US20140234088A1 - Modular blade or vane for a gas turbine and gas turbine with such a blade or vane - Google Patents
Modular blade or vane for a gas turbine and gas turbine with such a blade or vane Download PDFInfo
- Publication number
- US20140234088A1 US20140234088A1 US14/012,101 US201314012101A US2014234088A1 US 20140234088 A1 US20140234088 A1 US 20140234088A1 US 201314012101 A US201314012101 A US 201314012101A US 2014234088 A1 US2014234088 A1 US 2014234088A1
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- United States
- Prior art keywords
- shell
- carrying structure
- airfoil
- blade
- vane
- 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.)
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Classifications
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- 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
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
- F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
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- 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
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
-
- 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
- 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
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
-
- 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
- F01D5/34—Rotor-blade aggregates of unitary construction, e.g. formed of sheet laminae
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- 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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
-
- 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
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- 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
- F05D2230/50—Building or constructing in particular ways
- F05D2230/51—Building or constructing in particular ways in a modular way, e.g. using several identical or complementary parts or features
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- 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
-
- 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
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- 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
- F05D2240/80—Platforms for stationary or moving blades
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the present invention relates to the technology of gas turbines. It refers to a modular blade or vane according to the preamble of claim 1 . It further refers to a gas turbine with such a blade or vane.
- a near-wall cooling scheme for an airfoil of a gas turbine engine discloses the document U.S. Pat. No. 5,720,431.
- the disclosed airfoil includes a double-wall configuration in the mid-chord region thereof with a plurality of radial feed passages defined on each side of the airfoil between an inner wall and an outer wall.
- a central radially extending feed chamber is defined between the two inner walls.
- the trailing edge of the airfoil includes a conventional single wall configuration with two outer walls defining a sequence of trailing edge passages there between.
- this airfoil provides stress conditions in the interface area between the double wall configuration in the trailing edge.
- the experience in this field teaches that the yield rate of such near-wall cooling schemes is not commercially viable.
- FIG. 1 a which is a replica of FIG. 4 of that document, shows a modular turbine blade 10 , where additional parts made of different materials are joined with a central airfoil 11 , said additional parts comprising a leading edge 13 , a trailing edge 15 and a blade tip 12 .
- a buffer layer 14 may be provided between the leading edge part 13 and the central airfoil 11 for stress reduction between components with different thermal expansion coefficient.
- Document EP 2 189 626 A1 discloses another modular rotor blade arrangement 20 , especially for a gas turbine, which can be fastened on a blade carrier and includes in each case a blade airfoil element 21 and a platform element 22 , wherein the platform elements of a blade row form a continuous inner shroud.
- a mechanical decoupling which extends the service life, is achieved by the blade airfoil element 21 and the platform element 22 being formed as separate elements and by being able to be fastened in each case separately on the blade carrier. According to one configuration of this arrangement (see FIG.
- the blade arrangement 20 comprises an aerodynamically effective airfoil 21 , a shank 23 which adjoins the airfoil 21 at the bottom and is shrouded by the platform element 22 , and a blade root 24 which adjoins the shank 23 at the bottom, wherein the blade root 24 is provided for fastening the airfoil element 21 on the blade carrier.
- the blade airfoil element 21 may be constructed of different sections consisting of different materials. For example, the leading edge and the trailing edge consist of a material which is different to that of the remaining airfoil.
- the platform element 22 has a through-opening 25 through which the airfoil element 21 extends.
- the sections which consist of different materials extend downwards into the region of the blade airfoil element which is shrouded by the platform element 22 .
- the transition between regions of different material is then not exposed to the extreme temperature conditions which prevail directly in the region of the hot gas path.
- the stress level in the platform-airfoil transition region is significantly lowered.
- One aspect of the invention provides a modular blade or vane for a gas turbine, which essentially comprises the modular components of:
- a platform element with a planar or contoured surface defining a platform level and a through-opening therein
- the airfoil having:
- a carrying structure extending along a longitudinal axis of the airfoil, having a root portion for fastening on the blade or vane carrier of the gas turbine, having a tip portion and having at least one interior passage, extending from the root portion to the tip portion of the airfoil;
- a first joint in a region below the platform level of the platform element integrally joins the shell to said carrying structure
- At least one additional joint between shell and carrying structure is a form-fit joint, allowing relative movement in longitudinal direction between shell and carrying structure.
- the first joint between shell and carrying structure is a welded joint, a brazed joint or a retainer joint.
- this first joint is gastight.
- said at least one additional joint between shell and carrying structure comprises an edge of the shell fitting into a groove, formed in a component at the tip of the airfoil.
- the groove is formed in a shroud element or a tip cap.
- said additional joint allows leakage of a cooling medium from the gap into the hot gas path.
- the shell is laterally fixed to the carrying structure by a number of complementary form-fitting elements on the outer surface of the carrying structure and on the inner surface of the shell between root and tip.
- these form-fitting elements are designed as dovetail joint.
- this joint design allows relative movement between shell and carrying structure in longitudinal direction to compensate thermal expansion and—in case of rotating blades—expansion caused by centrifugal force. And on the other hand, this joint design avoids shell deformation in lateral direction, e.g. buckling inwards or outwards.
- a cooling medium preferably cooling air
- This at least one passage is extending from the root portion of the airfoil to its tip and along this passage a number of feed holes for the cooling air is arranged.
- the air from the interior passage flows through said holes into the gap, defined between the outer surface of the carrying structure and the inner surface of the aerodynamically shaped shell to cool the shell from inside.
- the air leaving the feed holes, impinges onto the inner surface of the shell for effective cooling.
- the cooling air flows along the gap towards the tip of the airfoil and leaves the gap at the tip into the hot gas path.
- the exterior surface of the carrying structure and/or the inner surface of the shell are equipped with turbulators, for example ribs or pedestals, for enhancing heat transfer.
- individual turbulators are designed to provide support to the shell and thus increase the structural integrity of the shell.
- the manufacturing of a blade or vane according to the present invention is less complex and more efficient.
- the load carrying structure has a simple straight core being easily cast.
- the cooling features at its outer surface can be either cast or easily machined. This leads to significant cost savings.
- the cooling medium delivered with low pressure loss through the straight interior passage, flows out through feed-holes into the gap, defined between load carrying structure and shell, and passes the gap with a high yield rate of convective cooling.
- the feed-holes can serve to impinge the inner wall of the shell thus additionally increasing the cooling effect with the result of an improved lifetime of the components and/or a reduced amount of cooling medium.
- a reduced use of cooling medium results in higher engine performance.
- FIG. 1 a,b show modular gas turbine blades known from the prior art (WO 2011/058043 and EP 2189626);
- FIG. 2 shows a sectional top view of a modular turbine blade or vane design suitable for the present invention
- FIG. 3 shows a sectional side view of a modular blade according to an embodiment of the invention
- FIG. 4 a,b,c show in detail tip fixation variants of a shell
- FIG. 5 shows a sectional side view of a detail of a modular vane according to the invention
- FIG. 6 a,b show a sectional side view of another detail of a modular vane according to the invention.
- FIG. 2 shows a sectional top view of a modular gas turbine blade or vane design in accordance with the present invention.
- the airfoil 32 of a blade 30 or vane 31 comprises a load carrying structure 33 , extending in longitudinal direction from the root portion 35 to the tip 36 , and an aerodynamically shaped shell 34 , extending in a distance 37 over the outer surface 38 of the carrying structure 33 and defining the outer contour 40 of the airfoil 32 .
- the carrying structure 33 may be formed as a single unit or may be assembled from separate portions.
- the airfoil 32 is sealed by a cap 41 .
- the shell 34 defines a pressure side 42 , a suction side 43 , a leading edge 44 and a trailing edge 45 .
- the shell 34 may be made from a different material compared to the carrying structure 33 depending on the specification of the gas turbine.
- the shell 34 can be made of a single unit or can be assembled from more than one single parts, e.g. two half-shells attached to the carrying structure 33 .
- the carrying structure 33 comprises at least one inner passage 46 for conducting a flow of cooling air 49 from a reservoir in the blade or vane carrier towards the airfoil tip 36 .
- the wall of the carrying structure 33 is equipped with a number of holes 47 connecting the at least one inner passage 46 with the gap 48 defined between the outer surface 38 of the carrying structure 33 and the inner surface 39 of the shell 34 .
- the cooling air 49 flows out of the holes 47 into the gap 48 .
- this air 49 serves to impinge onto the inner surface 39 of the shell 34 and to effectively cool it. Before the air 49 exits the gap 48 through exit holes 50 in the shell 34 , it effects convective cooling of the shell 34 when flowing along the gap 48 towards the exit holes 50 .
- exit holes 50 can include film cooling holes at the pressure side 42 and/or the suction side 43 and/or the leading edge 44 and/or the trailing edge 45 , or these exit holes 50 can be a clearance between an outer or inner edge 51 , 52 of the shell 34 and the carrying structure 33 , as shown in more detail in FIGS. 4 and 6 .
- FIG. 3 shows a sectional side view of a blade 30 according to the invention.
- the load carrying structure 33 forms a fir tree portion 54 that fits into a complementary fir tree groove of a blade carrier of the turbomachine (not shown).
- the platform 53 is formed separately and joined to the load carrying structure 33 by means of welding, brazing or a retainer connection as described e.g. in U.S. Pat. No. 5,797,725.
- the shell 34 extends longitudinally parallel to axis 55 from a section under the platform level 56 to the airfoil tip 36 .
- Under the platform level 56 the shell 34 is integrally joined with the load carrying structure 33 by an appropriate joining method, such as welding, brazing etc. Thus a fixed and gas-tight joint is achieved in this section.
- shell 34 and load carrying structure 33 are connected in a way that allows compensation of thermal expansion. Details of preferred embodiments for tip fixation are shown in FIGS. 4 a - c.
- shell 34 and load carrying structure 33 are equipped with positive locking elements to laterally support the shell 34 .
- a dovetail connection 57 supports the shell 34 and prevents its lateral deformation, but allows relative movement along the longitudinal axis 55 to compensate thermal expansion and expansion caused by centrifugal forces.
- Exterior surface 38 of the load carrying structure 33 and/or interior surface 39 of the shell 34 are equipped with turbulators 58 , such as ribs, for increasing the heat transfer between the shell ( 34 ) and the cooling medium 49 .
- the height of individual turbulators 58 may correspond to the width 37 of the gap 48 . In doing so, these turbulators 58 act as a mechanical stop (distance holder) and prevent inward deformation of the shell 34 , particularly at the pressure side 42 of the airfoil 32 .
- FIG. 4 a, b, c show details of preferred embodiments for shell 34 fixation at the airfoil tip 36 allowing relative movement between the shell 34 and the load carrying structure 33 .
- the exposure of the individual components of the modular blade or vane to thermal expansion and—in case of rotating blades—to centrifugal forces requires a fixation of the shell 37 with clearance in longitudinal direction at one end.
- the shell 34 ends in a groove 59 at the lower side of a tip component 41 .
- a component 41 may be, for example, a tip cap or a shroud element.
- Depth and width of the groove 59 are sized so that a leakage of cooling medium 49 is possible.
- the contour of the groove 49 may be machined by any appropriate machining method.
- FIG. 4 b An alternative embodiment of fixation of shell 34 at the airfoil tip 36 is shown in FIG. 4 b .
- the load carrying structure 33 broadens to the outer contour of the airfoil 32 thereby overlapping the upper edge 51 of the shell 34 .
- a shoulder 60 is machined so that the upper edge 51 fits into this shoulder 60 .
- a surrounding weld seam 61 at the outer edge of the shoulder 60 locks the shell 34 in its position.
- FIG. 4 c shows another variant of a tip fixation.
- the shoulder 60 at the lower side of a tip component 41 e.g. the shroud element, is limited outwards by a stop bar 62 .
- FIGS. 5 and 6 show sectional side views of a vane 31 according to the invention.
- FIG. 5 shows an example for an outer diameter fixation of the shell 34 , the carrying structure 33 , and the platform 53
- FIG. 6 a and b disclose two arrangement variants of the carrying structure 33 , the shell 34 and the platform 53 at the inner diameter of the vane 31 .
- the carrying structure 33 broadens towards the end of the airfoil 32 , thereby overlapping the outer edge 51 of the shell 34 .
- a shoulder 60 is machined so that the edge 51 of the shell 34 and a portion of its lateral surface 64 bear against the shoulder 60 .
- the shell 34 is integrally joined to the carrying structure 33 .
- the carrying structure 33 and shell 34 are arranged in a distance 37 , forming the longitudinal gap 48 .
- a number of cooling holes 47 delivers the cooling medium 49 from the inner passage 46 inside the carrying structure 33 into the gap 48 .
- Turbulators 58 on the outer surface 38 of the load carrying structure 33 support the turbulent flow of the cooling medium 49 , thus enhancing convective heat transfer. Individual turbulators 58 may serve as distance holders, as mentioned in connection with FIG. 3 .
- the overlapping part 63 of the carrying structure 33 comprises an outer lateral surface 65 .
- the platform 53 comprises a through-opening 66 with an inner surface 67 .
- Lateral surface 65 of the carrying structure 33 and inner surface 67 of the through-opening 66 are complementary designed.
- a suitable method e.g. bi-metal-casting, welding brazing etc. the platform 53 is connected to the airfoil 32 .
- the inner edge 52 of shell 34 fits into the shoulder 60 of carrying structure 33 .
- the platform 53 is connected to the airfoil 32 .
- the size of platform 53 exceeds the size of the overlapping part 63 of the carrying structure 33 thus forming a groove in which the shell 34 ends.
- An optional seal 68 between airfoil shell 34 and platform 53 prevents the penetration of hot gases from the hot gas path into this groove.
- the overlapping part 63 of the carrying structure 33 is broader, compared to the example of FIG. 6 a , and the inner edge 52 of the airfoil shell 34 ends in a groove 59 , machined into the overlapping part 63 .
- a seal 68 may be arranged to prevent the penetration of hot gases into the groove 59 .
- the platform 53 is connected to the outer lateral surface 65 of the carrying structure 33 by any suitable method.
Abstract
Description
- This application claims priority to European Application 12182327.2 filed Aug. 30, 2012, the contents of which are hereby incorporated in its entirety.
- The present invention relates to the technology of gas turbines. It refers to a modular blade or vane according to the preamble of
claim 1. It further refers to a gas turbine with such a blade or vane. - Ever increasing hot gas temperatures in gas turbines require the use of special materials and/or designs (e.g. cooling schemes) to prevent excessive usage of cooling air. The usage of special materials and/or designs is partly done by means of inlays and/or inserts to the main structure of the part.
- A near-wall cooling scheme for an airfoil of a gas turbine engine discloses the document U.S. Pat. No. 5,720,431. The disclosed airfoil includes a double-wall configuration in the mid-chord region thereof with a plurality of radial feed passages defined on each side of the airfoil between an inner wall and an outer wall. A central radially extending feed chamber is defined between the two inner walls. The trailing edge of the airfoil includes a conventional single wall configuration with two outer walls defining a sequence of trailing edge passages there between.
- Disadvantageously this airfoil provides stress conditions in the interface area between the double wall configuration in the trailing edge. In addition, the experience in this field teaches that the yield rate of such near-wall cooling schemes is not commercially viable.
- Document WO 2011 058 043 A1 proposes to use inserts (special material and/or design) in parent metal to cope with the special requirements in the hot gas path of a gas turbine.
FIG. 1 a, which is a replica ofFIG. 4 of that document, shows amodular turbine blade 10, where additional parts made of different materials are joined with acentral airfoil 11, said additional parts comprising a leadingedge 13, atrailing edge 15 and ablade tip 12. Abuffer layer 14 may be provided between the leadingedge part 13 and thecentral airfoil 11 for stress reduction between components with different thermal expansion coefficient. - As a result, particularly loaded regions of the blade airfoil can be differently designed with regard to materials than the remaining regions. The
central airfoil 11 and theedges platform 16, which borders the hot gas channel and protects theblade root 17 below. - However, this known solution is based on a design where the airfoil with its root and the platform are and remain separate elements with separate mounting means. The discontinuity which is associated with the transition between the regions of different materials is exposed to the extreme temperature conditions which prevail in the region of the blade airfoil.
- Document EP 2 189 626 A1 discloses another modular
rotor blade arrangement 20, especially for a gas turbine, which can be fastened on a blade carrier and includes in each case ablade airfoil element 21 and aplatform element 22, wherein the platform elements of a blade row form a continuous inner shroud. With such a blade arrangement, a mechanical decoupling, which extends the service life, is achieved by theblade airfoil element 21 and theplatform element 22 being formed as separate elements and by being able to be fastened in each case separately on the blade carrier. According to one configuration of this arrangement (seeFIG. 1 b) theblade arrangement 20 comprises an aerodynamicallyeffective airfoil 21, ashank 23 which adjoins theairfoil 21 at the bottom and is shrouded by theplatform element 22, and ablade root 24 which adjoins theshank 23 at the bottom, wherein theblade root 24 is provided for fastening theairfoil element 21 on the blade carrier. Theblade airfoil element 21 may be constructed of different sections consisting of different materials. For example, the leading edge and the trailing edge consist of a material which is different to that of the remaining airfoil. - The
platform element 22 has a through-opening 25 through which theairfoil element 21 extends. By this means the sections which consist of different materials extend downwards into the region of the blade airfoil element which is shrouded by theplatform element 22. The transition between regions of different material is then not exposed to the extreme temperature conditions which prevail directly in the region of the hot gas path. The stress level in the platform-airfoil transition region is significantly lowered. - It is an object of the present invention to provide a modular gas turbine blade or vane, wherein the individual components are arranged in a way to improve the mechanical stability and integrity of the blade or vane and in addition this invention should enable an improved internal cooling of the blade or vane with a lesser amount of cooling medium.
- It is another object of the invention to provide a gas turbine with such a blade or vane.
- This and other objects are obtained by means of the subject matter of the independent claims. Advantageous embodiments are given in the dependent claims.
- One aspect of the invention provides a modular blade or vane for a gas turbine, which essentially comprises the modular components of:
- a platform element with a planar or contoured surface defining a platform level and a through-opening therein,
- an airfoil, extending through the platform element,
- the airfoil having:
- a carrying structure extending along a longitudinal axis of the airfoil, having a root portion for fastening on the blade or vane carrier of the gas turbine, having a tip portion and having at least one interior passage, extending from the root portion to the tip portion of the airfoil;
- an aerodynamically shaped shell, extending in a distance over the carrying structure and defining the outer surface of the airfoil,
- a longitudinally extending gap between the carrying structure and the shell,
- a number of through-holes in the carrying structure for directing a cooling medium from the interior passage into the gap, wherein
- a first joint in a region below the platform level of the platform element integrally joins the shell to said carrying structure, and
- at least one additional joint between shell and carrying structure is a form-fit joint, allowing relative movement in longitudinal direction between shell and carrying structure.
- According to a first preferred aspect of the invention the first joint between shell and carrying structure is a welded joint, a brazed joint or a retainer joint.
- According to an additional aspect this first joint is gastight.
- According to another aspect of the invention said at least one additional joint between shell and carrying structure comprises an edge of the shell fitting into a groove, formed in a component at the tip of the airfoil.
- Preferably, the groove is formed in a shroud element or a tip cap.
- In a preferred embodiment said additional joint allows leakage of a cooling medium from the gap into the hot gas path.
- According to a further aspect of the invention the shell is laterally fixed to the carrying structure by a number of complementary form-fitting elements on the outer surface of the carrying structure and on the inner surface of the shell between root and tip. Preferably these form-fitting elements are designed as dovetail joint.
- On the one hand, this joint design allows relative movement between shell and carrying structure in longitudinal direction to compensate thermal expansion and—in case of rotating blades—expansion caused by centrifugal force. And on the other hand, this joint design avoids shell deformation in lateral direction, e.g. buckling inwards or outwards.
- A cooling medium, preferably cooling air, is admitted through an inlet at the root of the carrying structure into the one or more cooling air passages inside the carrying structure. This at least one passage is extending from the root portion of the airfoil to its tip and along this passage a number of feed holes for the cooling air is arranged. The air from the interior passage flows through said holes into the gap, defined between the outer surface of the carrying structure and the inner surface of the aerodynamically shaped shell to cool the shell from inside. Preferably the air, leaving the feed holes, impinges onto the inner surface of the shell for effective cooling.
- According to an additional aspect of the invention the cooling air flows along the gap towards the tip of the airfoil and leaves the gap at the tip into the hot gas path.
- In those areas of the gap, applied with cooling air, the exterior surface of the carrying structure and/or the inner surface of the shell are equipped with turbulators, for example ribs or pedestals, for enhancing heat transfer.
- In addition, individual turbulators are designed to provide support to the shell and thus increase the structural integrity of the shell.
- By means of this invention a modular blade or vane arrangement is created which has a couple of advantages compared to prior art solutions.
- The manufacturing of a blade or vane according to the present invention is less complex and more efficient. The load carrying structure has a simple straight core being easily cast. The cooling features at its outer surface can be either cast or easily machined. This leads to significant cost savings.
- Another important feature of this invention is the efficient cooling. The cooling medium, delivered with low pressure loss through the straight interior passage, flows out through feed-holes into the gap, defined between load carrying structure and shell, and passes the gap with a high yield rate of convective cooling.
- In addition, the feed-holes can serve to impinge the inner wall of the shell thus additionally increasing the cooling effect with the result of an improved lifetime of the components and/or a reduced amount of cooling medium. A reduced use of cooling medium results in higher engine performance.
- By way of example, an embodiment of the present disclosure is described more fully hereinafter with reference to the accompanying drawings, in which:
-
FIG. 1 a,b show modular gas turbine blades known from the prior art (WO 2011/058043 and EP 2189626); -
FIG. 2 shows a sectional top view of a modular turbine blade or vane design suitable for the present invention; -
FIG. 3 shows a sectional side view of a modular blade according to an embodiment of the invention; -
FIG. 4 a,b,c show in detail tip fixation variants of a shell; -
FIG. 5 shows a sectional side view of a detail of a modular vane according to the invention; -
FIG. 6 a,b show a sectional side view of another detail of a modular vane according to the invention. - Exemplary embodiments of the present invention are now described with references to the drawings, wherein like reference numerals are used to refer to like elements throughout. In the following description, for purposes of explanation, numerous specific details are set forth to provide a thorough understanding of the disclosure. However, the present disclosure may be practiced without these specific details, and is not limited to the exemplary embodiment disclosed herein.
-
FIG. 2 shows a sectional top view of a modular gas turbine blade or vane design in accordance with the present invention. Theairfoil 32 of ablade 30 orvane 31 comprises aload carrying structure 33, extending in longitudinal direction from the root portion 35 to thetip 36, and an aerodynamicallyshaped shell 34, extending in adistance 37 over theouter surface 38 of the carryingstructure 33 and defining theouter contour 40 of theairfoil 32. The carryingstructure 33 may be formed as a single unit or may be assembled from separate portions. At thetip 36 theairfoil 32 is sealed by acap 41. - The
shell 34 defines apressure side 42, asuction side 43, a leadingedge 44 and a trailingedge 45. Theshell 34 may be made from a different material compared to the carryingstructure 33 depending on the specification of the gas turbine. Theshell 34 can be made of a single unit or can be assembled from more than one single parts, e.g. two half-shells attached to the carryingstructure 33. - The carrying
structure 33 comprises at least oneinner passage 46 for conducting a flow of coolingair 49 from a reservoir in the blade or vane carrier towards theairfoil tip 36. The wall of the carryingstructure 33 is equipped with a number ofholes 47 connecting the at least oneinner passage 46 with thegap 48 defined between theouter surface 38 of the carryingstructure 33 and theinner surface 39 of theshell 34. The coolingair 49 flows out of theholes 47 into thegap 48. Preferably thisair 49 serves to impinge onto theinner surface 39 of theshell 34 and to effectively cool it. Before theair 49 exits thegap 48 through exit holes 50 in theshell 34, it effects convective cooling of theshell 34 when flowing along thegap 48 towards the exit holes 50. These exit holes 50 can include film cooling holes at thepressure side 42 and/or thesuction side 43 and/or the leadingedge 44 and/or the trailingedge 45, or these exit holes 50 can be a clearance between an outer orinner edge shell 34 and the carryingstructure 33, as shown in more detail inFIGS. 4 and 6 . -
FIG. 3 shows a sectional side view of ablade 30 according to the invention. Under theplatform 53 theload carrying structure 33 forms a fir tree portion 54 that fits into a complementary fir tree groove of a blade carrier of the turbomachine (not shown). Theplatform 53 is formed separately and joined to theload carrying structure 33 by means of welding, brazing or a retainer connection as described e.g. in U.S. Pat. No. 5,797,725. Theshell 34 extends longitudinally parallel toaxis 55 from a section under theplatform level 56 to theairfoil tip 36. Under theplatform level 56 theshell 34 is integrally joined with theload carrying structure 33 by an appropriate joining method, such as welding, brazing etc. Thus a fixed and gas-tight joint is achieved in this section. At theairfoil tip 36shell 34 andload carrying structure 33 are connected in a way that allows compensation of thermal expansion. Details of preferred embodiments for tip fixation are shown inFIGS. 4 a-c. - Additionally,
shell 34 andload carrying structure 33 are equipped with positive locking elements to laterally support theshell 34. Adovetail connection 57, as shown inFIG. 2 , supports theshell 34 and prevents its lateral deformation, but allows relative movement along thelongitudinal axis 55 to compensate thermal expansion and expansion caused by centrifugal forces. -
Exterior surface 38 of theload carrying structure 33 and/orinterior surface 39 of theshell 34 are equipped withturbulators 58, such as ribs, for increasing the heat transfer between the shell (34) and the coolingmedium 49. - In addition, the height of
individual turbulators 58 may correspond to thewidth 37 of thegap 48. In doing so, theseturbulators 58 act as a mechanical stop (distance holder) and prevent inward deformation of theshell 34, particularly at thepressure side 42 of theairfoil 32. -
FIG. 4 a, b, c show details of preferred embodiments forshell 34 fixation at theairfoil tip 36 allowing relative movement between theshell 34 and theload carrying structure 33. As mentioned above, the exposure of the individual components of the modular blade or vane to thermal expansion and—in case of rotating blades—to centrifugal forces requires a fixation of theshell 37 with clearance in longitudinal direction at one end. - In case of a rotating blade the fixed joint of the
shell 34 has to take up the centrifugal forces. From this reason the fixed joint betweenshell 34 and carryingstructure 33 is achieved in the root section 35, whereas at theairfoil tip 36 theshell 34 and theload carrying structure 33 are connected in a way that allows relative movement. As a consequence of this arrangement theshell 37 is subject to tension during rotation of theblade 30 and thereby additionally stabilized against deformation, such as buckling. - According to
FIG. 4 a theshell 34 ends in agroove 59 at the lower side of atip component 41. Such acomponent 41 may be, for example, a tip cap or a shroud element. Depth and width of thegroove 59 are sized so that a leakage of coolingmedium 49 is possible. The contour of thegroove 49 may be machined by any appropriate machining method. - An alternative embodiment of fixation of
shell 34 at theairfoil tip 36 is shown inFIG. 4 b. At thetip 36 theload carrying structure 33 broadens to the outer contour of theairfoil 32 thereby overlapping theupper edge 51 of theshell 34. At the lower side of this overlapping part 63 ashoulder 60 is machined so that theupper edge 51 fits into thisshoulder 60. A surroundingweld seam 61 at the outer edge of theshoulder 60 locks theshell 34 in its position. -
FIG. 4 c shows another variant of a tip fixation. Theshoulder 60 at the lower side of atip component 41, e.g. the shroud element, is limited outwards by astop bar 62. -
FIGS. 5 and 6 show sectional side views of avane 31 according to the invention.FIG. 5 shows an example for an outer diameter fixation of theshell 34, the carryingstructure 33, and theplatform 53, andFIG. 6 a and b disclose two arrangement variants of the carryingstructure 33, theshell 34 and theplatform 53 at the inner diameter of thevane 31. - In the embodiment according to
FIG. 5 the carryingstructure 33 broadens towards the end of theairfoil 32, thereby overlapping theouter edge 51 of theshell 34. At the inner side of the overlapping part 63 ashoulder 60 is machined so that theedge 51 of theshell 34 and a portion of itslateral surface 64 bear against theshoulder 60. By appropriate means, such as welding, brazing etc. theshell 34 is integrally joined to the carryingstructure 33. Beyond this fix joint the carryingstructure 33 andshell 34 are arranged in adistance 37, forming thelongitudinal gap 48. A number of cooling holes 47 delivers the cooling medium 49 from theinner passage 46 inside the carryingstructure 33 into thegap 48. For high efficient cooling the jets of the coolingmedium 49 impinge theinner surface 39 of theshell 34.Turbulators 58 on theouter surface 38 of theload carrying structure 33 support the turbulent flow of the coolingmedium 49, thus enhancing convective heat transfer.Individual turbulators 58 may serve as distance holders, as mentioned in connection withFIG. 3 . - The overlapping
part 63 of the carryingstructure 33 comprises an outerlateral surface 65. Theplatform 53 comprises a through-opening 66 with aninner surface 67.Lateral surface 65 of the carryingstructure 33 andinner surface 67 of the through-opening 66 are complementary designed. By a suitable method, e.g. bi-metal-casting, welding brazing etc. theplatform 53 is connected to theairfoil 32. - According to
FIG. 6 a theinner edge 52 ofshell 34 fits into theshoulder 60 of carryingstructure 33. At the outerlateral surface 65 of an overlappingpart 63 of the carryingstructure 33 theplatform 53 is connected to theairfoil 32. The size ofplatform 53 exceeds the size of the overlappingpart 63 of the carryingstructure 33 thus forming a groove in which theshell 34 ends. Anoptional seal 68 betweenairfoil shell 34 andplatform 53 prevents the penetration of hot gases from the hot gas path into this groove. - According to
FIG. 6 b the overlappingpart 63 of the carryingstructure 33 is broader, compared to the example ofFIG. 6 a, and theinner edge 52 of theairfoil shell 34 ends in agroove 59, machined into the overlappingpart 63. Optionally aseal 68 may be arranged to prevent the penetration of hot gases into thegroove 59. Theplatform 53 is connected to the outerlateral surface 65 of the carryingstructure 33 by any suitable method.
Claims (14)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP12182327.2A EP2703601B8 (en) | 2012-08-30 | 2012-08-30 | Modular Blade or Vane for a Gas Turbine and Gas Turbine with Such a Blade or Vane |
EP12182327.2 | 2012-08-30 |
Publications (1)
Publication Number | Publication Date |
---|---|
US20140234088A1 true US20140234088A1 (en) | 2014-08-21 |
Family
ID=46801332
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US14/012,101 Abandoned US20140234088A1 (en) | 2012-08-30 | 2013-08-28 | Modular blade or vane for a gas turbine and gas turbine with such a blade or vane |
Country Status (7)
Country | Link |
---|---|
US (1) | US20140234088A1 (en) |
EP (1) | EP2703601B8 (en) |
JP (1) | JP6016739B2 (en) |
KR (1) | KR101586210B1 (en) |
CN (1) | CN103696810B (en) |
AU (1) | AU2013221909B2 (en) |
RU (1) | RU2563046C2 (en) |
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US20160130951A1 (en) * | 2014-11-12 | 2016-05-12 | Alstom Technology Ltd. | Cooling for turbine blade platform-aerofoil joints |
US20170081966A1 (en) * | 2015-09-18 | 2017-03-23 | General Electric Company | Stator component cooling |
US20170122112A1 (en) * | 2014-04-16 | 2017-05-04 | Siemens Aktiengesellschaft | Controlling cooling flow in a cooled turbine vane or blade using an impingement tube |
EP3196410A1 (en) * | 2016-01-24 | 2017-07-26 | Rolls-Royce North American Technologies, Inc. | Turbine endwall and tip cooling for dual wall airfoils |
US20180100516A1 (en) * | 2016-10-12 | 2018-04-12 | Safran Aircraft Engines | Vane comprising an assembled platform and blade |
EP3318717A1 (en) * | 2016-11-08 | 2018-05-09 | Rolls-Royce Corporation | Undercut on airfoil coversheet support member |
US20180202295A1 (en) * | 2017-01-13 | 2018-07-19 | Rolls-Royce Corporation | Airfoil with Dual-Wall Cooling for a Gas Turbine Engine |
US20180371926A1 (en) * | 2014-12-12 | 2018-12-27 | United Technologies Corporation | Sliding baffle inserts |
US10668528B2 (en) | 2014-12-04 | 2020-06-02 | Siemens Aktiengesellschaft | Method for producing a rotor blade |
US11008878B2 (en) | 2018-12-21 | 2021-05-18 | Rolls-Royce Plc | Turbine blade with ceramic matrix composite aerofoil and metallic root |
US20230243267A1 (en) * | 2022-01-28 | 2023-08-03 | Raytheon Technologies Corporation | Components for gas turbine engines |
EP4310297A1 (en) * | 2022-07-20 | 2024-01-24 | General Electric Technology GmbH | Cooling circuit for a stator vane braze joint |
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US10605090B2 (en) * | 2016-05-12 | 2020-03-31 | General Electric Company | Intermediate central passage spanning outer walls aft of airfoil leading edge passage |
WO2018215143A1 (en) * | 2017-05-22 | 2018-11-29 | Siemens Aktiengesellschaft | Aerofoil |
GB201806542D0 (en) * | 2018-04-23 | 2018-06-06 | Rolls Royce Plc | A blade and a method of manufacturing a blade |
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US10668528B2 (en) | 2014-12-04 | 2020-06-02 | Siemens Aktiengesellschaft | Method for producing a rotor blade |
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EP3318717A1 (en) * | 2016-11-08 | 2018-05-09 | Rolls-Royce Corporation | Undercut on airfoil coversheet support member |
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US20180202295A1 (en) * | 2017-01-13 | 2018-07-19 | Rolls-Royce Corporation | Airfoil with Dual-Wall Cooling for a Gas Turbine Engine |
US11008878B2 (en) | 2018-12-21 | 2021-05-18 | Rolls-Royce Plc | Turbine blade with ceramic matrix composite aerofoil and metallic root |
US20230243267A1 (en) * | 2022-01-28 | 2023-08-03 | Raytheon Technologies Corporation | Components for gas turbine engines |
EP4310297A1 (en) * | 2022-07-20 | 2024-01-24 | General Electric Technology GmbH | Cooling circuit for a stator vane braze joint |
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Also Published As
Publication number | Publication date |
---|---|
JP2014047788A (en) | 2014-03-17 |
KR20140029282A (en) | 2014-03-10 |
JP6016739B2 (en) | 2016-10-26 |
EP2703601A1 (en) | 2014-03-05 |
RU2563046C2 (en) | 2015-09-20 |
KR101586210B1 (en) | 2016-01-18 |
RU2013140190A (en) | 2015-03-10 |
EP2703601B1 (en) | 2016-07-20 |
AU2013221909B2 (en) | 2015-05-21 |
CN103696810A (en) | 2014-04-02 |
EP2703601B8 (en) | 2016-09-14 |
CN103696810B (en) | 2016-09-28 |
AU2013221909A1 (en) | 2014-03-20 |
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