EP3106613A1 - Système de refroidissement pour moteur à turbine à gaz - Google Patents

Système de refroidissement pour moteur à turbine à gaz Download PDF

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Publication number
EP3106613A1
EP3106613A1 EP15001689.7A EP15001689A EP3106613A1 EP 3106613 A1 EP3106613 A1 EP 3106613A1 EP 15001689 A EP15001689 A EP 15001689A EP 3106613 A1 EP3106613 A1 EP 3106613A1
Authority
EP
European Patent Office
Prior art keywords
turbine
cooling
plenum
cooling air
rotor compartment
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Withdrawn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
EP15001689.7A
Other languages
German (de)
English (en)
Inventor
Gabriel L. Suciu
Brian D. Merry
James D. Hill
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.)
Raytheon Technologies 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
Application filed by United Technologies Corp filed Critical United Technologies Corp
Publication of EP3106613A1 publication Critical patent/EP3106613A1/fr
Withdrawn legal-status Critical Current

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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
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • 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
    • F01D5/084—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades the fluid circulating at the periphery of a multistage rotor, e.g. of drum type
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes

Definitions

  • the present disclosure relates to turbine engines, and more particularly to turbine engines having improved high pressure turbine cooling.
  • a gas turbine engine typically includes a multi-stage axial compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and directed to the combustor where it is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. To control the heat transfer induced by the hot combustion gases entering the turbine, typically cooling air is channeled through turbine cooling circuits and is used to cool various turbine components.
  • Maintaining sufficient cooling air within the gas turbine engine is critical to proper engine performance and component longevity.
  • the flow of cooling air across the turbine rotor and through the interior of the blades removes heat so as to prevent excessive reduction of the mechanical strength properties of the turbine blades and turbine rotor.
  • Typical cooling methods include directing cooling air from a variety of sources having different pressures and temperatures.
  • a cooling system for gas turbine engines includes a turbine rotor compartment defining a cooling air plenum.
  • a plurality of turbine discs are rotatably housed within the rotor compartment.
  • a cooling air inlet is in fluid communication with the plenum.
  • Each turbine disc includes a cooling outlet in fluid communication with the plenum for cooling the rotor compartment.
  • the cooling air inlet can be in fluid communication with a cooling air conduit aligned to deliver fluid output from a high pressure compressor.
  • the rotor compartment can include a plurality of seals proximate a hub of each turbine disc to fluidly seal the plenum.
  • the rotor compartment can have first and second turbine discs. Each turbine disc can be sealingly engaged to the rotor compartment.
  • a segmented seal can be disposed between the first and second turbine discs to isolate the plenum from a gas path outboard of the rotor compartment.
  • Each turbine disc can include a plurality of turbine blades mounted thereabout, wherein the turbine blades provide a cooling air outlet of the plenum.
  • Each turbine disc can also have at least one passage through the hub thereof to allow cooling air from the inlet to circulate through the plenum.
  • each turbine blade can have an axial passage therethrough to place the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs. The passage can be disposed proximate of the turbine blades of each respective turbine disc.
  • a method for providing cooling air to a rotor stage includes delivering fluid output from a high pressure compressor through a cooling air inlet into a cooling air plenum defined by a rotor compartment.
  • the method can further include sealing the cooling air plenum with a plurality of seals disposed near a hub of each turbine disc.
  • a segmented seal can be disposed between first and second turbine discs to isolate the cooling air plenum from a gas path outboard of the rotor compartment.
  • the method can further include cooling the plenum by flowing the fluid output through at least one passage of a hub of each turbine disc.
  • Each turbine disc may include a plurality of turbine blades, each turbine disc having at least one passage proximate the turbine blades to place the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs.
  • FIG. 1 a partial view of an exemplary embodiment of a cooling system for gas turbine engines in accordance with the disclosure is shown in Fig. 1 and is designated generally by reference character 100.
  • FIG. 2 Other embodiments of the incremental cooling system in accordance with the disclosure, or aspects thereof, are provided in Fig. 2 , as will be described.
  • a typical prior art method of cooling a gas turbine engine includes using cooling air to cool first and second rotors from a variety of sources.
  • one cooling line is typically provided from cooled air external to the system whereas another cooling line is typically for cooled compressor discharge air.
  • Each of the different cooling lines being directed to cool different parts of the rotors.
  • turbine discs in turbine engines of the prior art do not allow for cooled air to reach between or around each disc.
  • the sources of cooling air used in the typical configurations are at different levels of pressure and temperature. This variation leads to compromises and increased stress on the turbine components due to the thermal gradient.
  • Fig. 1 illustrates an exemplary embodiment of a turbine engine 100 of the present disclosure where first and second turbine discs 130, 132, respectively, of a rotor compartment 106 are cooled entirely from the same source of cooling air S1.
  • Uncombusted high pressure air typically referred to as T3 air, for example, is passed through a tangential on-board injector ("TOB1") nozzle.
  • TOB1 tangential on-board injector
  • the TOB1 nozzle reduces the relative total temperature of the T3 air.
  • the reduced temperature T3 air is directed to the rotor compartment 106 to cool the entire rotor system.
  • the flow paths within the rotor compartment and a plurality of seals around the rotor compartment are arranged to provide cooling air to the entire rotor compartment from the single source, without the need for additional sources of cooling air. This ensures that the entire rotor system is cooled from a single source with one temperature and pressure thereby improving the longevity of motor life compared to a traditional system.
  • the rotor compartment 106 defines a cooling air plenum 120 in fluid communication with a cooling air inlet 122. T3 cooling air is directed to the cooling air inlet via a cooling air conduit 124.
  • a plurality of turbine discs 130, 132 are rotatably housed within the rotor compartment 106. First and second turbine discs 130,132 are shown in Fig. 2 , however, the rotor compartment may contain any suitable number of turbine discs.
  • the cooling air plenum 120 is fluidly sealed by a plurality of seals 140, 142, 144 disposed around the turbine discs 130, 132 to cool the entire rotor compartment 106.
  • seals 140 and 142 are located near hubs 136B, 136A of the turbine discs 130, 132.
  • a segmented seal 144 is disposed between the first and second turbine discs 130, 132. The segmented seal 144 isolates the cooling air plenum 120 from a gas path 126 outboard of the rotor compartment 106. This allows the cooling air to flow inboard of the segmented seal 144 around the first and second turbine discs 130, 132.
  • each turbine disc 130, 132 of the present disclosure includes passages in fluid communication with the plenum for cooling the rotor compartment. More specifically, a plurality of passages 150, 152, 154, 156 are disposed within turbine discs 130, 132 to allow the cooling air to flow therethrough and reach the entire plenum. At least one passage 150, 152 is provided through hubs 136B, 136A of each of the first and second turbine discs 130, 132. This allows cooling air to flow through the hub and between the first and second turbine discs 130, 132 as indicated by the flow arrows in Fig. 1 . At least one axial passage 154, 156 is also provided within each turbine disc 130, 132.
  • Each turbine disc includes a plurality of turbine blades 137, 139 mounted thereabout.
  • the turbine blades 137, 139 provide cooling air outlets 138 to the plenum 120.
  • the seals and passages are positioned to fluidly seal the rotor compartment while allowing the cooling air to circulate within the entire plenum to cool components of the turbine engine. Additional seals and/or passages may be included. Further, the positioning of the seals and/or passages may be adjusted, changed or modified without departing from the scope of the present disclosure.
  • a method for providing cooling air to a rotor stage comprises, at box 202, delivering fluid output from a high pressure compressor through a cooling air inlet, e.g., air inlet 122, into a cooling air plenum, e.g., plenum 120, defined by a rotor compartment, e.g., rotor compartment 106.
  • the method further includes sealing the plenum with a plurality of seals, e.g., seals 140, 142, 144, at box 204. At least one seal disposed near a hub of each turbine disc and a segmented seal between first and second turbine discs.
  • the segmented seal isolating the plenum from a gas path outboard the rotor compartment, as shown in box 206.
  • at least one passage e.g., passages 150, 152, 154, 156, is positioned through the hub of each turbine disc.
  • placing the cooling air inlet in fluid communication with a portion of the plenum between the turbine discs is constructed by providing at least one passage proximate turbine blades, e.g., blades 137, 139.
  • any other suitable cooling fluid from any other suitable source can be used without departing from the scope of this disclosure.
  • any other suitable components e.g., compressor or combustor components, can also be cooled with the methods and delivery described herein.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP15001689.7A 2015-06-06 2015-06-08 Système de refroidissement pour moteur à turbine à gaz Withdrawn EP3106613A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US201562008619P 2015-06-06 2015-06-06

Publications (1)

Publication Number Publication Date
EP3106613A1 true EP3106613A1 (fr) 2016-12-21

Family

ID=53365703

Family Applications (1)

Application Number Title Priority Date Filing Date
EP15001689.7A Withdrawn EP3106613A1 (fr) 2015-06-06 2015-06-08 Système de refroidissement pour moteur à turbine à gaz

Country Status (1)

Country Link
EP (1) EP3106613A1 (fr)

Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
GB2137283A (en) * 1983-03-30 1984-10-03 United Technologies Corp Clearance control in turbine seals
EP0313826A1 (fr) * 1987-10-30 1989-05-03 BBC Brown Boveri AG Turbine à gaz avec flux axial
US5232339A (en) * 1992-01-28 1993-08-03 General Electric Company Finned structural disk spacer arm
US5795130A (en) * 1995-11-24 1998-08-18 Mitsubishi Jukogyo Kabushiki Kaisha Heat recovery type gas turbine rotor
EP1413711A1 (fr) * 2002-10-21 2004-04-28 Siemens Aktiengesellschaft Procédé et dispositif de refroidissement des aubes des turbines à gaz
WO2004113684A1 (fr) * 2003-06-16 2004-12-29 Siemens Aktiengesellschaft Turbomachine, en particulier turbine a gaz
US20110085905A1 (en) * 2009-10-14 2011-04-14 General Electric Company Turbomachine rotor cooling
EP2458147A2 (fr) * 2010-11-29 2012-05-30 Alstom Technology Ltd Turbine à gaz à flux axial
US20120148405A1 (en) * 2010-12-13 2012-06-14 General Electric Company Cooling circuit for a drum rotor
EP2586968A2 (fr) * 2011-10-28 2013-05-01 United Technologies Corporation Agencement d'écoulement secondaire pour rotor rainuré

Patent Citations (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2791091A (en) * 1950-05-15 1957-05-07 Gen Motors Corp Power plant cooling and thrust balancing systems
GB2137283A (en) * 1983-03-30 1984-10-03 United Technologies Corp Clearance control in turbine seals
EP0313826A1 (fr) * 1987-10-30 1989-05-03 BBC Brown Boveri AG Turbine à gaz avec flux axial
US5232339A (en) * 1992-01-28 1993-08-03 General Electric Company Finned structural disk spacer arm
US5795130A (en) * 1995-11-24 1998-08-18 Mitsubishi Jukogyo Kabushiki Kaisha Heat recovery type gas turbine rotor
EP1413711A1 (fr) * 2002-10-21 2004-04-28 Siemens Aktiengesellschaft Procédé et dispositif de refroidissement des aubes des turbines à gaz
WO2004113684A1 (fr) * 2003-06-16 2004-12-29 Siemens Aktiengesellschaft Turbomachine, en particulier turbine a gaz
US20110085905A1 (en) * 2009-10-14 2011-04-14 General Electric Company Turbomachine rotor cooling
EP2458147A2 (fr) * 2010-11-29 2012-05-30 Alstom Technology Ltd Turbine à gaz à flux axial
US20120148405A1 (en) * 2010-12-13 2012-06-14 General Electric Company Cooling circuit for a drum rotor
EP2586968A2 (fr) * 2011-10-28 2013-05-01 United Technologies Corporation Agencement d'écoulement secondaire pour rotor rainuré

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