EP3231999A1 - Aube directrice dote de pale refroidie par couche d'air - Google Patents

Aube directrice dote de pale refroidie par couche d'air Download PDF

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Publication number
EP3231999A1
EP3231999A1 EP16164813.4A EP16164813A EP3231999A1 EP 3231999 A1 EP3231999 A1 EP 3231999A1 EP 16164813 A EP16164813 A EP 16164813A EP 3231999 A1 EP3231999 A1 EP 3231999A1
Authority
EP
European Patent Office
Prior art keywords
airfoil
film cooling
blade
short
cooling holes
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Withdrawn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
EP16164813.4A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Roland Häbel
Daniela Koch
Radan RADULOVIC
Marco Schüler
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 EP16164813.4A priority Critical patent/EP3231999A1/fr
Publication of EP3231999A1 publication Critical patent/EP3231999A1/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Withdrawn legal-status Critical Current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
    • 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
    • F01D5/186—Film 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading edge of a stator vane
    • 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
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/124—Fluid guiding means, e.g. vanes related to the suction side of a stator vane
    • 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
    • F05D2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the invention relates to a guide vane with at least one vane platform, which is arranged in the intended mounted state at a radially outer position, and extending from the at least one blade platform and extending in a longitudinal direction of the vane blade which defines a leading edge and a trailing edge and in the interior of which is provided a cavity which is connected to the outside of the airfoil by a plurality of film cooling holes formed in the airfoil and spaced from each other, at least one long array of film cooling holes lined up in a row being provided in the region of the leading edge of the airfoil extending from the blade platform in the longitudinal direction of the airfoil over at least 60% of the length of the airfoil.
  • the invention relates to a method for producing such a guide vane. Furthermore, the invention relates to a turbomachine, in particular a gas turbine, with such guide vanes and a method for refurbishing or producing a turbomachine using such guide vanes.
  • a guide vane comprises at least one vane platform, which is arranged in the intended mounted state at a radially outer position.
  • a vane includes an airfoil that protrudes from the at least one vane platform and extends in a longitudinal direction of the vane. The airfoil defines a leading edge and a trailing edge.
  • turbomachines such as gas turbines
  • a housing in which a flow passage extends in an axial direction.
  • the flow channel is flowed through during the operation of the turbomachine by a hot gas.
  • a plurality of turbine stages are arranged one behind the other in the axial direction and spaced from each other.
  • Each turbine stage includes a stator vane ring (stator) connected to the housing and a rotor blade (rotor) connected to a centrally mounted tie rod passing through the housing in the axial direction.
  • the expanding hot gas flowing through the flow channel is deflected by the guide vanes in such a way that the blades arranged behind are flown in optimally.
  • the rotor is set in rotation.
  • the rotational energy can be converted for example by means of a generator into electrical energy.
  • vanes are provided with elaborate coating systems.
  • guide vanes are cooled during operation of the gas turbine.
  • cavities are formed in their interior, which are flowed through by a cooling fluid.
  • Common cooling methods are, for example, the impingement cooling, in which the cooling fluid is guided so that it impinges on the wall of the blade from the inside, or the film cooling, in which the cooling fluid forms a cooling film on the outside of the blade.
  • a plurality of film cooling holes formed in the airfoil and spaced apart connect the cavity to the outside of the airfoil.
  • at least one long arrangement of line-like film cooling holes is provided which extends from the blade platform in the longitudinal direction of the blade over at least 60% of the length of the airfoil.
  • vanes wear during operation despite these measures.
  • high thermal and mechanical stress can occur that the guide vanes erode in such a way that particles detach, for example, from the coating system. These particles are transported by the hot gas through the flow channel and can impinge on vanes of subsequent turbine stages, whereby the struck blades are additionally exposed to a strong particle-induced erosion. Due to the centrifugal force acting in the hot gas and due to so-called horseshoe vortices generated by the vanes, the particles are accelerated radially outward and strike predominantly near the radially outer platform the suction sides of the vanes of subsequent turbine stages.
  • the invention is based on the idea of providing the guide vanes with greater protection in the area particularly subject to particle impact by forming additional film cooling bores.
  • the additional film cooling holes cause a denser cooling film in the region of the blade leaf in question, whereby impinging particles due to flow-mechanical reasons are prevented from impacting.
  • the airfoil body is cooled more strongly in the region in question, whereby its thermal stress remains below maximum permissible values even if the heat-insulating layer has been partially or completely removed by impacting particles.
  • the at least one long arrangement can be over at least 80% and preferably over the entire Length of the airfoil extend.
  • Such vanes are installed in many common gas turbines.
  • the at least one short arrangement extends over at most 30% and preferably over at most 20% of the length of the airfoil. This allows a better concentration of the additional film cooling on the platform-near region of the airfoil.
  • the film cooling bores of the at least one short arrangement can be offset relative to the film cooling bores of the at least one long arrangement in the longitudinal direction. Such offset of the film cooling holes can improve the fluidic barrier formed by the cooling film.
  • the distances between adjacent film cooling holes of an arrangement are identical over the entire length of the arrangement, wherein in particular the distances within the at least one short arrangement are greater than the distances within the at least one long arrangement.
  • This is a particularly easy to produce way to achieve the above offset between film cooling holes of the at least one short assembly and film cooling holes of the at least one long arrangement.
  • the short arrangements are arranged on the suction side of the airfoil. This allows greater protection to be focused on the most vulnerable area of the airfoil.
  • the maximum distance of the short arrays from the leading edge of the airfoil is in the range of 5% to 20% of the suction-side profile circumference between the leading edge and the trailing edge of the airfoil and is preferably 10% of this profile circumference. It has been shown that the particle-induced erosion is greatest in the immediate vicinity of the leading edge. Therefore, it may be sufficient to provide only the front fifth or a lesser portion of the suction side perimeter of the perimeter with short arrays of line-like film cooling holes.
  • the arrangements of film cooling holes starting from the leading edge of the blade in the direction of the trailing edge, are arranged one behind the other and at a distance from one another.
  • a plurality of long arrangements and / or a plurality of short arrangements are provided, wherein in particular the short arrangements are arranged between long arrangements and in particular short arrangements and long arrangements alternate.
  • This variant lends itself to the subsequent formation of short arrangements in vanes, which initially have several long arrangements.
  • the short arrangements can be placed in particular centrally between the long arrangements.
  • the lengths of the short arrangements advantageously decrease starting from the leading edge in the direction of the trailing edge, in particular in such a way that the short arrangements define a trapezoidal film cooling area. In this choice, it can be taken into account that particle-induced erosion takes place closer to the blade platform as the distance from the leading edge increases.
  • At least one film cooling hole in particular all the film cooling holes of the at least one short arrangement with respect to the vertical on the outer surface of the blade are in particular Tilt angle inclined.
  • the angle of inclination of the at least one film cooling bore differs from the angles of inclination of the film cooling holes of an adjacent long assembly.
  • This choice of tilt angles can further optimize the cooling film in the region in question with regard to its fluidic barrier effect.
  • the guide vane may be provided with a coating system comprising an adhesive layer (BC) and a thermal barrier coating (TBC), wherein the protective layer is applied to the airfoil and the thermal barrier coating adheres to the adhesive layer.
  • a coating system comprising an adhesive layer (BC) and a thermal barrier coating (TBC), wherein the protective layer is applied to the airfoil and the thermal barrier coating adheres to the adhesive layer.
  • the vane platform is inclined with respect to a plane perpendicular to the longitudinal direction of the airfoil. With such inclined blade platforms can form a conically widening in the flow direction of the hot gas flow channel.
  • a further blade platform can be connected to the blade opposite to the at least one blade platform. This further blade platform limits the flow channel on its radially inner side.
  • the present invention provides a turbomachine, in particular a gas turbine with a housing in which extends in an axial direction, a flow channel during the operation of the turbomachine flows through a hot gas, and a plurality of turbine stages, each comprising a arranged in the flow channel vane ring and are arranged in the axial direction behind each other and spaced from each other, wherein at least one turbine stage according to the invention comprises guide vanes.
  • the at least one turbine stage is arranged exclusively from the second position in the turbomachine. This follows from the fact that the blades of the first turbine stage can not be subject to particle-induced erosion because of their forward position.
  • the flow channel widens conically in the direction of flow of the hot gas, wherein the at least one turbine stage has vanes with inclined blade platforms and the inclined blade platforms of these vanes delimit the flow channel to the outside.
  • the conical widening of the flow channel in the flow direction of the hot gas represents a design feature to be found in many turbomachines.
  • vanes can also be made originally with long and short arrangements of line-like juxtaposed film cooling holes.
  • the at least one short arrangement is formed on the suction side of the airfoil. Especially the suction side is exposed to a particle-induced erosion.
  • film cooling holes of the at least one long array can be closed and / or cross-sectional areas of film cooling holes of the at least one long array can be reduced such that the sum of the cross-sectional areas of all the film cooling holes of the manufactured vane is at most 10% greater than the sum of Cross sectional areas of all film cooling holes of the provided guide vane and is preferably identical to this.
  • the present invention provides a method for refurbishing or producing a turbomachine having a plurality of turbine stages, in particular a gas turbine, in which guide vanes according to the invention are arranged exclusively from the second turbine stage in the gas turbine.
  • the guide vanes according to the invention are not required in the first turbine stage, since the guide vanes of this turbine stage are not subject to particle-induced erosion.
  • the FIG. 1 shows a guide blade 1 for a turbomachine, not shown, in particular gas turbine.
  • the vane 1 comprises a blade platform 2, which is arranged in the intended mounted state at a radially outer position.
  • the guide vane 1 comprises an airfoil 3 which protrudes from the vane platform 2 and extends in a longitudinal direction L of the vane 1.
  • the airfoil 3 defines a leading edge 4 and a trailing edge 5 and a cavity 6 provided in its interior.
  • a plurality of film cooling holes 7 are arranged at a distance from each other, which connects the cavity 6 with the outside of the airfoil 3.
  • two long arrangements 8 of line-like film cooling holes 7 are provided, which extend from the blade platform 2 in the longitudinal direction L of the blade 3 over approximately 80% of the length of the blade 3.
  • a short arrangement 9 of line-like film cooling holes 7 is provided which extends from the blade platform 2 in the longitudinal direction L of the blade 3 over approximately 40% of the length of the blade 3.
  • the distance of the short assembly 9 from the leading edge 4 of the airfoil 3 is about 15% of the suction side profile circumference between the leading edge 4 and the trailing edge 5 of the airfoil.
  • Both the long arrays 8 and the short arrays 9 of film cooling holes 7 are arranged on the suction side 10 of the airfoil 3.
  • the long and short arrangements 8, 9 of film cooling holes 7 are arranged starting from the leading edge 4 of the blade 3 in the direction of the trailing edge 5 one behind the other and spaced from one another.
  • the short arrangement 9 is formed centrally between the two long arrangements 8, so that the long arrangements 8 and the short arrangement 9 are arranged alternately.
  • the distances between adjacent film cooling holes 7 are identical.
  • the distances within the short arrangement 9 are greater than the distances within the two long arrangements 8, so that the film cooling holes 7 of the short arrangement 9 are offset from the film cooling holes 7 of the two long arrangements 8.
  • the vane 1 is provided with a coating system (not shown) comprising an adhesive layer (Bond Coating BC) and a thermal barrier coating (TBC).
  • a coating system comprising an adhesive layer (Bond Coating BC) and a thermal barrier coating (TBC).
  • Bond Coating BC Ad Coating BC
  • TBC thermal barrier coating
  • the blade platform 2 is inclined to a plane perpendicular to the longitudinal direction L of the blade 3. Opposite the blade platform 2, a further blade platform 11 is connected to the blade 3.
  • the two blade platforms 2, 11 limit in the intended mounted state a flow channel of the turbomachine.
  • the turbomachine comprises a housing 12 in which a flow channel 13 extends in an axial direction. Furthermore, the turbomachine comprises a plurality of turbine stages 14, each of which comprises a guide vane ring 15 arranged in the flow channel 13 and arranged one behind the other in the axial direction A and at a distance from one another. Only the vane rings 15 from the second position in the turbomachine have inventive guide vanes 1 according to the first embodiment of the present invention. As a result of the inclined blade platform 2 of the guide vanes 1, which limit the flow channel 13 to the outside, the flow channel 13 widens conically in the flow direction of the hot gas.
  • the flow channel 13 is flowed through by a hot gas.
  • the guide vanes 1 are each traversed by a cooling fluid.
  • the cooling fluid flows from the cavity 8 through the film channel holes 7 on the outside 10 of the airfoil 3.
  • the effluent cooling fluid forms a cooling film that protects the outside 10 of the airfoil 3 from the flowing hot gas and reduces the thermal load on the airfoil 3 .
  • the additional film cooling holes 7 of the short assembly 9 cause a denser and stronger cooling film, so that particles that have detached from vanes 1 upstream turbine stages 14, are prevented from being impacted on the suction side 10 of the airfoil 3 or previously at least braked so that the damage caused to the coating system of the vane 1 is relatively small.
  • the short assembly 9 is arranged in this particularly affected area of the airfoil 3.
  • FIGS. 2 and 3 shows the airfoil 3 of a vane 1 according to a second embodiment of the present invention. It has the same basic structure as the one in FIG. 1 shown guide vane 1 according to the first embodiment. Deviating from this, a plurality of long arrangements and a plurality of short arrangements are provided on the suction side 10 of the airfoil 3, which are arranged alternately. In this case, the lengths of the short arrangements 9, starting from the leading edge 4 in the direction of the trailing edge 5, decrease such that the short arrangements 9 define a trapezoidal film cooling area. Another difference is that the distances between film cooling holes 7 within the short assembly 9 are as large as the distances of the film cooling holes 7 of the long assemblies. 8
  • FIG. 4 shows a vane 1 according to a third embodiment of the present invention. It has the same basic structure as the one in FIG. 1 Deviating from the distances between film cooling holes 7 within the short assembly 9 are as large as the distances of the film cooling holes 7 within the two long arrangements 8. Accordingly, the film cooling holes 7 of the short assembly 9 are not opposite the film cooling holes 7 of the two long arrangements 8 arranged offset in the longitudinal direction L of the airfoil 3. In addition, the long assemblies 8 extend over only about 60% of the length of the airfoil 3.
  • FIG. 5 finally shows a vane 1 according to a fourth embodiment of the present invention.
  • the guide vane 1 corresponds in terms of their basic structure of in FIG. 1
  • the film cooling holes 7 of both the long assemblies 8 and the short assemblies 9 are inclined relative to the vertical to the outer surface 10 of the airfoil 3 by certain inclination angle.
  • the angles of inclination within an arrangement 8, 9 are essentially identical, while the angles of inclination between long arrangements 8 and short arrangements 9 are different, so that the cooling fluid in each case flows out of the cooling fluid bores 7 in opposite directions.
  • the protective effect of the cooling fluid film on the outside 10 of the airfoil 3 can be enhanced in the area particularly affected by the particle-induced erosion.
  • a guide blade 1 having the features described above is provided, but without short arrangements 9 of film cooling holes.
  • film cooling holes 7 of the at least one long arrangement 8 can be closed.
  • the cross-sectional areas of film cooling holes 7 of the at least one long assembly 9 can be reduced. In this way it is achieved that the sum of the cross-sectional areas of all film cooling holes 7 of the manufactured vane 1 is at most 10% greater than the sum of the cross-sectional areas of all film cooling holes 7 of the provided guide vane and ideally identical to this.
  • guide vanes according to the invention are arranged exclusively from the second turbine stage in the gas turbine. This is associated with a longer life of the affected by a particle-induced erosion vanes 1, which allows longer maintenance intervals of the turbomachine. If the cooling fluid flow through the guide vanes 1 at least does not increase significantly compared with the cooling fluid flow through conventional guide vanes, the guide vanes 1 according to the invention can readily be installed, in particular without having to make changes in the cooling fluid circuit of the turbomachine.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP16164813.4A 2016-04-12 2016-04-12 Aube directrice dote de pale refroidie par couche d'air Withdrawn EP3231999A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP16164813.4A EP3231999A1 (fr) 2016-04-12 2016-04-12 Aube directrice dote de pale refroidie par couche d'air

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP16164813.4A EP3231999A1 (fr) 2016-04-12 2016-04-12 Aube directrice dote de pale refroidie par couche d'air

Publications (1)

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EP3231999A1 true EP3231999A1 (fr) 2017-10-18

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EP16164813.4A Withdrawn EP3231999A1 (fr) 2016-04-12 2016-04-12 Aube directrice dote de pale refroidie par couche d'air

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113803116A (zh) * 2021-09-18 2021-12-17 沈阳航空航天大学 一种具有收缩型端壁气膜孔冷却结构的涡轮转子叶片

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1226751A (fr) * 1966-02-09 1971-03-31
EP0641917A1 (fr) * 1993-09-08 1995-03-08 United Technologies Corporation Refroidissement du bord d'attaque d'une aube
GB2455899A (en) * 2007-12-31 2009-07-01 Gen Electric Turbine nozzle cooling
EP2093382A2 (fr) * 2008-02-20 2009-08-26 United Technologies Corporation Aube statorique pourvue de trous de refroidissement disposés en éventail sur un large raccordement entre profil et plateforme

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1226751A (fr) * 1966-02-09 1971-03-31
EP0641917A1 (fr) * 1993-09-08 1995-03-08 United Technologies Corporation Refroidissement du bord d'attaque d'une aube
GB2455899A (en) * 2007-12-31 2009-07-01 Gen Electric Turbine nozzle cooling
EP2093382A2 (fr) * 2008-02-20 2009-08-26 United Technologies Corporation Aube statorique pourvue de trous de refroidissement disposés en éventail sur un large raccordement entre profil et plateforme

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113803116A (zh) * 2021-09-18 2021-12-17 沈阳航空航天大学 一种具有收缩型端壁气膜孔冷却结构的涡轮转子叶片

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