrust balancing system for gas turbine engines
EP0785338A1 (fr) * 1996-01-18 1997-07-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Dispositif de refroidissement d'un disque de turbine
US8002285B2 (en) 2003-05-01 2011-08-23 Justak John F Non-contact seal for a gas turbine engine
US8172232B2 (en) 2003-05-01 2012-05-08 Advanced Technologies Group, Inc. Non-contact seal for a gas turbine engine
US8919781B2 (en) 2003-05-01 2014-12-30 Advanced Technologies Group, Inc. Self-adjusting non-contact seal
EP1785651A1 (fr) * 2005-11-15 2007-05-16 Snecma Léchette annulaire destinée à un labyrinthe d'étanchéité, et son procédé de fabrication
US20130195627A1 (en) * 2012-01-27 2013-08-01 Jorn A. Glahn Thrust balance system for gas turbine engine
EP3133241A1 (fr) * 2015-08-19 2017-02-22 United Technologies Corporation Ensemble de joint d'étanchéité sans contact pour équipement rotatif
EP3159480A1 (fr) * 2015-10-19 2017-04-26 United Technologies Corporation Joint d'étanchéité de rotor et commande d'équilibrage de poussée de rotor

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