WO2016206944A1 - Ensemble turbine - Google Patents

Ensemble turbine Download PDF

Info

Publication number
WO2016206944A1
WO2016206944A1 PCT/EP2016/062581 EP2016062581W WO2016206944A1 WO 2016206944 A1 WO2016206944 A1 WO 2016206944A1 EP 2016062581 W EP2016062581 W EP 2016062581W WO 2016206944 A1 WO2016206944 A1 WO 2016206944A1
Authority
WO
WIPO (PCT)
Prior art keywords
surface increasing
turbine assembly
increasing structure
cooling
aerofoil
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.)
Ceased
Application number
PCT/EP2016/062581
Other languages
English (en)
Inventor
Henrik Hull
Janos Szijarto
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
Siemens Corp
Original Assignee
Siemens AG
Siemens Corp
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, Siemens Corp filed Critical Siemens AG
Publication of WO2016206944A1 publication Critical patent/WO2016206944A1/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • F05D2230/211—Manufacture essentially without removing material by casting by precision casting, e.g. microfusing or investment casting
    • 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
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing 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
    • F05D2240/126—Baffles or ribs
    • 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
    • F05D2240/127—Vortex generators, turbulators, or the like, for mixing
    • 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
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade

Definitions

  • the present invention relates to a turbine assembly like an aerofoil-shaped turbine assembly such as turbine rotor blades and stator vanes.
  • the present invention further relates to a matrix component and method for manufacturing said turbine assembly .
  • high temperature turbines may include hollow blades or vanes comprising cool ⁇ ing passages for a cooling medium to cool the blades and vanes during operation of the turbine assembly.
  • Such blades or vanes with e.g. an inner serpentine geometry for the cooling passages are typically made by an investment casting process which uses a ceramic core to define the vari ⁇ ous internal passages and further structures. After casting, the ceramic core is removed from the aerofoil by a leaching process .
  • the cooling passages terminate into a vertical slot posi ⁇ tioned at a trailing edge of the aerofoil. Due to different flow dynamics inside the hollow aerofoil local heat distribu ⁇ tion along the slot may vary. Hence, selected regions may overheat and be subjected to detrimental heat effects which, in turn, can lead to failure of the aerofoil and thus the turbine engine. It is a first objective of the present invention to provide an aerofoil-shaped turbine assembly such as a turbine rotor blade and a stator vane with which the above-mentioned short- comings can be mitigated, and especially that an effective cooling concept can be realised which provides sufficient cooling and ensures effective heat transfer.
  • a third ob ⁇ jective of the invention is to provide a method for manufac ⁇ turing a turbine assembly that is reliable and time- efficient .
  • the present invention provides a turbine assembly comprising at least one basically hollow aerofoil with at least one trailing edge comprising at least two wall segments that each comprises an inner surface, wherein the inner surfaces face each other, and wherein the inner surfaces enclose at least one exit aperture extending basically in span wise direction of the at least one basically hollow aerofoil, wherein that at least one of the inner surfaces of the at least two wall segments of the trailing edge comprises at least one surface increasing structure extending partially into the at least one exit aperture from one inner surface to the other inner surface, wherein the turbine assembly further comprises at least one cooling passage for a cooling medium discharging in the at least one exit aperture and wherein the at least one cooling passage has an exit opening and wherein the at least one exit aperture has a depth in a direction ba ⁇ sically perpendicular to a span wise direction of the basi- cally hollow aerofoil and wherein the at least one surface increasing structure has a longitudinal extension with a forward end and a rear end and wherein the at least one surface increasing structure extend
  • a turbine assembly is intended to mean an assembly provided for a turbine engine, like a gas turbine, wherein the assem ⁇ bly possesses at least one aerofoil.
  • the turbine assembly may be a part of a turbine wheel or a turbine cascade with circumferential arranged aerofoils.
  • the turbine assembly may comprise in addition to the aerofoil further structures, like a root portion and/or an outer and/or an inner platform. The latter two would be arranged at opposed ends of the aero- foil (s) and/or the inner platform would be arranged between the aerofoil and the root portion.
  • a “basically hollow aerofoil” means an aero ⁇ foil with a casing, wherein the casing encases at least one cavity and/or an e.g. meandering cooling passage.
  • a structure like a rib, which divides different cavities/passages in the aerofoil from one another and for example extends in a span wise direction of the aerofoil, does not hinder the definition of "a basically hollow aerofoil".
  • the basically hollow aerofoil referred to as aerofoil in the following description, has two cooling regions, a channelled cooling region at a leading edge of the aerofoil and pin- fin/pedestal cooling region at the trailing edge. These re- gions could be separated from one another through a rib.
  • the wall segment is preferably a part of either a suction side or a pressure side of the aerofoil.
  • the sur ⁇ face increasing structure may also be positioned or arranged at or formed in a partitioning wall dividing the exit aperture in the direction from the pressure to the suction side.
  • a "surface increasing structure” is intended to mean any structure, like a knoll, rib, hoop or the like, which increases the surface of the wall to which the struc ⁇ ture is attached or where it is positioned or in which it is formed that is feasible for a person skilled in the art.
  • the attachment may be facilitated by any connection feasible for a person skilled in the art, like screwing, welding, gluing, etc.
  • an effective surface area comprises all sur ⁇ faces that interact with the cooling medium bypassing the surface increasing structure.
  • the effective surface comprises the surfaces extending along a width and along a length of the surface increasing structure as well as both sloping flanges of the surface in ⁇ creasing structure extending to the wall segment having the surface increasing structure.
  • the surface increasing structure has a width of 0.4 mm, a height of
  • the at least one surface increasing structure and the wall segment of the at least two wall segments that is embodied with the at least one surface increasing structure are embodied integrally with each other.
  • the surface increasing structure can be cast during the casing process of the tur- bine assembly and thus advantageously economising the manu ⁇ facturing process.
  • smooth junctions between the surface increasing structure and the wall segment can be pro ⁇ vided.
  • formed integrally is intended to mean that the surface increasing structure and the wall seg- ment are built from the same piece, that they are a one-piece part or that they cannot be separated from each other without damaging at least one of the structures or that there is a loss of function for at least one of the structures.
  • An "exit aperture” is intended to mean an opening or hole which primary function is to provide an exit for a selected structure or means, like a fluid, a gas, a liquid or the like. Generally, another function may be also possible, like a fastening function etc.
  • the selected structure may be any structure feasible for a person skilled in the art and is preferably a cooling medium traveling at least one cooling passage inside of the aerofoil for example with a flow path or direction leading basically from the leading edge to the trailing edge.
  • the exit aperture may have any shape or cross section feasible for a person skilled in the art, like circular, triangular, rectangular or oval etc.
  • the exit aperture is a slot.
  • the exit aperture or the slot, respectively, may extend over a part of the trail ⁇ ing edge viewed in span wise direction or basically all along a span wise extension of the trailing edge.
  • a span wise direction of the aerofoil is de- fined as a direction extending basically perpendicular, preferably perpendicular, to a direction from a leading edge to a trailing edge of the aerofoil or the direction from a root portion to a tip of the aerofoil or the direction from one platform to the opposed platform in case of the embodiment of the aerofoil as a vane.
  • the phrase "extending partially into” should be understood that the surface increasing structure originates at one wall segment and extends in direction to the opposed wall segment or its inner surface, respectively, thereby crossing a part of the exit aperture. Specifically, the surface increasing structure does not physically connect the two wall segments with each other. Thus, there is always a gap between the top of the surface increasing structure and the opposed arranged wall segment.
  • the dimensions of the exit aperture and of the surface in ⁇ creasing structure depend on the dimensions of the turbine assembly or its aerofoil, respectively. These dimensions may be specified as relative values or as absolute values.
  • the at least one exit aperture has a width ex ⁇ tending from one inner surface to the other inner surface and wherein the at least one surface increasing structure has a height that is maximally 50% of the width if the at least one exit aperture.
  • maximally 50% should lie a divergence from 50% of about 5% to 10%.
  • the at least two wall segments are spaced apart a distance from each other and wherein the at least one surface increasing structure extends maximally along 50% of the dis- tance between the at least two wall segments.
  • the exit aperture may have a width of about 0.5 mm to 0.9 mm, thus the height of the surface increasing struc ⁇ ture would maximally be about 0.25 mm to 0.45 mm.
  • the at least one aerofoil has a span in span wise direction and wherein a width of the at least one surface increasing structure is greater than 0.4% of the span of the at least one aerofoil or preferably of about 0.4% to 1% of the span and most preferably of about 0.5% of the span.
  • a width of the at least one surface increasing structure is greater than
  • a height of the at least one surface in ⁇ creasing structure is greater than 0.3% of the span of the at least one aerofoil or preferably of about 0.3% to 1% and most preferably of about 0.4% of the span.
  • the height of the at least one surface increasing structure is greater than 0.3 mm.
  • a length of the at least one surface increasing structure is greater than 1% of the span of the at least one aerofoil or preferably of about 1% to 3% and most preferably of about 2% of the span.
  • the length of the at least one surface increasing structure is greater than 1.4 mm.
  • each surface increasing structure may be dimension as well as an exit opening or cross section of each cooling passage individually.
  • the turbine assembly comprises at least one cooling passage for the cooling medium discharging in the at least one exit aperture, thus, providing a selected flow path for the cooling medium.
  • the at least one cooling passage has a cross sectional area and the at least one sur ⁇ face increasing structure has a cross sectional area.
  • the cross sec ⁇ tional area of the at least one surface increasing structure is less than 60% of the cross sectional area of the at least one cooling passage and preferably about 40% to 60% and most preferably the cross sectional area of the at least one sur- face increasing structure is maximally 50% of the cross sec ⁇ tional area of the at least one cooling passage.
  • the portion of the cross sectional area of the cooling passage where the flow of the cooling medium is unhindered by the surface increasing structure is large enough to allow a suf ⁇ ficient amount of cooling medium to be discharged from the cooling passage to bypass the surface increasing structure for heat transfer and thus cooling.
  • the at least one surface increasing structure is positioned at the hottest point of the at least one aerofoil during operation of the turbine assembly.
  • the surface increasing structure can take effect at the region of the aerofoil where cooling is needed the most and the risk for detriment is highest.
  • This location can be determined before ⁇ hand of the casting of the aerofoil with specific software under consideration of characteristics of the turbine assem ⁇ bly as well as the used turbine engine, like used materials, position of the turbine assembly along the flow path of the cooling medium, temperatures at selected positions of the turbine engine, flow capacity and velocity, speed of the tur ⁇ bine engine etc.
  • the at least one cooling passage has an exit opening and the at least one exit aperture has a depth in a direction basically perpendicular to a span wise direction of the basically hollow aerofoil.
  • an "exit open ⁇ ing" is intended to mean the region where the cooling medium, like air, is discharged from the cooling passage and specifi- cally into the exit aperture to be released from the turbine assembly into the surrounding environment.
  • the portions of the flow medium dis ⁇ charged from different cooling passages can be merged after passing the exit opening.
  • the scope of a "basically perpen- dicular" arrangement should also lie an arrangement of the depth in respect to the span wise direction that deviates up to 10% from the strict span wise orientation.
  • the at least one surface increasing structure has a longitudinal extension with a forward end and a rear end and wherein the at least one surface increasing structure extends from its forward end to its rear end along the depth of the at least one exit aperture and wherein the forward end of the at least one surface increasing structure is arranged in the exit opening of the at least one cooling passage. Due to this arrangement the cooling of the discharged cooling me ⁇ dium can start to take effect directly after the cooling me- dium exits the cooling passage.
  • the arrangement as "arranged in” should be understood as positioned at or
  • the exit opening and the forward end are basically at the same depth viewed in direction from the leading edge to the trailing edge.
  • the surface increasing structure with its forward and rear ends is preferably embod ⁇ ied as a rib.
  • An especially effective heat transfer can be provided when the turbine assembly comprises a plurality of surface in ⁇ creasing structures. Furthermore, it is advantageous when the plurality of surface increasing structures extend basically in parallel towards each other, thus minimising the risk of flow perturbation along the flow path of the cooling medium after exiting the cooling passage (s) .
  • a homogeneous discharge can be provided when the surface increasing structures of the plurality of surface increasing structures are arranged equally spaced towards each other. It is possible to arrange the surface increasing structures in span wise direction of the aerofoil all along the exit ap ⁇ erture. However, to allow a sufficient amount of the cooling medium/air to exit the aerofoil unhindered the positioning of the surface increasing structures (only) at selected areas might be more advantageous.
  • the surface increasing structures can be arranged in one or more groups with a se ⁇ lected number of surface increasing structures. For example, it would be possible to arrange a group of six surface in ⁇ creasing structures at the region where the aerofoil will be hottest during operation of the turbine engine, like near the tip of the aerofoil.
  • each of the cooling passages of the plurality of cooling passages comprises at least one surface increasing structure of the plurality of surface increasing structures.
  • good heat transfer can be provided at each cooling passage.
  • each of these cooling pas ⁇ sages would have at least one surface increasing structure.
  • the cooling passage (s) might have a dimension that would al ⁇ low to arrange more than one surface increasing structure per cooling passage.
  • the at least one cooling passage has a longitudinal extension and the at least one surface increasing structure has a longitudinal extension.
  • the longitudinal extension of the at least one surface increasing structure and the longitudinal extension of the at least one cooling passage are arranged aligned with each other.
  • the longitudinal ex ⁇ tension of the at least one surface increasing structure and the axial extension of the at least one cooling passage ex ⁇ tend basically coaxial with each other or the direction of the flow path of the cooling medium is coaxial to the extension of the surface increasing structure.
  • the turbine assembly has a suction side and a pressure side.
  • the surface increasing structure can be arranged at either side or even at both sides.
  • the wall seg ⁇ ment of the at least two wall segments that is embodied with the at least one surface increasing structure is positioned at the pressure side of the at least one hollow aerofoil.
  • the surface increasing structure is a part of or formed in the pressure side of the aerofoil or its inner sur- face, respectively.
  • the at least one surface increasing struc ⁇ ture has a rear end.
  • the trailing edge has a termination rim.
  • rim should be also understood as edge or border.
  • the rear end of the surface increas ⁇ ing structure terminates flush with the trailing edge termi ⁇ nation rim.
  • the length of the surface increasing struc ⁇ ture is basically equal to the depth of the exit aperture.
  • the cooling passage has a cross section and two cooling pas ⁇ sages have a span wise distance towards each other. This span wise distance is basically the cross section of the cooling passage. Furthermore, two surface increasing structures have a span wise distance towards each other that is basically the cross section of the cooling passage. In other words, the spacing between two surface increasing structures is the same as or equal to the spacing between two cooling passages.
  • the aerofoil is a tur- bine blade or turbine vane, and especially a turbine blade.
  • the matrix component comprises at least one cavity, wherein dimensions of the at least one cavity are embodied in such a way so that during the casting process with the at least one casting material at least one surface increasing structure of the turbine assembly is formed.
  • the surface increasing structure can be formed reliably and easily during the manufacturing or casting process, respectively, of the turbine assembly, what further ad ⁇ vantageously economising the manufacturing process. Moreover, smooth junctions between the surface increasing structure and the wall segment can be provided.
  • the term "matrix” should be understood as tem ⁇ plate or blueprint.
  • the casting material may be any material feasible for a person skilled in the art to form the shape and structures of the turbine assembly, like a metal, a blended metal mixture or an alloy etc.
  • the matrix component may be manufactured or built of any material feasible for a person skilled in the art that can be removed after the cast ⁇ ing process and may for example be a ceramic material.
  • the matric component may be named ceramic core.
  • the invention further provides a method for manufacturing the inventive turbine assembly by using the inventive matrix com ⁇ ponent . It is proposed that the method comprises at least the steps of: Casting a casting material around the matrix component and removing the matrix component and thus obtaining the turbine assembly. Due to this manufacturing process the surface increasing structure can be integrated in the turbine assembly effec ⁇ tively and reliably. Furthermore, an additional manufacturing step for integration of the surface increasing structure can be omitted saving time and costs.
  • FIG 1 shows a schematically and sectional view of a gas turbine engine comprising several inventive turbine assemblies ,
  • FIG 2 shows a perspective view of one turbine assembly from FIG 1 with an aerofoil comprising a trailing edge with an exit aperture
  • FIG 3 shows two wall segments and the exit aperture of the aerofoil from FIG 2 with surface increasing structures at one wall segment
  • FIG 4 shows one surface increasing structure from FIG 3 in an enlarged view
  • FIG 5 shows a matrix component for a casting process for the turbine assembly from FIG 2 with cavities cor ⁇ responding to the surface increasing structures from FIG 4
  • FIG 6 shows a trailing edge region of the matrix compo ⁇ nent from FIG 5 with cavities in an enlarged view.
  • upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine 68 unless otherwise stated. If used and not otherwise stated, the terms axial, radial and circumferential are made with reference to a rotational axis 78 of the engine 68.
  • FIG 1 shows an example of a gas turbine engine 68 in a sec- tional view.
  • the gas turbine engine 68 comprises, in flow se ⁇ ries, an inlet 70, a compressor section 72, a combustion section 74 and a turbine section 76, which are generally ar ⁇ ranged in flow series and generally in the direction of a longitudinal or rotational axis 78.
  • the gas turbine engine 68 further comprises a shaft 80 which is rotatable about the ro ⁇ tational axis 78 and which extends longitudinally through the gas turbine engine 68.
  • the shaft 80 drivingly connects the turbine section 76 to the compressor section 72.
  • air 82 which is taken in through the air inlet 70 is compressed by the com ⁇ pressor section 72 and delivered to the combustion section or burner section 74.
  • the burner section 74 comprises a burner plenum 84, one or more combustion chambers 86 defined by a double wall can 88 (not shown in detail) and at least one burner 90 fixed to each combustion chamber 86.
  • the combustion chamber (s) 86 and the burner (s) 90 are located inside the burner plenum 84.
  • the compressed air passing through the compressor section 72 enters a diffuser 92 and is discharged from the diffuser 92 into the burner plenum 84 from where a portion of the air enters the burner 90 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 94 or working gas from the combustion is channelled via a transition duct 96 to the turbine section 74.
  • the turbine section 76 comprises a number of blade carrying production discs 98 or turbine wheels attached to the shaft 80.
  • the turbine section 76 comprises two discs 98 each carry an annular array of turbine assem ⁇ blies 10 which each comprises a basically hollow aerofoil 12 (see FIG 2) embodied as a turbine blade.
  • the number of blade carrying production discs 98 could be different, i.e. only one production disc 98 or more than two production discs 98.
  • turbine assemblies 10 embodied as aerofoils 12 from turbine cascades 100 are disposed between the turbine blades.
  • Each turbine cascade 100 carries an annu ⁇ lar array of aerofoils 12 in the form of guiding vanes, which are fixed to a stator 102 of the gas turbine engine 68. Be ⁇ tween the exit of the combustion chamber 86 and the leading turbine blades inlet guiding vanes or nozzle guide vanes 104 are provided. The combustion gas 94 from the combustion chamber 86 enters the turbine section 76 and drives the turbine blades which in turn rotate the shaft 80. The guiding vanes 104 serve to op ⁇ timise the angle of the combustion or working gas 94 on to the turbine blades.
  • the compressor section 72 comprises an axial series of guide vane stages 106 and rotor blade stages 108 with turbine assemblies 10 or turbine blades or vanes, respectively .
  • FIG 2 shows in a perspective view a turbine assembly 10 or turbine blade of the gas turbine engine 68 with an aerofoil 12.
  • the turbine assembly 10 comprises a basically hollow aerofoil 12, with two cooling regions, specifically, a channelled cooling region and a fin-pin/pedestal cooling region (not shown in detail) .
  • the former is located at a leading edge 110 and comprises a meandering cooling channel 112 for a cooling medium 34 (see FIG 5) that extend in span wise direction 26 of the aerofoil 12 and different parts of the cooling channel 12 are separated by ribs 114.
  • the latter is located at a trailing edge 14 of the aerofoil 12.
  • the cooling channel 112 is spliced at the trailing edge 14 in several cooling pas ⁇ sages 32, 32' (see FIG 5) which were also separate by ribs 114.
  • the cooling channel 112 and the cooling passages 32, 32' may be in flow communication with each other or with other cooling features of the aerofoil 12, like film cooling holes, impingement devices or the like (not specified or shown) .
  • the turbine assembly 10 further comprises a platform 116 and a root portion 118, wherein the platform 116 is arranged in span wise direction 26 between the aerofoil 12 and the root portion 118.
  • the aerofoil 12 may comprise, if embod ⁇ ied as guide vane, an outer platform, embodied as a shroud, at its tip, which is not shown in FIG 2.
  • the trailing edge 14 comprises two wall segments 16, 18, wherein the wall segment 16 is a part of a pressure side 52 and the wall segment 18 is a part of a suction side 120 of the aerofoil 12.
  • the wall segments 16, 18 extend in a span wise direction 26 of the aerofoil 12.
  • the wall segments 16, 18 each comprise an inner surface 20, 22 (see FIG 3 and 4) that face each other and that enclose an exit aperture 24, wherein the cooling medium 34 is discharged from the cooling passages 32, 32' into the exit aperture 24 and subsequently in the surrounding environment, like the gas path in the compressor section 72 or the turbine section 76.
  • the exit aperture 24 extends in span wise direction 26 of the aerofoil 12 basically all along a span 30 of the aerofoil 12.
  • the exit aperture 24 has a depth D in a direc ⁇ tion 42 basically perpendicular to a span wise direction 26 or the direction 42 from the leading edge 110 to the trailing edge 52 of the aerofoil 12.
  • the inner surface 20 of the wall segment 16 of the pressure side 52 of the aerofoil 12 comprises several, in this exemplary embodiment six, surface increasing structures 28, 28'. These surface increasing structures 28, 28' extend partially into the exit aperture 24 and specifically they originate from the inner surface 20 of the pressure side 52 (wall segment 16) to the opposed inner surface 22 of the suc ⁇ tion side 120 (wall segment 18) .
  • FIG 5 shows a matrix component 62 for the manufacturing of the turbine assembly 10 and FIG 6 an enlarged part of a trailing edge region of the matrix component 62.
  • the matrix component 62 is used as a tem ⁇ plate to form the innards of the aerofoil 12, like the cool ⁇ ing passages 32, 32', the dividing ribs 114, the exit aper- ture 24 or the surface increasing structures 28, 28' etc.
  • Structures which are shown in FIG 5 and 6 in solid form are structures that will be holes, openings or empty apertures in the resulting aerofoil 12. Vacant spaces, in turn, will be solid structures in the resulting aerofoil 12 or turbine as- sembly 10, respectively.
  • the reference numerals of the end structures (12, 16, 20, 24, 28, 28', 32, 32', 114) according to FIG 1 to 4 are used in FIG 5 and 6 even if they are not actually shown.
  • the method for manufacturing the turbine assembly by using the matrix component 62 comprises at least the steps of:
  • the matrix component 62 comprises cavities 66, 66' (note that the cavities are also marked with the reference numerals for the corresponding surface increasing structures 28, 28' in FIG 5 and 6) which dimensions are embodied in such a way so that during the casting process with the casting material 64 the surface increasing structures 28, 28' of the turbine as- sembly 10 are formed.
  • the surface increasing structures 28, 28' and the wall segment 16 are embodied inte ⁇ grally with each other.
  • the features of the surface increasing structures will be de ⁇ scribed in reference to FIG 3 to 6. Even if the surface in ⁇ creasing structures 28, 28' as well as the cooling passages 32, 32' and the exit aperture 24 are only indirectly shown through their corresponding structures in the matrix component 62 of FIG 5 and 6 some features will be described in reference to these FIG because these FIG depict the features more clearly than FIG 3 and 4.
  • the surface increasing structures 28, 28' extend basically in parallel towards each other and are arranged equally spaced towards each other (see FIG 6) . As can be seen in FIG 3 and 6 each of the cooling passages
  • each 32, 32' of the plurality of cooling passage 32, 32' comprises one surface increasing structure 28, 28' of the plurality of surface increasing structures 28, 28'. Further, each surface increasing structure 28, 28 has a longitudinal extension 44 and each cooling passage 32, 32' has a longitudinal extension 50. The longitudinal extension 44 of each surface increasing structure 28, 28' and the longitudinal extension 50 of each cooling passage 32, 32' are arranged aligned with each other or are arranged coaxial towards each other.
  • each cooling passage 32, 32' has a cross section 56 and two cooling passages 32, 32' have a span wise distance 58 towards each other.
  • the span wise distance 58 between two cooling passages 32, 32' is basically the cross section 56 of one cooling passage 32, 32'.
  • two surface in ⁇ creasing structures 28, 28' have a span wise distance 60 to ⁇ wards each other that is basically the cross section 56 and thus the span wise distance 58 of the (two) cooling pas ⁇ sage ( s ) 32 , 32 ' .
  • each surface increas- ing structure 28, 28' has a height h that is maximally 50% of the width W of the exit aperture 24.
  • a cross sectional area 38 of each surface increasing structure 28, 28' is maximally 50% of a cross sectional area 36 of each cooling passage 32, 32' (see FIG 3 and 6) .
  • a width w of the surface in ⁇ creasing structure 28 is about 0.5% of the span 30 of the aerofoil 12.
  • a height h of the surface increasing structure 28 is about 0.4% of the span 30 and a length 1 of the surface increasing structure 28 is about 2% of the span 30.
  • the width w is greater than 0.4 mm
  • the height h is greater than 0.3 mm
  • the length 1 is greater than 1.4 mm.
  • an effective surface 122 of the surface increas ⁇ ing structure 28 is 1.587 mm 2 (see dotted area in FIG 4) .
  • Each cooling passage 32, 32' has an exit opening 40 to dis- charge the cooling medium 34 into the exit aperture 24 (see FIG 6) .
  • the longitudinal extension 44 of each surface in ⁇ creasing structure 28, 28' has a forward end 46 (towards the leading edge 110) and a rear end 48 (towards the trailing edge 52) .
  • the surface increasing structure 28, 28' extends from its forward end 46 to its rear end 48 along the depth D of the exit aperture 24.
  • the forward end 46 of each surface increasing structure 28, 28' is arranged in the exit opening 40 of the corresponding cooling passage 32, 32' (see also FIG 6) .
  • each surface increasing structure 28, 28' terminates flush with a termination rim 54 of the trailing edge 14 (see FIG 6) .
  • the term "comprising” does not ex ⁇ clude other elements or steps and "a” or “an” does not ex ⁇ clude a plurality. Also elements described in association with different embodiments may be combined. It should also be noted that reference signs in the claims should not be con ⁇ strued as limiting the scope of the claims.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

La présente invention concerne un ensemble turbine (10), lequel ensemble comprend au moins un profil aérodynamique fondamentalement creux (12) avec au moins un bord de fuite (14) comprenant au moins deux segments de paroi (16, 18) qui comprennent chacun une surface interne (20, 22), les surfaces internes (20, 22) se faisant mutuellement face, et les surfaces internes (20, 22) renfermant au moins une ouverture de sortie (24) s'étendant fondamentalement dans une direction d'étendue (26) du ou des profils aérodynamiques fondamentalement creux (12), au moins l'une des surfaces internes (20, 22) des au moins deux segments de paroi (16, 18) du bord de fuite (14) comprenant au moins une structure d'accroissement de surface (28, 28') s'étendant partiellement dans la ou les ouvertures de sortie (24) à partir d'une surface interne (20, 22) jusqu'à l'autre surface interne (20, 22), l'ensemble turbine comprenant en outre au moins un passage de refroidissement (32, 32') pour un milieu de refroidissement (34) évacué dans la ou les ouvertures de sortie (24), le ou les passages de refroidissement (32, 32') ayant une ouverture de sortie (40) et la ou les ouvertures de sortie (24) ayant une profondeur (D) dans une direction (42) fondamentalement perpendiculaire à une direction d'étendue (26) du profil aérodynamique fondamentalement creux (12), la ou les structures d'accroissement de surface (28, 28') ayant une extension longitudinale (44) avec une extrémité avant (46) et une extrémité arrière (48), la ou les structures d'accroissement de surface (28, 28') s'étendant à partir de son extrémité avant (46) jusqu'à son extrémité arrière (48) le long de la profondeur (D) de la ou des ouvertures de sortie (24), et l'extrémité avant (46) de la ou les structures d'accroissement de surface (28, 28') étant disposée dans l'ouverture de sortie (40) du ou des passages de refroidissement ( 32, 32'). Ceci produit un transfert de chaleur, et, par conséquent, un refroidissement, suffisants, de l'ensemble turbine (10).
PCT/EP2016/062581 2015-06-23 2016-06-03 Ensemble turbine Ceased WO2016206944A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP15173387.0 2015-06-23
EP15173387.0A EP3109404A1 (fr) 2015-06-23 2015-06-23 Ensemble de turbine

Publications (1)

Publication Number Publication Date
WO2016206944A1 true WO2016206944A1 (fr) 2016-12-29

Family

ID=53476784

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2016/062581 Ceased WO2016206944A1 (fr) 2015-06-23 2016-06-03 Ensemble turbine

Country Status (2)

Country Link
EP (1) EP3109404A1 (fr)
WO (1) WO2016206944A1 (fr)

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1895098A2 (fr) * 2006-08-30 2008-03-05 Honeywell International Inc. Pale de turbine refroidie améliorée à haute effectivité
US20100074763A1 (en) * 2008-09-25 2010-03-25 Siemens Energy, Inc. Trailing Edge Cooling Slot Configuration for a Turbine Airfoil
US20130302179A1 (en) * 2012-05-09 2013-11-14 Robert Frederick Bergholz, JR. Turbine airfoil trailing edge cooling hole plug and slot
EP2787173A1 (fr) * 2011-11-30 2014-10-08 IHI Corporation Pale de turbine

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1895098A2 (fr) * 2006-08-30 2008-03-05 Honeywell International Inc. Pale de turbine refroidie améliorée à haute effectivité
US20100074763A1 (en) * 2008-09-25 2010-03-25 Siemens Energy, Inc. Trailing Edge Cooling Slot Configuration for a Turbine Airfoil
EP2787173A1 (fr) * 2011-11-30 2014-10-08 IHI Corporation Pale de turbine
US20130302179A1 (en) * 2012-05-09 2013-11-14 Robert Frederick Bergholz, JR. Turbine airfoil trailing edge cooling hole plug and slot

Also Published As

Publication number Publication date
EP3109404A1 (fr) 2016-12-28

Similar Documents

Publication Publication Date Title
JP4311919B2 (ja) ガスタービンエンジン用のタービン翼形部
JP5503140B2 (ja) 発散型タービンノズル
US10815789B2 (en) Impingement holes for a turbine engine component
US10260354B2 (en) Airfoil trailing edge cooling
US12215602B2 (en) Component for a turbine engine with a cooling hole
JP2012102726A (ja) タービンロータブレードのプラットフォーム領域を冷却するための装置、システム、及び方法
US20170211393A1 (en) Gas turbine aerofoil trailing edge
KR20210103391A (ko) 에어포일에서 충돌 공기를 재사용하기 위한 충돌 인서트, 충돌 인서트를 포함하는 에어포일, 터보머신 구성요소, 및 이를 포함하는 가스 터빈
CN204312137U (zh) 涡轮发动机及涡轮机喷嘴轮叶段
JP2017145826A (ja) ガスタービンエンジン後縁排出穴
US11549377B2 (en) Airfoil with cooling hole
US10760431B2 (en) Component for a turbine engine with a cooling hole
EP3647544B1 (fr) Profil d'aube statorique refroidie d'une turbine à gaz
KR20210103392A (ko) 에어포일에서 충돌 공기를 재사용하기 위한 삼중 벽 충돌 인서트, 충돌 인서트를 포함하는 에어포일, 터보머신 구성요소, 및 이를 포함하는 가스 터빈
US10724391B2 (en) Engine component with flow enhancer
CN107084002B (zh) 一种涡轮机组件和制备方法
RU2688124C2 (ru) Способ для изготовления узла турбины
CN108506048A (zh) 用于涡轮发动机的膜孔布置
US20170234145A1 (en) Accelerator insert for a gas turbine engine airfoil
EP3241991A1 (fr) Ensemble de turbine
WO2016206944A1 (fr) Ensemble turbine
US11788417B2 (en) Turbine blade and gas turbine
EP3875735A1 (fr) Surface portante pour turbine à gaz

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 16728886

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

122 Ep: pct application non-entry in european phase

Ref document number: 16728886

Country of ref document: EP

Kind code of ref document: A1