US20140321981A1 - Turbine engine shutdown temperature control system - Google Patents

Turbine engine shutdown temperature control system Download PDF

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Publication number
US20140321981A1
US20140321981A1 US13/871,080 US201313871080A US2014321981A1 US 20140321981 A1 US20140321981 A1 US 20140321981A1 US 201313871080 A US201313871080 A US 201313871080A US 2014321981 A1 US2014321981 A1 US 2014321981A1
Authority
US
United States
Prior art keywords
air
turbine
cavity
exhaust
longitudinal axis
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.)
Abandoned
Application number
US13/871,080
Other languages
English (en)
Inventor
Jose L. Rodriguez
David A. Little
Jiping Zhang
Patrick M. Pilapil
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.)
Siemens AG
Original Assignee
Siemens AG
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 AG filed Critical Siemens AG
Priority to US13/871,080 priority Critical patent/US20140321981A1/en
Assigned to SIEMENS ENERGY, INC reassignment SIEMENS ENERGY, INC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: RODRIGUEZ, JOSE L., LITTLE, DAVID A., PILAPIL, Patrick M., ZHANG, JIPING
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS ENERGY, INC.
Priority to EP14723657.4A priority patent/EP2989297A1/en
Priority to RU2015145806A priority patent/RU2015145806A/ru
Priority to CN201480020972.5A priority patent/CN105121789A/zh
Priority to JP2016510702A priority patent/JP2016516941A/ja
Priority to PCT/US2014/034296 priority patent/WO2014176085A1/en
Publication of US20140321981A1 publication Critical patent/US20140321981A1/en
Abandoned 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • 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
    • F01D19/00—Starting of machines or engines; Regulating, controlling, or safety means in connection therewith
    • F01D19/02—Starting of machines or engines; Regulating, controlling, or safety means in connection therewith dependent on temperature of component parts, e.g. of turbine-casing
    • 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
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/26—Double casings; Measures against temperature strain in casings
    • 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
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F02C7/18—Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air

Definitions

  • This invention is directed generally to turbine engines, and more particularly to systems enabling warm startups of the gas turbine engines without risk of compressor and turbine blade interference with radially outward sealing surfaces.
  • gas turbine engines typically include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power.
  • Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit.
  • Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures.
  • the engines take a long time to cool down after shutdown.
  • Many of the components cool at different rates and as a result, interferences develop between various components.
  • the clearance between turbine blade tips and blade rings positioned immediately radially outward of the turbine blades is such a configuration in which an interference often develops.
  • the turbine vane carriers with blade rings typically cool faster than the turbine rotor assembly including the turbine blades.
  • the turbine vane carriers reduce in diameter more than the turbine rotor assembly.
  • the turbine vane carriers reduce in diameter more than the turbine rotor assembly.
  • the turbine engine shutdown temperature control system may include a turbine engine component positioned within a turbine case such that a cavity is positioned therebetween and at least one air amplifier.
  • the air amplifier may include a hollow chamber, wherein the air amplifier may extend into the cavity and may have a longitudinal axis that is nonparallel with a longitudinal axis of the turbine case. An exhaust outlet of the air amplifier may be directed to exhaust air in a direction that is nonparallel with the longitudinal axis of the turbine case.
  • the air amplifier may be in fluid communication with an air supply source enabling the air amplifier to exhaust air into the cavity.
  • the air amplifier may be positioned in a turbine case that may be, but is not limited to being, a turbine exhaust casing forward outer diameter cavity or a turbine combustion casing midframe cavity.
  • the turbine engine component positioned concentrically within the turbine case may be, but is not limited to being, a turbine vane carrier.
  • the air amplifier may be offset in the cavity such that the air amplifier is positioned radially outward from the longitudinal axis of the turbine case.
  • the air exhausted from the exhaust outlet of the air amplifier is able to create a circumferential flow within the cavity.
  • the exhaust outlet of the air amplifier may be positioned no further than a distance from an inner surface of an outer wall equating to 20 percent of a radially extending distance from the inner surface of the turbine case forming the outer wall of the cavity to an inner surface forming an inner wall of the cavity.
  • the exhaust outlet of the air amplifier when viewed axially downstream, may be offset circumferentially from top dead center, bottom dead center, left side center and right side center.
  • the turbine engine shutdown temperature control system may include first and second air amplifiers.
  • the turbine engine shutdown temperature control system may include a first air amplifier extending into the cavity and having a longitudinal axis that is nonparallel with the longitudinal axis of the turbine case.
  • An exhaust outlet of the first air amplifier may be directed to exhaust air in a direction that is nonparallel with the longitudinal axis of the turbine case.
  • the second air amplifier may extend into the cavity and having a longitudinal axis that is nonparallel with the longitudinal axis of the turbine case.
  • An exhaust outlet of the second air amplifier may be directed to exhaust air in a direction that is nonparallel with the longitudinal axis of the turbine case and in a same circumferential direction as air exhausted from the exhaust outlet of the first air amplifier.
  • the first and second air amplifiers may be positioned on opposite sides of the cavity from each other to entrain a circumferential airflow and maintain a consistent airflow temperature throughout the cavity.
  • An advantage of this invention is that the turbine engine shutdown temperature control system creates consistent air temperature within cavities surrounding compressor and turbine blade assemblies to eliminate turbine and compressor blade tip rub during warm restarts of gas turbine engines.
  • the turbine engine shutdown temperature control system creates a circumferential airflow within the cavity formed by the turbine case by entraining the air already within the turbine case in the cavity, thereby creating an air mass within the cavity having a generally consistent temperature without having to inject large quantities of air into cavity.
  • Still another advantage of this invention is that the turbine engine shutdown temperature control system reduces the temperature gradient between the top dead center and other aspects of the turbine case.
  • Another advantage of this invention is that the turbine engine shutdown temperature control system reduces the temperature gradient between the top dead center and other aspects of the blade ring to eliminate turbine airfoil rubbing.
  • Yet another advantage of this invention is that due to a large amplification ratio of the turbine engine shutdown temperature control system, there is no need for large conduits, such as four inches to six inches. Rather, an air supply source may be coupled to one or more air amplifiers with 3 ⁇ 8 inch diameter tubing.
  • Another advantage of this invention is that only a very small amount of compressed air is needed to entrain air within the cavity in the turbine case.
  • Still another advantage of this invention is that the air exhausted by one or more air amplifiers has a temperature nearly equal to a temperature of the air within the cavity, thereby eliminating the possibility of thermal shock to the system.
  • Another advantage of this invention is that the turbine engine shutdown temperature control system operates during turning gear operation and thus, there is no impact to normal gas turbine engine operation.
  • FIG. 1 is a cross-sectional side view of a gas turbine engine.
  • FIG. 3 is a schematic diagram of a gas turbine engine including the turbine engine shutdown temperature control system taken at section line 3 - 3 in FIG. 2 .
  • FIG. 4 is a perspective view of the turbine engine shutdown temperature control system positioned in a turbine exhaust casing.
  • FIG. 5 is a detailed view of an air amplifier installed in a turbine exhaust casing.
  • FIG. 6 is another detailed view of an air amplifier installed in a turbine exhaust casing.
  • FIG. 7 is a partial perspective view of the turbine engine shutdown temperature control system positioned in a turbine combustion casing.
  • FIG. 8 is a detailed view of the turbine engine shutdown temperature control system positioned in the turbine combustion casing.
  • FIG. 9 is a partial perspective view of the turbine engine shutdown temperature control system positioned in a turbine vane carrier.
  • FIG. 10 is a partial cross-sectional view of the turbine engine shutdown temperature control system positioned a horizontal joint of the turbine vane carrier.
  • FIG. 11 is a top view of a portion of a turbine case in which an air amplifier is mounted.
  • this invention is directed to a turbine engine shutdown temperature control system 10 configured to foster consistent air temperature within cavities 12 surrounding compressor and turbine blade assemblies 14 , 16 to eliminate turbine and compressor blade 18 , 20 tip rub during warm restarts of gas turbine engines 22 .
  • the turbine engine shutdown temperature control system 10 may include one or more air amplifiers 24 positioned in a turbine engine component.
  • An exhaust outlet 26 of the air amplifier 24 may extend into a cavity 12 created by a turbine case 34 and may be configured to exhaust air in a generally circumferential direction to entrain air within the cavity 12 to flow circumferentially to establish a consistent air temperature within the cavity 12 thereby preventing uneven cooling of turbine engine components after shutdown and prevent damage to turbine components during a warm restart.
  • the turbine engine shutdown temperature control system 10 may include one or more air amplifiers 24 configured to exhaust air into the cavity 12 .
  • the air amplifier 24 may include one or more hollow chambers 28 .
  • the air amplifier 24 may extend into the cavity 12 and may have a longitudinal axis 30 that is nonparallel with a longitudinal axis 32 of the turbine case 34 .
  • the air amplifier 24 may be positioned such that an exhaust outlet 26 of the air amplifier 24 may be directed to exhaust air in a direction that is nonparallel with the longitudinal axis 32 of the turbine case 34 .
  • the air amplifier 24 may be in fluid communication with an air supply source 36 enabling the air amplifier 24 to exhaust air into the air amplifier 24 .
  • the turbine engine shutdown temperature control system 10 may include a turbine engine component 38 positioned concentrically within a turbine case 34 such that a cavity 12 is positioned therebetween.
  • the cavity 12 may be a cylindrically shaped cavity 12 defined outwardly by the turbine case 34 and inwardly by the turbine component 38 .
  • the turbine component 38 may be, but is not limited to being, a compressor blade assembly 14 , a turbine blade assembly 16 and combustor 40 .
  • the turbine case 34 may be a turbine exhaust casing 42 forming a forward outer diameter cavity 44 .
  • the turbine case 34 may be a turbine combustion casing 46 forming a turbine combustion casing midframe cavity 48 .
  • the cavity may also be defined inwardly by outer surfaces of a turbine component 38 .
  • the turbine component 38 may be a turbine vane carrier 50 positioned concentrically within a turbine case 34 .
  • the air amplifier 24 positioned within the cavity 12 such that the air amplifier 24 is offset within the cavity 12 .
  • the air amplifier 24 may be offset in the cavity 12 such that the air amplifier 24 may be positioned radially outward from the longitudinal axis 32 of the turbine case 34 .
  • the exhaust outlet 26 of the air amplifier 24 when viewed axially downstream, may be offset circumferentially from top dead center 52 , bottom dead center 54 , left side center 56 and right side center 58 .
  • the exhaust outlet 26 of the air amplifier 24 when viewed from above the turbine engine 22 , may be directed orthogonal to the longitudinal axis 32 of the engine 22 and the turbine case 34 .
  • the exhaust outlet 26 of the air amplifier 24 when viewed from above the turbine engine 22 , may be directed nonorthogonal to the longitudinal axis 32 of the engine 22 and the turbine case 34 .
  • air exhausted out of the exhaust outlet 26 may flow torodially within the cavity 12 such that it flows nonorthogonally relative to the longitudinal axis 32 of the engine 22 and the turbine case 34 .
  • the exhaust outlet 26 of the air amplifier 24 may be positioned to entrain air within the cavity 12 while maintaining no more than a limited profile to limit disruption of the air flow in the cavity 12 .
  • the exhaust outlet 26 of the air amplifier 24 may be positioned no further than a distance 60 from an inner surface 62 of an outer wall 64 equating to 20 percent of a radially extending distance 60 from the inner surface 62 of the turbine case 34 forming the outer wall 64 of the cavity 12 to an inner surface 66 forming an inner wall 68 of the cavity 12 .
  • the air amplifier 24 may be formed from a support engaging extension 70 and an exhaust region 72 housing the exhaust outlet 26 .
  • the support engaging extension 70 may extend through the turbine case 34 , and the exhaust region 72 may be positioned in the cavity 12 .
  • the exhaust region 72 may be nonparallel and nonorthogonal with the support engaging extension 70 .
  • the exhaust region 72 may be orthogonal with the support engaging extension 70 .
  • the exhaust region 72 may be aligned tangentially with the inner surface 62 of the turbine case 34 forming the cavity 12 . In such a position, the air exhausted from the exhaust outlet 26 causes minimal disruption with the air flow within the cavity 12 and continues to keep the air flowing in a circumferential direction within the cavity 12 .
  • One or more air amplifiers 24 may be positioned within the turbine case 34 in varying configurations to entrain the air within the cavity 12 via air exhausted from the air amplifiers 24 .
  • the turbine engine shutdown temperature control system 10 may be formed from a first air amplifier 74 extending into the cavity 12 and having a longitudinal axis 30 that is nonparallel with the longitudinal axis 32 of the turbine case 34 .
  • the exhaust outlet 26 of the first air amplifier 74 may be directed to exhaust air in a direction that is nonparallel with the longitudinal axis 32 of the turbine case 34 .
  • the turbine engine shutdown temperature control system 10 may also include a second air amplifier 76 extending into the cavity 12 and having a longitudinal axis 30 that is nonparallel with the longitudinal axis 32 of the turbine case 34 .
  • the exhaust outlet 26 of the second air amplifier 76 may be directed to exhaust air in a direction that is nonparallel with the longitudinal axis 32 of the turbine case 34 and in a same circumferential direction as air exhausted from the exhaust outlet 26 of the first air amplifier 74 .
  • the first and second air amplifiers 74 , 76 may be positioned on opposite sides of the cavity 12 from each other to entrain a circumferential airflow and maintain a consistent airflow temperature throughout the cavity 12 .
  • the exhaust regions 72 of the first and second air amplifiers 74 , 76 may be aligned tangentially with an inner surface 62 of the turbine case 34 forming the cavity 12 .
  • the exhaust outlets 26 of the first and second air amplifiers 74 , 76 may be positioned no further than a distance from the inner surface 62 of the outer wall 64 equating to 20 percent of a radially extending distance 20 from the inner surface 62 of the turbine engine component forming the outer wall 62 of the cavity 12 to an inner surface 66 forming an inner wall 68 of the cavity 12 .
  • the air amplifiers 24 may be positioned in close proximity to horizontal extending joints 78 of the turbine case 34 .
  • One or more first air amplifiers 74 may be positioned proximate to a first horizontal extending joint 78 on a first side of the turbine case 34
  • one or more second air amplifiers 76 may be positioned proximate to a second horizontal extending joint 78 positioned on a second side of the turbine case 34 whereby the second side is on an opposite side of the turbine case 34 from the first side.
  • the turbine engine shutdown temperature control system 10 may be used to foster consistent air temperature within cavities 12 surrounding compressor and turbine blade assemblies 14 , 16 to eliminate turbine and compressor blade 18 , 20 tip rub during warm restarts of gas turbine engines 22 .
  • the turbine engine shutdown temperature control system 10 creates a consistent air temperature within the air cavity 12 by exhausting air from the exhaust outlet 26 of the air amplifiers 24 into the cavity 12 .
  • the exhaust air is exhausted in a circumferential direction to create a circumferential flow of air within the cavity 12 and thereby prevent the hottest gases within the cavity 12 from collecting at the top dead center 52 and causing unequal cooling and unequal thermal contraction while the turbine engine 22 cools after shutdown.
  • the turbine engine shutdown temperature control system 10 operates during turning gear operation and thus, there is no impact to normal gas turbine engine operation.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Exhaust Silencers (AREA)
  • Control Of Turbines (AREA)
US13/871,080 2013-04-26 2013-04-26 Turbine engine shutdown temperature control system Abandoned US20140321981A1 (en)

Priority Applications (6)

Application Number Priority Date Filing Date Title
US13/871,080 US20140321981A1 (en) 2013-04-26 2013-04-26 Turbine engine shutdown temperature control system
EP14723657.4A EP2989297A1 (en) 2013-04-26 2014-04-16 Turbine engine shutdown temperature control system
RU2015145806A RU2015145806A (ru) 2013-04-26 2014-04-16 Система управления температурой останова турбинного двигателя
CN201480020972.5A CN105121789A (zh) 2013-04-26 2014-04-16 涡轮发动机停机温度控制系统
JP2016510702A JP2016516941A (ja) 2013-04-26 2014-04-16 タービンエンジンシャットダウン温度制御システム
PCT/US2014/034296 WO2014176085A1 (en) 2013-04-26 2014-04-16 Turbine engine shutdown temperature control system

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/871,080 US20140321981A1 (en) 2013-04-26 2013-04-26 Turbine engine shutdown temperature control system

Publications (1)

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US20140321981A1 true US20140321981A1 (en) 2014-10-30

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Application Number Title Priority Date Filing Date
US13/871,080 Abandoned US20140321981A1 (en) 2013-04-26 2013-04-26 Turbine engine shutdown temperature control system

Country Status (6)

Country Link
US (1) US20140321981A1 (zh)
EP (1) EP2989297A1 (zh)
JP (1) JP2016516941A (zh)
CN (1) CN105121789A (zh)
RU (1) RU2015145806A (zh)
WO (1) WO2014176085A1 (zh)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP3862542A1 (en) * 2020-02-05 2021-08-11 Raytheon Technologies Corporation Cooling system for power cables in a gas turbine engine
US11274604B2 (en) * 2016-02-12 2022-03-15 Raytheon Technologies Corporation Bowed rotor start mitigation in a gas turbine engine using aircraft-derived parameters
US11384690B2 (en) 2015-12-30 2022-07-12 General Electric Company System and method of reducing post-shutdown engine temperatures
US11578657B2 (en) 2020-10-27 2023-02-14 Raytheon Technologies Corporation Power cable cooling system in a gas turbine engine
US11585291B2 (en) 2020-09-11 2023-02-21 Raytheon Technologies Corporation Tail cone ejector for power cable cooling system in a gas turbine engine
US11668206B1 (en) * 2022-03-09 2023-06-06 General Electric Company Temperature gradient control system for a compressor casing

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US11035251B2 (en) * 2019-09-26 2021-06-15 General Electric Company Stator temperature control system for a gas turbine engine
CN114198153B (zh) * 2020-09-17 2024-05-03 中国航发商用航空发动机有限责任公司 涡轮叶片冷却系统及航空发动机

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Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11384690B2 (en) 2015-12-30 2022-07-12 General Electric Company System and method of reducing post-shutdown engine temperatures
US11274604B2 (en) * 2016-02-12 2022-03-15 Raytheon Technologies Corporation Bowed rotor start mitigation in a gas turbine engine using aircraft-derived parameters
EP3862542A1 (en) * 2020-02-05 2021-08-11 Raytheon Technologies Corporation Cooling system for power cables in a gas turbine engine
US11719113B2 (en) 2020-02-05 2023-08-08 Raytheon Technologies Corporation Cooling system for power cables in a gas turbine engine
EP4368817A3 (en) * 2020-02-05 2024-06-05 RTX Corporation Cooling system for power cables in a gas turbine engine
US12044149B2 (en) 2020-02-05 2024-07-23 Rtx Corporation Cooling system for power cables in a gas turbine engine
US11585291B2 (en) 2020-09-11 2023-02-21 Raytheon Technologies Corporation Tail cone ejector for power cable cooling system in a gas turbine engine
US11578657B2 (en) 2020-10-27 2023-02-14 Raytheon Technologies Corporation Power cable cooling system in a gas turbine engine
US11668206B1 (en) * 2022-03-09 2023-06-06 General Electric Company Temperature gradient control system for a compressor casing

Also Published As

Publication number Publication date
JP2016516941A (ja) 2016-06-09
EP2989297A1 (en) 2016-03-02
WO2014176085A1 (en) 2014-10-30
RU2015145806A (ru) 2017-06-02
CN105121789A (zh) 2015-12-02

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AS Assignment

Owner name: SIEMENS ENERGY, INC, FLORIDA

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:RODRIGUEZ, JOSE L.;LITTLE, DAVID A.;ZHANG, JIPING;AND OTHERS;SIGNING DATES FROM 20130307 TO 20130419;REEL/FRAME:030293/0546

AS Assignment

Owner name: SIEMENS AKTIENGESELLSCHAFT, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:SIEMENS ENERGY, INC.;REEL/FRAME:032026/0784

Effective date: 20130904

STCB Information on status: application discontinuation

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