EP2378071A1 - Turbineneinheit mit Kühlanordnung und Kühlverfahren - Google Patents

Turbineneinheit mit Kühlanordnung und Kühlverfahren Download PDF

Info

Publication number
EP2378071A1
EP2378071A1 EP10004074A EP10004074A EP2378071A1 EP 2378071 A1 EP2378071 A1 EP 2378071A1 EP 10004074 A EP10004074 A EP 10004074A EP 10004074 A EP10004074 A EP 10004074A EP 2378071 A1 EP2378071 A1 EP 2378071A1
Authority
EP
European Patent Office
Prior art keywords
sidewall
segment
perforations
cooling air
face
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Withdrawn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
EP10004074A
Other languages
English (en)
French (fr)
Inventor
Sergey Shukin
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 EP10004074A priority Critical patent/EP2378071A1/de
Publication of EP2378071A1 publication Critical patent/EP2378071A1/de
Withdrawn legal-status Critical Current

Links

Images

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
    • 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
    • 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
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving blades
    • F05D2240/81—Cooled platforms
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

Definitions

  • the present invention relates turbomachines such as gas turbines.
  • the present invention relates to cooling of blade or vane endwalls (platforms) and heat shields in turbomachines.
  • Modern gas turbines operate under extremely high gas temperature conditions. This requires the use of heavy cooling of the airfoils and the endwalls (also referred to as platforms) of turbine blades and vanes to ensure sufficient lifetime of these blades and vanes.
  • the efficiency of a gas turbine can be increased by optimizing certain parameters of the operation of the turbine.
  • the most relevant parameters that affect the efficiency are temperature and pressure of the gas medium which forces the rotation of the turbine rotor, referred to herein as drive gas.
  • the normal operating temperature of the drive gas nowadays, especially at the turbine inlet region is already significantly higher than the admissible material temperatures of the turbine components exposed to this fluid. Material temperatures that are too high lead to drop in strength of such heat exposed components. Temperature exceeding said limits cause melting and/or formation of cracks in the component, which may eventually cause local or complete destruction of the component. In order that the component temperatures do not exceed the admissible material temperatures, turbine components exposed to high temperatures are therefore cooled by a cooling medium.
  • a turbine assembly generally includes a plurality of stationary vanes and rotating blades.
  • Each blade or vane includes an airfoil portion that extends into the flow path of the drive gas flowing axially through the turbine.
  • the base of the airfoil is arranged on a platform or an endwall.
  • platforms are set side by side in an annular manner.
  • stationary vanes such platforms or endwalls are also arranged to form an annular shroud at the tip of the blading.
  • structures commonly referred to as heat shields are often arranged annularly above the tip of the rotor blades to protect the turbine stator from high temperature of the turbine fluid.
  • Another known method for endwall cooling involves providing longitudinal holes through the endwall at the sides to allow cooling air to penetrate these holes, to effect a convective cooling.
  • An example of such a technique is disclosed in the document US6309175B1 .
  • this method involves a risk of clogging of these holes by dust and exposure of the material around these holes to thermo-mechanical fatigue. Additionally, the drilling of such longitudinal holes also presents manufacturing complexities due to the inherent geometry of these components.
  • the object of the present invention is to provide an improved method for cooling blade or vane endwalls and heat shields that addresses the above mentioned problems posed by the existing state of the art.
  • the present invention provides an improved cooling of blade and vane endwalls or heat shields that are circumferentially arranged side by side with a separating gap between the sidewalls any two adjacent segments of these endwalls or heat shields.
  • the adjacent segments are referred to herein as first and second segments.
  • the underlying idea of the present invention is to "close" the separating gap between the sidewalls by feeding the separating gap with cooling air, which, in turn, is used for cooling the sidewalls of adjacent endwall or heat shield segment. This is done by providing a perforation through at least one of the sidewalls, namely the first sidewall. The perforation conducts cooling air and ejects the same into the separating gap to impinge on the second sidewall.
  • One advantage of the present invention is that the parasitic leakage of the drive gas through the separating gap is eliminated or minimized due to the "closing" of this gap by the ejected cooling air. This improves the turbine stage efficiency and overall gas turbine efficiency.
  • a second advantage of the present invention is that the ejected cooling air provides impingement cooling of the second sidewall facing the first sidewall across the separating gap, thereby increasing the lifetime of the neighboring second segment.
  • a fourth advantage of the present invention is that the two-fold convective and impingement cooling provides a reduction in the amount of cooling air used for sidewall cooling, thus contributing to an increase in turbine stage and overall efficiency.
  • multiple such perforations are provided on both, the first and the second sidewall. These perforations on the first and second sidewalls have a relative arrangement, such that the cooling air ejected through the perforations on each sidewall impinges on spaces in between the perforations on the opposite sidewall.
  • FIG 1 an exemplary turbine assembly 1 in a high pressure turbine stage in a gas turbine engine.
  • the assembly 1 is understood to be generally symmetrical in cross-sectional view about a longitudinal machine axis 2.
  • the turbine assembly 1 includes a set of stationary guide vanes 3, one of which is shown in the cross-sectional view to the left of FIG 1 .
  • the turbine assembly 1 further includes a set of rotating blades 4, one of which is shown in the cross-sectional view to the right of FIG 1 .
  • the set of guide vanes 3 and the set of blades 4 are each mounted in annular formation around the machine axis 2 with each guide vane 3 and each blade 4 extending radially outwardly from the axis 2.
  • Each guide vane 3 includes an airfoil 3a which extends radially into the axial flow path of the drive gas between an outer endwall 5 and an inner endwall 6. Multiple segments of outer 5 and inner 6 endwalls are arranged side by side along a circumferential direction to effect an annular formation of outer 5 and inner 6endwalls endwalls respectively.
  • Each blade 4 includes an airfoil 4a which extends radially into the axial flow path of the drive gas.
  • a blade platform 8 extends circumferentially from the blade 4 at the base of the airfoil 4a.
  • One or more blades 4 may be arranged on a single platform 8. Multiple such segments of endwalls are arranged side by side along a circumferential direction to effect an annular formation of endwalls around the axis 2.
  • segments 8 referred to as heat shields are annularly mounted above the tip of the blades 4.
  • the heat shield segments 7, similar to the endwall 5,6,8 segments are also arranged side by side along a circumferential direction.
  • one or more perforations 14a is provided through the first sidewall 9a for conducting cooling air, such that the cooling air is ejected from the perforations 14a into the gap 13 to impinge on the neighboring second sidewall 9b.
  • This provides impingement cooling of the second sidewall 9b and convective cooling of the first sidewall 9a thanks to the perforations penetrating through the material of the first sidewall 9a.
  • the two-fold convective and impingement cooling advantageously provides a reduction in the amount of cooling air used for sidewall cooling, thus contributing to an increase in turbine stage and overall efficiency.
  • one or more perforations 14b are provided through the second sidewall 9b to conduct cooling air to likewise provide impingement cooling of the neighboring first sidewall 9a as well as convective cooling of the second sidewall 9b.
  • the segments 5a and 5b each respectively includes a first face 15a and 15b exposed directly to the axial flow path of the drive gas, and a second face 19a and 19b radially opposite to the respective first surfaces 15a and 15b.
  • the second faces 19a and 19b define respective cavities 23a and 23b wherein cooling air is supplied, for example, from the last stages of a compressor of the gas turbine engine.
  • the one or more perforations 14a and 14b through the sidewalls 9a and 9b are configured such as to conduct the cooling air from the respective cavities 23a and 23b.
  • a similar cooling arrangement may be advantageously used for cooling the guide vane inner endwalls 6.
  • two adjacent segments of the inner endwall 6 are denoted as a first segment 6a and a second segment 6b.
  • the segments 6a and 6b are arranged side by side with a separating gap 13 between them.
  • the first segment 6a has a first tangential sidewall 10a that faces a second tangential sidewall 10b of the second segment 6b, such that the gap 13 extends between these sidewalls 10a and 10b.
  • one or more perforations 14a and 14b are provided through the first sidewall 10a and second sidewall 10b for conducting cooling air therethrough, and ejecting the cooling air into the gap 13 to impinge on the respective neighboring sidewall 10b and 10a, for convective cooling of the respective sidewall 10a and 10b and impingement cooling of the neighboring sidewall 10b and 10a.
  • the one or more perforations 14a and 14b conduct cooling air from a respective cavity 23a and 23b defined on faces 20a and 20b of the segments 6a and 6b that are radially opposite to faces 16a and 16b that are directly exposed to the drive gas.
  • FIG 3 is a cross-sectional view along the section line B-B through the guide vanes 3 in FIG 1 .
  • FIG 3 is a cross-sectional view along the section line B-B through the guide vanes 3 in FIG 1 .
  • two adjacent segments of the heat shield 7 are denoted as a first segment 7a and a second segment 7b.
  • two adjacent segments of the blade platform or endwall 8 are denoted as a first segment 8a and a second segment 8b.
  • the respective first segments 7a,8a and second segments 7b,8b are arranged side by side with a separating gap 13 between them.
  • the respective first segment 7a,8a has a respective first tangential sidewall 11a,12a that faces a second tangential sidewall 11b,12b of the respective second segment 7b,8b, such that the gap 13 extends between the sidewalls 11a and 11b and between 12a and 12b.
  • one or more perforations 14a are provided through the first sidewall 11a,12a and one or more perforations 14b are provided through the second sidewall 11b,12b for conducting cooling air therethrough.
  • the perforations 14a and 14b eject cooling air into the gap 13 to impinge on the sidewalls of the neighboring segment.
  • the perforations 14a and 14b conduct cooling air from a respective cavity 23a and 23b defined on faces 21a and 21b of the segments 7a and 7b that are radially opposite to faces 17a and 17b that are directly exposed to the drive gas.
  • the perforations 14a and 14b conduct cooling air from a respective cavity 23a and 23b defined on faces 22a and 22b of the segments 8a and 8b that are radially opposite to faces 18a and 18b that are directly exposed to the drive gas.
  • FIG 4 is illustrated an arrangement of the perforations 14a and 14b on opposite sidewalls relative to each other.
  • FIG 4 is a schematic illustration wherein the first segment depicted on the left side of FIG 4 could be any of the segments 5a,6a,7a,8a mentioned above. Accordingly the second segment depicted on the right side of FIG 4 would include any of the respective adjacent segments 5b,6b,7b,8b.
  • the respective tangential sidewalls are shown as 9a,10a,11a,12a and 9b,10b,11b,12b.
  • the streams of cooling air 24 ejected through the perforations 14b impinge on the sidewall 9a,10a,11a,12a on spaces between the perforations 14a on the sidewall 9a,10a,11a,12a.
  • the perforations 14a and 14b have a staggered arrangement relative to each other as shown, i.e., the perforations 14a and 14b are spaced apart with an axial shift relative to each other.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP10004074A 2010-04-16 2010-04-16 Turbineneinheit mit Kühlanordnung und Kühlverfahren Withdrawn EP2378071A1 (de)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP10004074A EP2378071A1 (de) 2010-04-16 2010-04-16 Turbineneinheit mit Kühlanordnung und Kühlverfahren

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP10004074A EP2378071A1 (de) 2010-04-16 2010-04-16 Turbineneinheit mit Kühlanordnung und Kühlverfahren

Publications (1)

Publication Number Publication Date
EP2378071A1 true EP2378071A1 (de) 2011-10-19

Family

ID=42676856

Family Applications (1)

Application Number Title Priority Date Filing Date
EP10004074A Withdrawn EP2378071A1 (de) 2010-04-16 2010-04-16 Turbineneinheit mit Kühlanordnung und Kühlverfahren

Country Status (1)

Country Link
EP (1) EP2378071A1 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106194277A (zh) * 2015-05-29 2016-12-07 通用电气公司 冲击冷却的花键密封件

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1022437A1 (de) * 1999-01-19 2000-07-26 Siemens Aktiengesellschaft Bauteil zur Verwendung in einer thermischen Machine
EP1176285A2 (de) * 2000-07-27 2002-01-30 General Electric Company Kühlung für einen Turbinenmantelring
US20040047725A1 (en) * 2002-09-06 2004-03-11 Mitsubishi Heavy Industries, Ltd. Ring segment of gas turbine
US20050100437A1 (en) * 2003-11-10 2005-05-12 General Electric Company Cooling system for nozzle segment platform edges
US20060093484A1 (en) * 2004-11-04 2006-05-04 Siemens Westinghouse Power Corp. Cooling system for a platform of a turbine blade
EP1749967A2 (de) * 2005-08-02 2007-02-07 Rolls-Royce plc Kühlsystem für eine Gasturbine

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1022437A1 (de) * 1999-01-19 2000-07-26 Siemens Aktiengesellschaft Bauteil zur Verwendung in einer thermischen Machine
EP1176285A2 (de) * 2000-07-27 2002-01-30 General Electric Company Kühlung für einen Turbinenmantelring
US20040047725A1 (en) * 2002-09-06 2004-03-11 Mitsubishi Heavy Industries, Ltd. Ring segment of gas turbine
US20050100437A1 (en) * 2003-11-10 2005-05-12 General Electric Company Cooling system for nozzle segment platform edges
US20060093484A1 (en) * 2004-11-04 2006-05-04 Siemens Westinghouse Power Corp. Cooling system for a platform of a turbine blade
EP1749967A2 (de) * 2005-08-02 2007-02-07 Rolls-Royce plc Kühlsystem für eine Gasturbine

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN106194277A (zh) * 2015-05-29 2016-12-07 通用电气公司 冲击冷却的花键密封件
CN106194277B (zh) * 2015-05-29 2020-09-22 通用电气公司 冲击冷却的键槽密封件

Similar Documents

Publication Publication Date Title
US9863254B2 (en) Turbine airfoil with local wall thickness control
US8281604B2 (en) Divergent turbine nozzle
EP2825748B1 (de) Kühlkanal für ein gasturbinentriebwerk und gasturbinentriebwerk
EP3244011B1 (de) System zum kühlen der dichtungsschienen einer schaufelspitzenabdeckung einer turbinenschaufel
US8858175B2 (en) Film hole trench
EP1959097B1 (de) Prallkühlung für Verkleidung und Innenteil einer Gasturbinenmotorschaufel
RU2577688C2 (ru) Лопатка для турбомашины и турбомашина, содержащая такую лопатку.
EP3064713B1 (de) Turbinenlaufschaufel und zugehöriger turbinenabschnitt
US8979481B2 (en) Turbine bucket angel wing features for forward cavity flow control and related method
US20190085705A1 (en) Component for a turbine engine with a film-hole
US9938835B2 (en) Method and systems for providing cooling for a turbine assembly
EP2586974B1 (de) Turbinenrotorschaufelplattform mit Vorderkantenwellung für Leistung und Sekundärströmung, zugehörige Turbinenrotor und Verfahren zur Steuerung der sekundären Abblasströmung
WO2007063128A1 (en) Blade platform cooling in turbomachines
US11773729B2 (en) Component for a gas turbine engine with a film hole
US9869185B2 (en) Rotating turbine component with preferential hole alignment
US9932835B2 (en) Airfoil cooling device and method of manufacture
US10458291B2 (en) Cover plate for a component of a gas turbine engine
US20170298743A1 (en) Component for a turbine engine with a film-hole
EP2378071A1 (de) Turbineneinheit mit Kühlanordnung und Kühlverfahren
EP2952682A1 (de) Schaufel für ein gasturbinenkraftwerk mit einer gekühlten plattform
CN110249112B (zh) 具有***件的涡轮发动机部件
US20130236329A1 (en) Rotor blade with one or more side wall cooling circuits

Legal Events

Date Code Title Description
AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO SE SI SK SM TR

AX Request for extension of the european patent

Extension state: AL BA ME RS

PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20120420