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EP1635042A1 - Refroidissement par impact d'un élément d'une turbomachine et méthode de fabrication d'une plaque de refroidissement par impact - Google Patents

Refroidissement par impact d'un élément d'une turbomachine et méthode de fabrication d'une plaque de refroidissement par impact Download PDF

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Publication number
EP1635042A1
EP1635042A1 EP04021701A EP04021701A EP1635042A1 EP 1635042 A1 EP1635042 A1 EP 1635042A1 EP 04021701 A EP04021701 A EP 04021701A EP 04021701 A EP04021701 A EP 04021701A EP 1635042 A1 EP1635042 A1 EP 1635042A1
Authority
EP
European Patent Office
Prior art keywords
passage
impingement cooling
component
impingement
cooling
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Withdrawn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
EP04021701A
Other languages
German (de)
English (en)
Inventor
Stefan Dr. Baldauf
Michael HÄNDLER
Gernot Lang
Christian Lerner
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 EP04021701A priority Critical patent/EP1635042A1/fr
Publication of EP1635042A1 publication Critical patent/EP1635042A1/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Withdrawn legal-status Critical Current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
    • 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
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/232—Three-dimensional prismatic conical
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/292—Three-dimensional machined; miscellaneous tapered
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/75—Shape given by its similarity to a letter, e.g. T-shaped
    • 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 invention relates to a cooled component of a turbomachine with a hot gasbeaufschlagbaren outer wall, the back is spaced to a provided with at least one passage impingement cooling element to form a gap, through which passage for cooling the outer wall, a cooling fluid as an impingement cooling jet flowable and on the back of the outer wall is alsprallbar , Moreover, the invention relates to a turbomachine with a hot gas-charged component and a method for producing a baffle cooling element with at least one passage.
  • Gas turbines with cooled components are well known. They burn a fuel with compressed air to a high pressure hot gas, which then relaxes work in a turbine unit on the rotor. In order to withstand the thermal stresses, the hot gas charged components are cooled. For this purpose, different cooling methods are known. In addition to convective cooling and film cooling, the impingement cooling of a component, for example a turbine blade, is known to be particularly efficient. To form a gap, each impingement-cooled component has at the back of its outer wall, which is exposed to hot gas, a baffle plate which is spaced apart from it and which is provided with vertically extending impingement-cooling bores. Through this flows an impingement cooling jet, which impinges on the back of the outer wall and this cools.
  • the impingement cooling is mostly very effective for flat, flat outside and rear sides.
  • the edge regions or angled corner regions of the intermediate spaces in particular in acute-angled corner regions, the most corresponding thereto executed outer wall can be effectively cooled only by means of the impact cooling effectively.
  • the object of the invention is to provide a cooled component with an impingement cooling, in which the effectiveness of the impingement cooling, in particular in the corner regions of a gap, is improved. It is another object of the invention to provide for this purpose a turbomachine and a method for producing an impingement cooling element with at least one passage.
  • the first object is achieved by the features of claim 1 and by the features of claim 5.
  • the object of the turbomachine is solved by the features of claim 8 and the object directed to the method by the features of claim 10 and by the features of claim 12.
  • the first inventive solution of the object directed to the component proposes that for individual adjustment of the impingement cooling jet at least one passage has an opening for the cooling fluid inlet opening with a cross-sectional area whose surface area is greater than the area cross-section of the cross-sectional area of the outlet-side opening of the passage.
  • the first solution ensures that the parameters such as, for example, the pressure and the speed of the cooling fluid flowing through the passage can advantageously be set as an impingement cooling jet.
  • an acceleration or a deceleration of the impingement cooling jet can be achieved in order to adapt the impingement cooling jet to the local, thermal requirements.
  • the second solution also enables targeted application of the cooling fluid to the rear side of the outer wall.
  • the two solutions are based on the common knowledge that the impingement cooling jet can be advantageously influenced to increase its cooling effect if the passage, that is to say the impingement cooling opening, deviates from the cylindrical, in particular the circular cylindrical cross-sectional shape.
  • the passage extends along a central axis which is inclined to a straight line which is perpendicular to a surface disposed on the impingement cooling element, on the rear side facing surface in the region of the passage.
  • a larger area of the outer wall can be achieved by the inclined impingement cooling jet. This is particularly advantageous in the edge or in the angled corner regions of a bump-cooled component.
  • the impingement cooling element is arranged locally closer to the back of the outer wall in the region of the passage than the areas of the impingement cooling element which have no passage.
  • the locally smaller distance between the back of the outer wall and the impingement cooling element in the immediate area of the passage increases the efficiency of the impingement cooling radiation.
  • the contour has a plurality of peaks in the manner of a cross or star, an equalization of the cooling effect of the impingement cooling jet can be achieved. This is particularly advantageous if it is to cool a flat surface, which, however, is thermally stressed differently and which, accordingly, applies different cooling.
  • the punctual cooling effect of the previously formed by drilling impingement cooling jet is replaced by a flat cooling effect. If necessary, necessary temperature gradients along the outer wall can be provided or avoided.
  • the component may be embodied as a turbine guide vane, as a turbine blade, as a guide ring, as a platform of a turbine blade or as a combustion chamber heat shield of a gas turbine.
  • the advantages directed to the component apply mutatis mutandis to the turbomachine.
  • a component according to the invention with a baffle cooling element can be produced in a particularly simple manner if the baffle cooling element is locally deformed in the immediate region of the passage or in the region in which a passage is provided, then that a trough-like depression results in the direction of the flow.
  • the local deformation can be inexpensively produced by means of an embossing process. In this case, both the nozzle shape and the inclination of the passage can be produced by the embossing process.
  • FIG. 1 shows a gas turbine 1 with a rotor 5 rotatably mounted about a rotation axis 3.
  • the gas turbine 1 has an intake chamber 7, a compressor 9, a toroidal annular combustion chamber 11 and a turbine unit 13.
  • Both in the compressor 9 and in the turbine unit 13 guide vanes 15 and blades 17 are each arranged in rings.
  • a blade ring 19 is followed by a blade ring 21.
  • the rotor blades 17 are fastened to the rotor 5 by means of rotor disks 23, whereas the stator blades 15 are fixedly mounted on the housing 25.
  • air 29 is sucked through the intake manifold 7 and compressed by the compressor.
  • the compressed air is guided to the burners 33, which are provided on a ring lying on the annular combustion chamber 11.
  • the compressed air 29 is mixed with a fuel 35, which Mixture in the annular combustion chamber 11 is burned to a hot gas 37.
  • the hot gas 37 flows through the flow channel 27 of the turbine unit 13 past guide vanes 15 and blades 17.
  • the hot gas 37 relaxes on the blades 17 of the turbine unit 13 to perform work.
  • the rotor 5 of the gas turbine 1 is set in a rotational movement, which serves to drive the compressor 9 and to drive a working engine, not shown.
  • FIG. 2 shows a section of the cross section of a component 41 of a gas turbine 1, which may be formed, for example, as a vane 15 or blade 17 of the turbine unit 13.
  • the component 41 has an outer wall 43 acted upon by the hot gas 37, on the rear side 45 of which an opposing space 51 is opposed by an impingement cooling element 47.
  • the impingement cooling element 47 is designed as an impingement cooling plate.
  • a plurality of passages 49 are provided as impact cooling holes.
  • a cooling fluid 53 for example cooling air or cooling steam, flows as an impingement cooling jet 55 through each passage 49.
  • Each impingement cooling jet 55 impinges on the rear side 45 of the outer wall 43 and cools it.
  • the passages 49 are designed to increase the cooling fluid flow in the manner of a nozzle.
  • the passages have a for the cooling fluid 53 on the inlet side cross-section C, which is greater than its exit-side cross-section D.
  • the built-in impingement cooling element 47 in the region 57 adjoining the passage 49 to the rear side 45 has a smaller distance B than the distances A, in which no passages 49 are provided.
  • the nozzle shown in FIG 2 can also be formed inclined or asymmetric, so that larger angular ranges can be achieved specifically deviating from a normal N by the impingement cooling jet 55. Consequently, for the passage 49 along its extension, an axis 61 is defined, which is inclined relative to the normal N, which is perpendicular to the sheet-like impingement cooling element 47.
  • the passage 63 has, contrary to the passages 49 of the prior art, no circular-cylindrical cross-section, but that of an n-sided polygon, here in a cross shape, which may be produced, for example, by a punching process.
  • the cross-shaped passages 63 are advantageously used for a surface cooling advantageous when they are arranged in a grid.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP04021701A 2004-09-13 2004-09-13 Refroidissement par impact d'un élément d'une turbomachine et méthode de fabrication d'une plaque de refroidissement par impact Withdrawn EP1635042A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP04021701A EP1635042A1 (fr) 2004-09-13 2004-09-13 Refroidissement par impact d'un élément d'une turbomachine et méthode de fabrication d'une plaque de refroidissement par impact

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP04021701A EP1635042A1 (fr) 2004-09-13 2004-09-13 Refroidissement par impact d'un élément d'une turbomachine et méthode de fabrication d'une plaque de refroidissement par impact

Publications (1)

Publication Number Publication Date
EP1635042A1 true EP1635042A1 (fr) 2006-03-15

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP04021701A Withdrawn EP1635042A1 (fr) 2004-09-13 2004-09-13 Refroidissement par impact d'un élément d'une turbomachine et méthode de fabrication d'une plaque de refroidissement par impact

Country Status (1)

Country Link
EP (1) EP1635042A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2955890A1 (fr) * 2010-02-02 2011-08-05 Snecma Secteur d'anneau de turbine de turbomachine
GB2492374A (en) * 2011-06-30 2013-01-02 Rolls Royce Plc Gas turbine engine impingement cooling
US11112113B2 (en) 2018-05-30 2021-09-07 Raytheon Technologies Corporation And manufacturing process for directed impingement punched plates

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1285369A (en) * 1969-12-16 1972-08-16 Rolls Royce Improvements in or relating to blades for fluid flow machines
US3844343A (en) * 1973-02-02 1974-10-29 Gen Electric Impingement-convective cooling system
US5083422A (en) * 1988-03-25 1992-01-28 General Electric Company Method of breach cooling
US6000908A (en) * 1996-11-05 1999-12-14 General Electric Company Cooling for double-wall structures
US20030021685A1 (en) * 2001-07-13 2003-01-30 Reinhard Fried Base material with cooling air hole
DE10202783A1 (de) * 2002-01-25 2003-07-31 Alstom Switzerland Ltd Gekühltes Bauteil für eine thermische Maschine, insbesondere eine Gasturbine

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1285369A (en) * 1969-12-16 1972-08-16 Rolls Royce Improvements in or relating to blades for fluid flow machines
US3844343A (en) * 1973-02-02 1974-10-29 Gen Electric Impingement-convective cooling system
US5083422A (en) * 1988-03-25 1992-01-28 General Electric Company Method of breach cooling
US6000908A (en) * 1996-11-05 1999-12-14 General Electric Company Cooling for double-wall structures
US20030021685A1 (en) * 2001-07-13 2003-01-30 Reinhard Fried Base material with cooling air hole
DE10202783A1 (de) * 2002-01-25 2003-07-31 Alstom Switzerland Ltd Gekühltes Bauteil für eine thermische Maschine, insbesondere eine Gasturbine

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2955890A1 (fr) * 2010-02-02 2011-08-05 Snecma Secteur d'anneau de turbine de turbomachine
GB2492374A (en) * 2011-06-30 2013-01-02 Rolls Royce Plc Gas turbine engine impingement cooling
US11112113B2 (en) 2018-05-30 2021-09-07 Raytheon Technologies Corporation And manufacturing process for directed impingement punched plates
EP3575688B1 (fr) * 2018-05-30 2022-06-29 Raytheon Technologies Corporation Procédé de conception et de fabrication pour plaques d'impact perforées dirigées

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