EP2921649A1 - Partie de profil aérodynamique d'une pale de rotor ou aube directrice d'une turbomachine - Google Patents

Partie de profil aérodynamique d'une pale de rotor ou aube directrice d'une turbomachine Download PDF

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Publication number
EP2921649A1
EP2921649A1 EP14160697.0A EP14160697A EP2921649A1 EP 2921649 A1 EP2921649 A1 EP 2921649A1 EP 14160697 A EP14160697 A EP 14160697A EP 2921649 A1 EP2921649 A1 EP 2921649A1
Authority
EP
European Patent Office
Prior art keywords
insert
airfoil portion
airfoil
cavity
trailing edge
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
EP14160697.0A
Other languages
German (de)
English (en)
Other versions
EP2921649B1 (fr
Inventor
Emanuele Facchinetti
Guillaume Wagner
Marc Henze
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.)
Ansaldo Energia IP UK Ltd
Original Assignee
Alstom Technology 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 Alstom Technology AG filed Critical Alstom Technology AG
Priority to EP14160697.0A priority Critical patent/EP2921649B1/fr
Priority to CN201510118463.2A priority patent/CN104929695B/zh
Priority to US14/661,318 priority patent/US20150267557A1/en
Priority to KR1020150037331A priority patent/KR20150109279A/ko
Priority to JP2015054860A priority patent/JP2015178833A/ja
Publication of EP2921649A1 publication Critical patent/EP2921649A1/fr
Application granted granted Critical
Publication of EP2921649B1 publication Critical patent/EP2921649B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
    • 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/31—Application in turbines in steam turbines
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • 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
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • 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
    • 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
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Definitions

  • the present invention relates to the field of rotor blades or guide vanes of a turbomachine, especially of a gas or steam turbine.
  • the final aim of the present invention is providing adequate cooling in a rotor blade or guide vane airfoil improving the cooling flow control and enabling insert fits.
  • the present invention relates to a rotor blade or guide vane airfoil assembling of a gas or steam turbine and refers fundamentally to a specific or modular arrangement of airfoil inserts within the cavity of the respective airfoil portion.
  • the specific or modular arrangement of airfoil inserts within the cavity of the respective airfoil portion consisting of replaceable and/or non-replaceable inserts.
  • the rotor blade or guide vane comprising additionally substitutable and non-substitutable flow-applied and no flow-applied elements.
  • the present invention relates to a turbine blade, namely as rotor blade or guide vane, with a hollow airfoil portion having an outer wall that defines a cavity for receiving cooling air, the airfoil portion comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, a convex suction side, a concave pressure side.
  • At least one insert disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion.
  • the insert further comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto.
  • US 8,182,203 B2 discloses a turbine blade including an airfoil; a supply channel extending through the interior of the airfoil in the span direction, through which cooling fluid flows; a pin fin channel extending from the supply channel along the center line of the airfoil toward the trailing edge of the airfoil and opening at the trailing edge to the exterior of the airfoil; a plurality of gap pin fins projecting from a pair of opposing inner walls that constitute the pin fin channel at a region at the supply channel side of the pin fin channel and forming a gap there between extending in the span direction; pin fins connecting the pair of opposing inner walls at a region at the trailing edge side of the pin fin channel; and an insertion portion disposed in the gap to decrease the area of the channel of the cooling fluid at the region at the supply channel side of the pin fin channel.
  • the insertion portion is disposed in the gap formed between the gap pin fins. Therefore, the cross-sectional area of the channel at the supply channel side of the pin fin channel, through which cooling fluid flows, decreases as compared with a case in which the insertion portion is not disposed, so that the velocity of the cooling fluid at the region at the supply channel side increases. This increases the cooling efficiency at the region at the supply channel side, which improves the cooling efficiency of the pin fin channel, thus improving the cooling performance of the turbine blade.
  • a vane for directing hot gases in a gas turbine engine.
  • the vane includes a hollow aerofoil portion, which in use spans the working gas annulus of the engine.
  • the vane further includes an impingement tube which forms a covering over the interior surface of the aerofoil portion and which has jet-forming apertures formed therein for the production of impingement cooling jets.
  • the impingement tube includes two tube portions which are separately insertable into position into the aerofoil portion to form the covering.
  • the impingement tube further includes an expansion member which, when the tube portions are in position in the aerofoil portion, is locatable in the aerofoil portion to urge each tube portion outwardly and thereby holds the tube portions in position against the aerofoil portion.
  • US 7,452,182 B2 relates to a modular guide vane assembly.
  • the vane assembly includes an airfoil portion, an outer platform and an inner platform.
  • the airfoil portion can be made of at least two segments.
  • the components are connected together so as to permit assembly and disassembly of the vane.
  • repair involves the replacement of only the damaged subcomponents as opposed to the entire vane.
  • the modular design facilitates the use of various materials in the vane, including materials that are dissimilar. Thus, suitable materials can be selected to optimize component life, cooling air usage, aerodynamic performance, and cost.
  • one end of the airfoil can be received within a recess in one of the inner and outer platforms.
  • the assembly can further include a seal provided between the recesses and at least one of the radial ending of the airfoil and the outer peripheral surface of the airfoil proximate the radial end.
  • one or more of the airfoil segments, the inner platform and/or the outer platform can be made of Intermetallics, Oxide Dispersion Strengthened (ODS) alloys, single-crystal metals, advanced Superalloys, metal matrix composites, ceramics or CMC.
  • ODS Oxide Dispersion Strengthened
  • the inventive idea of the present invention leaves the use of typical rotor blade or guide vanes assembling consisting of an airfoil portion, an inner and an outer platform, also called shroud, made in one piece as depicted and explained in connection with notorious state of the art.
  • a rotor blade or guide vane which can be assembled by at least two separate parts, i.e. a separate airfoil portion and outer platform and a separate inner platform, on the one hand preconditions are created to provide interchange ability or repairing and/or reconditioning of the identified separate parts, modules, elements without replacing the whole rotor blade or guide vane.
  • the present invention describes an improved rotor blade or guide vane assembling of a gas or steam turbine on the basis of a modular structure comprising fundamentally an airfoil portion, inner platform, outer platform, whereas the airfoil portion and/or the platforms having at its one end means for the purpose of an interchangeable connection of rotor blade or vane guide elements, whereas the connection of rotor blade or guide vane elements having a permanent or semi-permanent fixation with respect to the airfoil portion in radially or quasi-radially extension and with respect to the axis of the gas or steam turbine, whereas the assembling of the airfoil portion in connection with platforms based on a friction-locked bonding actuated by adherence interconnecting, or the assembling of the airfoil portion in connection with platforms based on the use of a metallic and/or ceramic surface the fixing guide vane elements to each other, or the assembling of the airfoil portion in connection with platforms based on force closure means with a detachable or permanent connection,
  • the present invention uses same or similar assemblies to determine the various possible connection of various configured airfoil inserts within the cavity of the airfoil portion.
  • one or more additional airfoil insert(s) can be used at the end of the respective airfoil insert.
  • the additional airfoil insert(s) could be inserted and slide in the trailing edge region before to put in place a main airfoil insert.
  • the additional airfoil insert(s) could optionally be cast in.
  • the insertion of the airfoil additional insert(s) is performed by a cascade principle with respect to their size, namely:
  • the additional insert (see Figure 1 , item 200) is inserted from the outside (see Figure 1 , item 201) into the cavity (see Figure 1 , item 202) and then moved (see Figure 1 , item 202) in direction to the trailing edge (see Figure 1 , item 103) and fixed to predetermined position.
  • At least one main airfoil insert may be inserted afterwards. It is also possible to proceed conversely.
  • At least one main insert comprising at least one additional insert which is inserted from the outside and transferred into the cavity at an intermediate position, and then moved in direction to the trailing edge and fixed to predetermined position, wherein the additional insert forms the size of the trailing edge channel inlet at the end of the main insert.
  • At least one main insert comprising at least one additional insert which forms the size of the trailing edge channel inlet at the end of the main insert.
  • This additional insert consists of a structured unitary body.
  • Different sized inserts can be arranged in the transverse direction of rotor blade or guide vane.
  • Various gaps between the airfoil inserts can be provided in all directions within the airfoil cavity on a case by case basis.
  • a joining assembling referring to airfoil inserts can be mechanically secured, or the joining assembling can use a shrinking process.
  • the detachable or permanent connection comprising a force closure with bolt or rivet, or by HT brazing, active brazing, soldering.
  • an individual insert can be made of one piece or of a composite structure.
  • the inserts are able to resist the caloric and physical stresses, wherein the mentioned means are holistically or on their part interchangeable among one another.
  • one of the basic idea of the invention consists to split one or more inserts within the cavity of the airfoil portion in multiple inserts in order to better adapt at the rotor blade or guide vane geometries, regardless of whether the respective rotor blade or guide vane consist of a unique body or a modular structure.
  • the invention provides adequate cooling in the airfoil, improving the flow control and enabling the insert fits.
  • one or more additional inserts can be used.
  • the additional insert(s) could be inserted and slide in the trailing edge region before to put in place the main insert(s).
  • the additional insert(s) could optionally be cast in.
  • the inserts can be made of the same material as the respective airfoil portion in which they are intercalated, such as IN939 alloy and ECY768 alloy.
  • the inserts can be made of a material that may or may not have a greater resistance to heat compared to the material of the airfoil portion.
  • the inserts can be made of a material with a lower heat resistance than the material of the receiving airfoil portion.
  • the inserts can be made from an inexpensive material so that the cost of a replacement insert would not significantly add to the overall costs over the life of the engine.
  • the mentioned airfoil portion inserts For insertion or removal purpose of the airfoil portion inserts it is possible to handle the mentioned airfoil portion inserts only at its radially outwards directed end which is a remarkable feature for performing maintenance work at the turbine stage.
  • radial is intended to mean radial to the turbine when the rotor blade or guide vane assembling is installed in its operational position.
  • a manner of attaching the airfoil portion and their insert portions to the inner respectively outer platform consists in doing the fact, that the radially end of the mentioned element can be received in a recess provided in the respective platform.
  • the mentioned recesses can be substantially airfoil-shaped so as to correspond to the outer contour of the airfoil portion and airfoil inserts.
  • the airfoil portion assembly including optionally an outer shell arrangement, can be trapped between the inner platform and the outer platform.
  • One of the most important solutions of the invention is to provide at least one outer shell and, if necessary and needed and according to individual operative requirements or different operating regimes, at least one no flow-applied intermediate shell in connection with the airfoil inserts for modular variants of the original airfoil portion.
  • Function of the airfoil carrier is to carry mechanical load from the airfoil module.
  • an outer and, additionally, an intermediate hot gas path shells are introduced.
  • the mentioned shells can be made of at least two segments.
  • the components forming the shell are connected together so as to permit assembly and disassembly of shell, shell components, airfoil portion and airfoil inserts of rotor blade or guide vane.
  • the seals can be at least one of rope seals, W-shaped seals, C-shaped seals, E-shaped seals, a flat plate, and labyrinth seals.
  • the seals can be made of various materials including, for example, metals and ceramics.
  • the airfoil portion 100 of a rotor blade or guide vane of a turbo-machinery is formed in a blade shape in cross section and extends in the span direction, namely in vertical direction of the blade.
  • the airfoil portion 100 has an integral cavity 101, which is a hollow formed at the leading edge 102 and extending in the flow direction of the airfoil portion 100 to the trailing edge 103. At least in the region of the leading edge 102 the external wall 104 of the airfoil portion 100 comprising a number of film-cooling holes 105 communicating with the front cavity 101.
  • the airfoil portion 100 has, in its interior, a first integrally cavity 101 extending in the flow direction or the airfoil portion 100.
  • the inner cavity 101 can be provided with at least one partition (not shown) in the manner that the partition may be divided the hollow portion into a front cavity and a rear cavity.
  • a cooling fluid derived from the exterior for example compressed air extracted from the compressor, cools adequately the structure of the airfoil portion 100.
  • a main hollow (205) insert 106 is disposed at a predetermined space from the inner wall of the cavity 101.
  • a rear insert is also disposed at a predetermined space from the inner wall of the rear cavity.
  • the film-cooling holes 105 are through-holes that connect the front cavity 101 and the exterior of the airfoil portion 100 and are provided with such film-cooling holes at intervals in the direction in a suction surface 107 and pressure surface 108.
  • the film cooling holes 105 are formed from the front cavity 101 to the exterior as slanting holes inclined from the leading edge 102 to the trailing edge 103
  • the rear cavity of the airfoil portion 100 is provided with a pin fin channel 109, which is a hollow extending from the rear cavity 101 toward the trailing edge 103 along a center line of the airfoil 100 (not shown) and which is a region in which gap pin fins 110 and pin fins 111 are provided.
  • the gap pin fins 110 are a plurality of substantially columnar members protruding from regions at the rear cavity side of the pin fin channel 109, the regions being a pair of inner walls constituting the pin fin channel 109.
  • the amount of protrusion of the gap pin fins 110 from the above-described inner walls is set so as to form a gap between the gap pin fins 110 into which the end portion of the rear or additional insert 200 can be inserted.
  • the pin fins 111 are a plurality of substantially columnar members that connect regions at the trailing edge 103 side of the pin fin channel 109, the regions being the pair of inner walls constituting the pin fin channel 109.
  • the shape and arrangement of the pin fins 111 can be known ones and are not particularly limited.
  • the pin fin channel 109 is a channel in the rear cavity in the region of the trailing edge 103, through which cooling fluid flows after being used for impinging cooling, and constitutes a structure related to pin fin cooling for cooling the vicinity of the trailing edge 103 of the airfoil portion 100 and opens to the exterior at the trailing edge 103.
  • pin fins 112 are provided, at least along the middle region of the airfoil portion, in the cavity between the inner wall of the airfoil portion 100 and the external wall of the main insert 106, along both the suction side 107 and the pressure side 108. Also the shape and arrangement of the pin fins 112 can be known ones and are not particularly limited.
  • the front of the main insert 106 constitutes a structure related to impinging cooling for cooling the leading edge 102 and the other inner wall of the airfoil portion 100, together with the front and the subsequent cavity 101.
  • the front of the main insert 106 consists of a substantially cylindrical member having a cross-sectional form similar to the cross-sectional form of the front cavity 101. Furthermore, the front of the main insert 106 has a plurality of discharge holes 113 through which the cooling fluid flowing there through spouts against the inner wall of the front cavity 101
  • the rear part of the insert constitutes also a structure related to impinging cooling, like the front insert, for cooling the respective side of the airfoil portion 100.
  • the rear insert consists also of a substantially cylindrical member having a cross-sectional form similar to the cross-sectional form of the rear part of the cavity.
  • an additional airfoil insert 200 is used.
  • the additional airfoil insert 200 is inserted and slide in the trailing edge region 103 before to put in place the main airfoil insert 106.
  • the additional airfoil insert 200 could optionally be cast in.
  • the insertion of the airfoil additional insert 200 is performed by a cascade principle with respect to their size, namely:
  • the additional airfoil insert 200 is inserted from the outside 201 and transferred 201 a into the cavity 101 at an intermediate position 202, and then moved 202a in direction to the trailing edge 103 and fixed to predetermined position 203.
  • the main airfoil portion insert 106 may be inserted afterwards. But it is also possible to proceed conversely.
  • the additional insert 200 is provided with a transversally elasticity 204, so that it can be pushed over the end side constriction of the main insert 106 until it reaches its final position 203.
  • the connection between the main 106 and the additional insert 200 is designed accordingly, even in the case in which the additional insert 200 does not have any transversally elasticity 204.
  • the connection can be obtained mechanically, for example with introduction of fixing members (not shown), positioned in the region of the both inserts.
  • the additional insert 200 which is inserted from the outside 201 and transferred 201 a into the cavity 101 at an intermediate position, and then it can be moved alternatively in direction to the leading edge 102 and fixed to final predetermined position. Additionally, the additional insert 200 can be inserted from the underside of the airfoil portion or is an element of the cavity of the airfoil portion, and then it can be moved in the direction of the trailing edge or leading edge and fixed to final predetermined position. Accordingly, the additional insert 200 forms the size of the trailing edge cavity at the end of the main insert 106, or the additional insert forms the leading edge cavity between inner wall of the airfoil portion and subsequent main insert.
  • an airfoil portion 100 of a rotor blade or guide vane of a turbo-machinery having an outer wall that defines a cavity for receiving cooling air, the airfoil portion comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, suction side, a pressure side.
  • At least one insert disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion.
  • the insert comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto.
  • At least one main insert 106 comprising at least one additional insert 200 which is inserted as a first option from the outside 201 and transferred 201 a into the cavity 101 at an intermediate position 202, and then the additional insert is moved 202a in direction to the trailing edge 103 or leading edge 102 and fixed to final predetermined position 203.
  • a second option consists in the fact that the additional insert 200 can be inserted from the underside of the airfoil portion or consists of an element of the cavity of the airfoil portion.
  • the additional insert is moved in the direction of the trailing edge or leading edge and fixed to final predetermined position, wherein the additional insert 200 forms at least one size of the trailing edge cavity at the end of the main insert 106, or the additional insert forms at least one leading edge cavity between inner wall of the airfoil portion and subsequent disposed main insert.
  • the airfoil portion 100a comprising a main insert 106a and an additional insert 250.
  • the additional insert is cast in.
  • FIG 2 illustrates a conventional air-cooled airfoil portion 100a.
  • the airfoil 100a includes an external wall 104, and has a leading edge 102, a pressure side 108 a suction side 107 and a trailing edge 103.
  • the airfoil 100a is generally hollow and, is divided into a main insert 106a and an attached no-hollow insert 250 in a preselected position.
  • the cooling structure is in generally the same like Figure 1 .
  • High pressure cooling air from the turbine compressor is directed into the main insert 106a per conventional systems and methods, and is exhausted through a number of discharge holes 113 to form jets of air striking the inner walls of the chambers 205 for impingement cooling 260.
  • the discharge holes 113 of main insert 106a in the cavity 101 are located to impinge on the external wall 104 opposite the main insert 106a.
  • the cooling air forced into the chamber 205 and through the main insert 106a is exhausted after a convective cooling 261 through radially spaced rows of film cooling 262 apertures 105 that pass through the external wall 104 of the airfoil portion 100a.
  • an airfoil portion 100a of a rotor blade or guide vane of a turbo-machinery having an outer wall that defines a cavity for receiving cooling air, the airfoil portion comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, a convex suction side, a concave pressure side, and at least one insert disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion.
  • the insert further comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto.
  • At least one main insert 106a comprising at least one additional insert 250 which forms the size of the trailing edge cavity at the end of the main insert 106a and/or at least one main insert comprising at least one additional insert which form the size of the leading edge cavity at the initiation of the main insert.
  • an airfoil portion 100b of a rotor blade or guide vane of a turbo-machinery having an outer wall 104 that defines a cavity (see also Figures 1 and 2 ) for receiving cooling air.
  • the airfoil portion comprising a leading edge 102 that resides in an upstream direction, a trailing edge 103 that resides in a downstream direction, a suction side and a pressure side.
  • Other design features can be taken from Figures 1 and 2 .
  • the main insert 106b together with a downstream side predisposed additional insert 300 (see also Figure 2 ), a number of inside predisposed additional inserts 320, 320a, 320b forming various cavities in which the flow of the cooling medium ensures an individual or subsequent and/or multiple cooling 263, 264 along these cavities.
  • the airfoil portion comprising a leading edge 102 that resides in an upstream direction, a trailing edge 103 that resides in a downstream direction, a convex suction side, a concave pressure side.
  • At least one main insert 106b is disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion.
  • the main insert 106b further comprising a configuration of a number of additional inserts 320, 320a, 320b that generally conforms to the inside contour of the main insert 106b.
  • the internal side of additional inserts 320, 320a, 320b forming a number of sub-cavities of the main insert 106b.
  • the main and/or additional inserts extend to radially or quasi-radially and/or transversally or quasi-transversally direction of the airfoil portion and are sectioned and having different shapes or profiles along one or more orientations of the airfoil portion, see Figures 4 and 5 , item 401, 402; 501-504.
  • the mentioned different shapes correspond to a regular or irregular triangular (see Figure 4 ), quadrangular, pentagonal, tapered body.
  • the airfoil portion having in radially or quasi-radially direction compared to the axis of the turbo-machinery a pronounced or swirled or tailored aero-dynamic profile. (see Figure 5 , item 500).

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP14160697.0A 2014-03-19 2014-03-19 Partie de profil aérodynamique d'une pale de rotor ou aube directrice d'une turbomachine Active EP2921649B1 (fr)

Priority Applications (5)

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EP14160697.0A EP2921649B1 (fr) 2014-03-19 2014-03-19 Partie de profil aérodynamique d'une pale de rotor ou aube directrice d'une turbomachine
CN201510118463.2A CN104929695B (zh) 2014-03-19 2015-03-18 涡轮机的转子叶片或导叶的翼型件部分
US14/661,318 US20150267557A1 (en) 2014-03-19 2015-03-18 Airfoil portion of a rotor blade or guide vane of a turbo-machine
KR1020150037331A KR20150109279A (ko) 2014-03-19 2015-03-18 터보 기계의 회전자 블레이드 또는 가이드 베인의 에어포일부
JP2015054860A JP2015178833A (ja) 2014-03-19 2015-03-18 ターボ機械のロータブレードまたはガイドベーンの翼部分

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WO2017108661A1 (fr) * 2015-12-23 2017-06-29 Siemens Aktiengesellschaft Aube de turbine pour une turbomachine thermique
US20170292195A1 (en) 2016-04-12 2017-10-12 United Technologies Corporation Light weight component with internal reinforcement and method of making
EP3231602A1 (fr) * 2016-04-12 2017-10-18 United Technologies Corporation Boîtier léger pour composant interne et procédé de fabrication
EP3323995A3 (fr) * 2016-11-17 2018-07-25 United Technologies Corporation Profil aérodynamique à déflecteur insérable latéralement
US10302017B2 (en) 2016-04-12 2019-05-28 United Technologies Corporation Light weight component with acoustic attenuation and method of making
US10323325B2 (en) 2016-04-12 2019-06-18 United Technologies Corporation Light weight housing for internal component and method of making
US10619949B2 (en) 2016-04-12 2020-04-14 United Technologies Corporation Light weight housing for internal component with integrated thermal management features and method of making
US10724131B2 (en) 2016-04-12 2020-07-28 United Technologies Corporation Light weight component and method of making
FR3094743A1 (fr) * 2019-04-03 2020-10-09 Safran Aircraft Engines Aube améliorée pour turbomachine
EP4006303A3 (fr) * 2020-11-12 2022-06-08 Solar Turbines Incorporated Ailette de refroidissement interne de paroi d'aube
CN114607471A (zh) * 2018-07-19 2022-06-10 通用电气公司 具有可调节冷却构造的翼型件
EP4283094A3 (fr) * 2022-05-27 2024-01-17 General Electric Technology GmbH Composant de turbine avec circuit de refroidissement pour la relaxation des contraintes
CN117489418A (zh) * 2023-12-28 2024-02-02 成都中科翼能科技有限公司 一种涡轮导向叶片及其前冷气腔的冷气导流件

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WO2018222326A1 (fr) * 2017-05-30 2018-12-06 Siemens Aktiengesellschaft Aube de turbine à bout aminci et couche renforcée par dispersion d'oxydes densifiée
KR101937589B1 (ko) * 2017-09-18 2019-04-09 두산중공업 주식회사 터빈의 터빈 블레이드와 터빈 베인 및 이를 포함하는 터빈 및 가스터빈
WO2019108216A1 (fr) * 2017-12-01 2019-06-06 Siemens Energy, Inc. Élément de transfert de chaleur brasé pour composants de turbine refroidis
DE102020103648A1 (de) * 2020-02-12 2021-08-12 Doosan Heavy Industries & Construction Co., Ltd. Pralleinsatz zur Wiederverwendung von Prallluft in einem Schaufelblatt, Schaufelblatt, das einen Pralleinsatz umfasst, Turbomaschinenkomponente und damit versehende Gasturbine
US11867085B2 (en) * 2020-03-25 2024-01-09 Mitsubishi Heavy Industries, Ltd. Turbine blade
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CN112943384A (zh) * 2021-05-14 2021-06-11 成都中科翼能科技有限公司 一种用于涡轮导向叶片的冷气导管结构

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WO2017082868A1 (fr) * 2015-11-10 2017-05-18 Siemens Aktiengesellschaft Surface portante stratifiée pour une turbine à gaz
US10370979B2 (en) 2015-11-23 2019-08-06 United Technologies Corporation Baffle for a component of a gas turbine engine
EP3170981A1 (fr) * 2015-11-23 2017-05-24 United Technologies Corporation Chicane pour un composant d'un moteur de turbine à gaz
US11035236B2 (en) 2015-11-23 2021-06-15 Raytheon Technologies Corporation Baffle for a component of a gas turbine engine
WO2017108661A1 (fr) * 2015-12-23 2017-06-29 Siemens Aktiengesellschaft Aube de turbine pour une turbomachine thermique
US10399117B2 (en) 2016-04-12 2019-09-03 United Technologies Corporation Method of making light weight component with internal metallic foam and polymer reinforcement
US11040372B2 (en) 2016-04-12 2021-06-22 Raytheon Technologies Corporation Light weight component with internal reinforcement
US10323325B2 (en) 2016-04-12 2019-06-18 United Technologies Corporation Light weight housing for internal component and method of making
US10335850B2 (en) 2016-04-12 2019-07-02 United Technologies Corporation Light weight housing for internal component and method of making
EP3231602A1 (fr) * 2016-04-12 2017-10-18 United Technologies Corporation Boîtier léger pour composant interne et procédé de fabrication
US10619949B2 (en) 2016-04-12 2020-04-14 United Technologies Corporation Light weight housing for internal component with integrated thermal management features and method of making
US10724131B2 (en) 2016-04-12 2020-07-28 United Technologies Corporation Light weight component and method of making
US10302017B2 (en) 2016-04-12 2019-05-28 United Technologies Corporation Light weight component with acoustic attenuation and method of making
US20170292195A1 (en) 2016-04-12 2017-10-12 United Technologies Corporation Light weight component with internal reinforcement and method of making
EP3323995A3 (fr) * 2016-11-17 2018-07-25 United Technologies Corporation Profil aérodynamique à déflecteur insérable latéralement
US10502070B2 (en) 2016-11-17 2019-12-10 United Technologies Corporation Airfoil with laterally insertable baffle
CN114607471A (zh) * 2018-07-19 2022-06-10 通用电气公司 具有可调节冷却构造的翼型件
FR3094743A1 (fr) * 2019-04-03 2020-10-09 Safran Aircraft Engines Aube améliorée pour turbomachine
US11795828B2 (en) 2019-04-03 2023-10-24 Safran Aircraft Engines Blade for a turbine engine, associated turbine engine distributor and turbine engine
EP4006303A3 (fr) * 2020-11-12 2022-06-08 Solar Turbines Incorporated Ailette de refroidissement interne de paroi d'aube
EP4283094A3 (fr) * 2022-05-27 2024-01-17 General Electric Technology GmbH Composant de turbine avec circuit de refroidissement pour la relaxation des contraintes
CN117489418A (zh) * 2023-12-28 2024-02-02 成都中科翼能科技有限公司 一种涡轮导向叶片及其前冷气腔的冷气导流件
CN117489418B (zh) * 2023-12-28 2024-03-15 成都中科翼能科技有限公司 一种涡轮导向叶片及其前冷气腔的冷气导流件

Also Published As

Publication number Publication date
KR20150109279A (ko) 2015-10-01
JP2015178833A (ja) 2015-10-08
EP2921649B1 (fr) 2021-04-28
CN104929695A (zh) 2015-09-23
CN104929695B (zh) 2021-09-24
US20150267557A1 (en) 2015-09-24

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