EP1219780A2 - Refroidissement d'un élément d'une turbomachine par jet - Google Patents

Refroidissement d'un élément d'une turbomachine par jet Download PDF

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Publication number
EP1219780A2
EP1219780A2 EP01129066A EP01129066A EP1219780A2 EP 1219780 A2 EP1219780 A2 EP 1219780A2 EP 01129066 A EP01129066 A EP 01129066A EP 01129066 A EP01129066 A EP 01129066A EP 1219780 A2 EP1219780 A2 EP 1219780A2
Authority
EP
European Patent Office
Prior art keywords
turbine blade
cooled
flow
cooling
wall section
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.)
Granted
Application number
EP01129066A
Other languages
German (de)
English (en)
Other versions
EP1219780B1 (fr
EP1219780A3 (fr
Inventor
Bernhard Dr. Prof. Weigand
James P. Downs
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.)
GE Vernova GmbH
Original Assignee
Alstom Technology AG
Alstom Power NV
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 Alstom Technology AG, Alstom Power NV filed Critical Alstom Technology AG
Publication of EP1219780A2 publication Critical patent/EP1219780A2/fr
Publication of EP1219780A3 publication Critical patent/EP1219780A3/fr
Application granted granted Critical
Publication of EP1219780B1 publication Critical patent/EP1219780B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection 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
    • 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 device for impingement cooling in one Fluid machine of heat-exposed component with a component to be cooled Wall section, the at least one side at least one impact cooling air flow exposed through a flow channel within a surface element, which is arranged at a distance from the wall section to be cooled, passes through and strikes the wall section to be cooled. Further relates the invention to a related cooling method and a method for Production of a heat-resistant component.
  • Cooling measures have been taken to precisely expose those to the hot gases Cool components, which reduces the thermal load as well as the Component life is to be increased.
  • those components that are supplied with cooling air are targeted Side of the compressor branched off in partial flows and by accordingly provided cooling channels is fed directly to the components to be cooled.
  • the maximum of the so-called external Heat transfer coefficients are typically at those points of the Turbine blade leading edge reached, on which the hot gases meet vertically and thus lead to a maximum accumulation effect on the front edge of the turbine blade. It is precisely these areas that need to be cooled with maximum efficiency in order to ensure that not to exceed material-related temperature limits.
  • a preferred technique for cooling the turbine blade leading edge is based on the targeted supply of cooling air within the turbine blade along in itself Inside the blade cooling channels, the cooling air directly on the The inside of the turbine blade leading edge is passed around the front edge to cool in a convective manner.
  • a turbine blade designed in this way is known, for example, from US Pat. No. 5,603,606 can be seen in the cross section through the Front region of a turbine blade is shown, the cooling air duct 166th having a connecting gap 180 with a front cooling volume 168 is connected, which is directly inside the turbine blade on the Turbine blade leading edge adjoins.
  • the connection channel 180 is on one side of the turbine blade inner wall is limited, which means that in the front Volume of cooling air introduced tangentially the inside of the Turbine blade leading edge overflows. This will make the entire area the turbine blade leading edge is subjected to an internal cooling air flow, however this can just the hot areas described above along the Insufficient cooling of the turbine blade leading edge.
  • FIG Area of a turbine blade 1 shown a cross section through the front is in FIG Area of a turbine blade 1 shown.
  • FIG Area of a turbine blade 1 shown schematically streamlines 2, which on the Turbine blade leading edge 3 should represent hot gases.
  • the turbine blade 1 provides an inner cooling duct 4, which passes through at least one connecting channel 5, which is located in a partition 8, with a front volume 6 is connected, in the same one with the Extending outlet duct 7 connected to the turbine blade top.
  • Cooling air which is supplied at high pressure via the cooling channel 4, passes through the Flow channel 5 in the volume 6 at high speed and hits the Area 3 of the turbine blade leading edge.
  • This is also known as impingement cooling
  • cooling technology can reinforces that area on the turbine blade leading edge that is cooled by the Hot gases are most thermally stressed. Closer examination of the known impingement air flow through the connecting channel in the direction of cooling turbine blade leading edge, however, show that the rectilinear flow channel only an expansion of the cooling jet when Allows leaving the connection channel. This will only cover small areas effectively applied cooling air to the inside of the turbine blade leading edge and the cooling effect is limited to a very restricted area.
  • Another disadvantage of the straight-line connecting channels is that that the emerging cooling jet massively cools a very small area and therefore too strong temperature gradients and caused by it Stress gradient in the material contributes.
  • the invention has for its object a device for impingement cooling in a turbomachine exposed to heat, preferably one Turbine blade, according to the type mentioned above evolve that area of heat-stressed Turbine blade leading edge can be cooled as effectively and optimally as possible that the pursuit of higher combustion chamber temperatures or a reduction the cooling air requirement used can be taken into account accordingly.
  • claim 12 The object underlying the invention is achieved in claim 1 specified.
  • the subject of claim 12 is an inventive method for Impingement cooling.
  • Claim 14 relates to a manufacturing method according to the invention Component according to claim 1.
  • Advantageously further developing the inventive concept Features are the subject of the subclaims and the entire description, in particular with reference to the exemplary embodiments according to drawings, refer to.
  • the device for impingement cooling is one in one Fluid machine of heat-exposed component with a component to be cooled Wall section, the at least one side at least one impact air cooling flow is exposed through a flow channel within a surface element is spaced from the wall section to be cooled, passes through and on strikes the wall section to be cooled, designed such that the Flow channel has an inlet and an outlet opening that the Outlet opening directly faces the wall section to be cooled and that the inlet opening has a flow cross-section that is smaller than that Flow cross section of the outlet opening is.
  • FIG. 1 shows the front area of a turbine blade 1 in Cross-sectional view with a main cooling channel 4, which via a flow channel 5 is connected to a front cooling volume 6, the rest of a Partition 8 is separated from the main cooling duct 4.
  • the front cooling volume 6 has an outlet channel 7, which opens on the surface of the turbine blade 1.
  • the Cooling air supplied through the main cooling duct 4 passes through the at high pressure Flow channel 5 through and meets the flow channel 5 opposite lying inner wall of the turbine blade 1 in the area of Turbine blade leading edge 3.
  • the flow channel 5 is included a flow cross section widening in the flow direction, so that in the form of an impact air cooling flow through the flow channel 5 cooling air passing therethrough emerges divergent from the flow channel 5 and thus applied to a larger area of the turbine blade leading edge 3.
  • FIG. 2 shows a detailed representation of the geometric design of the Flow channel 5, which is provided within the partition 8, the Main cooling duct 4 separates from the front volume 6.
  • the flow channel 5 has an inlet opening 9 and an outlet opening 10, the inlet opening 9 has a smaller cross section or smaller opening diameter than that Outlet opening 10.
  • the flow channel 5 flared and has straight cut Boundary walls. In principle, however, it is also possible to have a funnel shape to provide curved boundary wall contours. It is essential that the Baffle air cooling flow in the longitudinal direction propagating through the flow channel 5 Flow profile widens divergent after passing through the flow channel in this way the largest possible area of the turbine blade leading edge 3 to be acted upon by impingement air cooling.
  • the opening angle ⁇ shown in FIG. 2 is between the Center axis through the flow channel 5 and from a boundary wall included, values between 2 ° and 9 °.
  • Typical mean diameters for the flow channel 5 are in the range between 0.5 and 7 mm.
  • FIG. 3 shows a longitudinal section through the front area of a turbine blade 1 shown, with the main cooling channel 4, the front volume 6 and the Partition 8, in the radial to the turbine blade 1 distributed a large number of individual Flow channels 5 is provided. All individual flow channels 5 are of this type oriented relative to the turbine blade leading edge 3, so that the through the Individual impingement cooling streams passing through flow channels immediately Able to cool the inner wall of the turbine blade leading edge.
  • the conventional casting process which involves training within a mold of the flow channels designed according to the invention are heat-resistant Insert forms are provided which are subsequently removed from the casting, to expose the free flow channels.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP01129066A 2000-12-22 2001-12-07 Dispositif et procédé de refroidissement d'un élément d'une turbomachine par jet Expired - Lifetime EP1219780B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE10064271A DE10064271A1 (de) 2000-12-22 2000-12-22 Vorrichtung zur Prallkühlung eines in einer Strömungskraftmaschine hitzeexponierten Bauteils sowie Verfahren hierzu
DE10064271 2000-12-22

Publications (3)

Publication Number Publication Date
EP1219780A2 true EP1219780A2 (fr) 2002-07-03
EP1219780A3 EP1219780A3 (fr) 2004-08-11
EP1219780B1 EP1219780B1 (fr) 2006-11-02

Family

ID=7668444

Family Applications (1)

Application Number Title Priority Date Filing Date
EP01129066A Expired - Lifetime EP1219780B1 (fr) 2000-12-22 2001-12-07 Dispositif et procédé de refroidissement d'un élément d'une turbomachine par jet

Country Status (3)

Country Link
US (1) US6634859B2 (fr)
EP (1) EP1219780B1 (fr)
DE (2) DE10064271A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2196625A1 (fr) * 2008-12-10 2010-06-16 Siemens Aktiengesellschaft Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé

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GB2402715B (en) * 2003-06-10 2006-06-14 Rolls Royce Plc Gas turbine aerofoil
US7708525B2 (en) * 2005-02-17 2010-05-04 United Technologies Corporation Industrial gas turbine blade assembly
US7520725B1 (en) 2006-08-11 2009-04-21 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall leading edge multi-holes cooling
US20090255268A1 (en) * 2008-04-11 2009-10-15 General Electric Company Divergent cooling thimbles for combustor liners and related method
GB0811391D0 (en) * 2008-06-23 2008-07-30 Rolls Royce Plc A rotor blade
US9296039B2 (en) 2012-04-24 2016-03-29 United Technologies Corporation Gas turbine engine airfoil impingement cooling
DE102012016493A1 (de) * 2012-08-21 2014-02-27 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit prallgekühlten Bolzen der Brennkammerschindeln
US9394798B2 (en) * 2013-04-02 2016-07-19 Honeywell International Inc. Gas turbine engines with turbine airfoil cooling
US10309228B2 (en) * 2016-06-09 2019-06-04 General Electric Company Impingement insert for a gas turbine engine
KR102028804B1 (ko) * 2017-10-19 2019-10-04 두산중공업 주식회사 가스 터빈 디스크
US20190277501A1 (en) * 2018-03-07 2019-09-12 United Technologies Corporation Slot arrangements for an impingement floatwall film cooling of a turbine engine
US11391161B2 (en) 2018-07-19 2022-07-19 General Electric Company Component for a turbine engine with a cooling hole
US11998974B2 (en) 2022-08-30 2024-06-04 General Electric Company Casting core for a cast engine component
GB2639585A (en) * 2024-03-18 2025-10-01 Siemens Energy Global Gmbh & Co Kg A wall section of a hot gas path part for a gas turbine and a method of additive manufacturing of a wall section

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US5603606A (en) 1994-11-14 1997-02-18 Solar Turbines Incorporated Turbine cooling system

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US3844343A (en) * 1973-02-02 1974-10-29 Gen Electric Impingement-convective cooling system
US4775296A (en) 1981-12-28 1988-10-04 United Technologies Corporation Coolable airfoil for a rotary machine
US4770608A (en) * 1985-12-23 1988-09-13 United Technologies Corporation Film cooled vanes and turbines
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US5720431A (en) * 1988-08-24 1998-02-24 United Technologies Corporation Cooled blades for a gas turbine engine
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US5688104A (en) * 1993-11-24 1997-11-18 United Technologies Corporation Airfoil having expanded wall portions to accommodate film cooling holes
US5931638A (en) * 1997-08-07 1999-08-03 United Technologies Corporation Turbomachinery airfoil with optimized heat transfer
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GB2343486B (en) * 1998-06-19 2000-09-20 Rolls Royce Plc Improvemnts in or relating to cooling systems for gas turbine engine airfoil
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Patent Citations (1)

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Publication number Priority date Publication date Assignee Title
US5603606A (en) 1994-11-14 1997-02-18 Solar Turbines Incorporated Turbine cooling system

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2196625A1 (fr) * 2008-12-10 2010-06-16 Siemens Aktiengesellschaft Aube de turbine dotée d'un passage agencé dans une paroi de séparation et noyau de coulage associé

Also Published As

Publication number Publication date
EP1219780B1 (fr) 2006-11-02
US6634859B2 (en) 2003-10-21
DE50111357D1 (de) 2006-12-14
DE10064271A1 (de) 2002-07-04
EP1219780A3 (fr) 2004-08-11
US20020168264A1 (en) 2002-11-14

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