EP3447248A1

EP3447248A1 - Turbine blade assembly comprising a sealing element made of adhesive material

Info

Publication number: EP3447248A1
Application number: EP17187088.4A
Authority: EP
Inventors: Quentin Luc Balandier
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2017-08-21
Filing date: 2017-08-21
Publication date: 2019-02-27

Abstract

A turbine blade assembly includes a first and a second blade root of a first and a second turbine blade inserted adjoining to each other, into a rim section of a rotor disc thereby defining a common cavity between a first and a second non-aerofoil sections of the turbine blades. A sealing element is located inside the common cavity, adhered to an inner surface of the common cavity. The sealing element extends between the first and the second non-aerofoil sections to at least partially fill the common cavity. The sealing element includes an adhesive material, or is formed solely from the adhesive material. The sealing element seals the gap between the non-aerofoil sections and consequently protects the non-aerofoil sections, particularly the shanks of the turbine blades and the rim section of the rotor disc, from the harmful effects of the hot working gases.

Description

The present invention relates to a turbine blade assembly for a gas turbine, and particularly to sealing techniques used for turbine blade assemblies. Furthermore, the present invention relates to a method for manufacturing a turbine blade assembly for a gas turbine.
In gas turbines, a plurality of turbine blades is mounted on a rotor disc to form a stage of the turbine section. Each stage of the turbine section of the gas turbine is a turbine blade assembly comprising a rotor disc and a plurality of turbine blades mounted on the rotor disc, typically using 'firtrees' serrations or one or several lobes per blade. The 'firtrees' serrations or the lobes form the blade root. The turbine blades are mounted around the rotor disc one after another by inserting roots of the turbine blades into corresponding slots formed in a rim section of the rotor disc such that platforms, and therefore the aerofoils extending out from the platforms, of the turbine blades are arranged circumferentially and serially around the rim of the rotor disc. The platforms of the adjoining turbine blades are arranged such that a portion of a hot gas path, that confines a mainstream flow of the hot working gases within the gas turbine, is defined by the circumferentially arranged blade platforms. A gap may exist between adjoining turbine blades, particularly between adjoining blade platforms, arranged around the rotor disc and may lead to leakage of hot working gas from the mainstream flow of the gas turbine into a non-aerofoil section of the turbine blades arranged around the rotor disc.
Conventionally, the blades can be axially positioned and retained into the disc rim slots using front and/or back locking plates, which prevent relative axial movement between the blades and the disc.
The turbine blade has two sections or parts: an aerofoil section and a non-aerofoil section. The non-aerofoil section of the turbine blade is a section of the turbine blade below the aerofoil, and that extends in a direction radially opposite to the emanating aerofoils. The non-aerofoil section includes a blade platform, a shank, and a blade root. The shank generally links the blade platform, wherefrom the aerofoil emanates, and the blade root which is inserted into the slot of the rim section of the rotor disc.
To avoid or reduce leakage of the hot working gas from the mainstream flow onto the blade root and the shank of the turbine blade and consequently onto the rim section of the rotor disc, conventionally a sealing arrangement, such as a seal strip or wire, is used between adjacent turbine blades. By positioning a seal strip or wire between adjacent turbine blades, an injection of hot working gas through the gaps between adjacent turbine blades arranged on the rotor disc and onto surfaces and components of the non-aerofoil section of the turbine blades is intended to be prevented.
Furthermore the conventional seal strip or wire is intended to prevent injection of the hot working gas onto the rotor disc and especially onto the rim section of the rotor disc. Moreover, cooling air which flows through cavities inside the non-aerofoil section of the turbine blade and/or through cavities inside or associated with the rotor disc supporting the turbine blades is prevented from disappearing out unintentionally into the mainstream flow of the hot working gas.
WO 2009/053169 A1 shows a turbine blade assembly for a gas turbine, wherein a seal strip is mounted between a platform and a root cavity. The seal strip is further in contact with a locking plate for locking the airfoil to an airfoil disc.
WO 2008/046684 A1 discloses a turbine blade assembly for a gas turbine. A seal strip is attached to a respective platform of adjacent turbine blades in order to minimize the leakage of hot working gas into the space under the platforms.
The conventionally known seal strips (or other known similar devices such as seal wires for example) are inserted into slots or grooves that are formed in the adjacent turbine blades arranged on the rotor disc, particularly in the platforms of the adjacent turbine blades. The seal strips when assembled generally form small gaps or clearances within the grooves. The gaps or clearances may be greater due to manufacturing or assembling inaccuracies, and may further be increased due to thermal expansion and physical deformations that the turbine blades and the conventionally known seal strips are subjected to during operation of the gas turbine. Consequently, the seal strips inserted into the grooves are not completely air tight sealing arrangement and thus hot working gas from the mainstream flow leaks from around the seal strip(s) due to the gaps between the seal strips and the walls of the grooves or the slots into which the seal strips are inserted. This results in temperature increase at the shanks, at the blade roots and the disc rims supporting the turbine blades, hence reducing lives of the components, i.e. of the turbine blades and the rotor discs.
Furthermore, the hot gases have corrosive effects which results into further weakening of the components' structural integrity and reduction in component lives, especially of the shanks of the turbine blades. Therefore, there exists a need for better sealing arrangements that protect turbine blade non-aerofoil sections, particularly the shank, and the structures wherefrom the turbine blades are supported i.e. the rotor discs, from the abovementioned harmful effects of the hot working gases.
Thus, an object of the present invention is to provide a sealing technique between adjacent turbine blades of a gas turbine that at least partially obviates the hot gas leakage that occurs in the conventionally known turbine blade assemblies with seal strips or wires due to the gaps that result from manufacturing and assembly of such seal strips or wires into the grooves of the adjacent turbine blades. The sealing technique is desired to at least partially, and preferably substantially or completely, protect the non-aerofoil sections of the turbine blades, particularly to protect the shanks of the turbine blades, and/or to protect the rim section of the rotor disc on which the turbine blades are arranged.
The above object is achieved by a turbine blade assembly for a gas turbine according to claim 1, a gas turbine comprising a turbine blade assembly according to claim 9, and a method for manufacturing a turbine blade assembly according to claim 10, of the present technique. Advantageous embodiments of the present technique are provided in dependent claims.
In a first aspect of the present technique, a turbine blade assembly for a gas turbine is presented. The turbine blade assembly includes a rotor disc, a first and a second turbine blades and a sealing element. The rotor disc includes a rim section.
The first turbine blade includes a first aerofoil section and a first non-aerofoil section. The first aerofoil section, hereinafter also referred to as the first aerofoil, has a leading edge and a trailing edge. The first non-aerofoil section has a first platform, a first front face (on side of the leading edge of the first aerofoil), a first back face (on side of trailing edge of the first aerofoil), a first shank and a first blade root. The first shank, also referred to as the first shank section, extends between the first blade root and the first platform. The first turbine blade is mounted on the rotor disc by the first blade root being inserted into the rim section of the rotor disc.
The second turbine blade includes a second aerofoil section and a second non-aerofoil section. The second aerofoil section, hereinafter also referred to as the second aerofoil, has a leading edge and a trailing edge. The second non-aerofoil section has a second platform, a second front face (on side of the leading edge of the second aerofoil), a second back face (on side of trailing edge of the second aerofoil), a second shank and a second blade root. The second shank, also referred to as the second shank section, extends between the second blade root and the second platform. The second turbine blade is mounted on the rotor disc by the second blade root being inserted into the rim section of the rotor disc.
The first and the second turbine blades are mounted on the rotor disc such that the first non-aerofoil section adjoins the second non-aerofoil section, in other words the first and the second turbine blades are next to each other. The first and the second non-aerofoil sections define a common cavity thereinbetween. In other words, at least one of the first platform, the first front face, the first back face, and the first shank in combination with at least one of the second platform, the second front face, the second back face, and the second shank, at least partially encloses a space or volume thereinbetween, the space or volume so enclosed is referred to as the common cavity. Preferably, the first shank in combination with the second shank at least partially encloses the space or volume thereinbetween forming the common cavity. One or more of the first platform, the first front face, the first back face, and the first shank and one or more of the second platform, the second front face, the second back face and the second shank form walls or inner surface of the common cavity.
The sealing element is located inside the common cavity. The sealing element is adhered to the inner surface of the common cavity. The sealing element extends between the first and the second non-aerofoil sections to at least partially fill the common cavity. The sealing element includes an adhesive material, or is formed solely from the adhesive material. The sealing element seals the gap between the non-aerofoil sections, and thereby obviates, completely or partially, leakage of the hot working gas from the mainstream flow into the space beneath the platform i.e. onto the non-aerofoil sections of the turbine blades, consequently protecting the turbine blade root and shank sections, and the rotor disc on which the turbine blades are mounted, especially the rim section of the rotor disc, from the abovementioned harmful effects of the hot working gases.
In an embodiment of the turbine blade assembly, a rim surface i.e. a circumferentially extending surface perpendicular to a central axis of the rotor disc forms a base of the common cavity i.e. the rim surface forms a part of the inner surface of the common cavity. Thus the rim surface of the rotor disc is protected from the aforementioned harmful effects of the hot working gases. Furthermore, being supported and being adhered to the rim surface provides the sealing element with stability in position when the gas engine is operated, i.e. when the rotor disc is rotated.
In another embodiment of the turbine blade assembly, a surface of a first lower platform of the first turbine blade and a surface of a second lower platform of the second turbine blade, together form a base of the common cavity i.e. the surfaces of the first and the second lower platforms form a part of the inner surface of the common cavity. Thus the surfaces of the lower platforms are protected from the aforementioned harmful effects of the hot working gases. Furthermore, being supported and being adhered to the surfaces of the lower platforms provides the sealing element with stability in position when the gas engine is operated, i.e. when the rotor disc is rotated.
In another embodiment of the turbine blade assembly, the first shank of the first turbine blade and the second shank of the second turbine blade define the common cavity. In a related embodiment of the turbine blade assembly, the first shank includes a first cavity and/or the second shank includes a second cavity, and wherein the common cavity includes the first cavity and/or the second cavity.
In another embodiment, the turbine blade assembly further includes at least one annular flat disc adhered on a side surface of the rim section of the rotor disc. The annular flat disc at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first and the second blade roots and/or the shanks, and the rim section of the rotor disc thereinbetween. The annular flat disc includes the adhesive material, or is solely formed from the adhesive material. In a related embodiment, the turbine blade assembly includes the annular flat disc one adhered on each side surface of the rim section of the rotor disc. Thus, the first and the second blade roots, and/or the first and the second shanks, and/or the rim section of the rotor disc thereinbetween are protected from the aforementioned harmful effects of the hot working gas.
In another embodiment of the turbine blade assembly, the annular flat disc at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first blade root and the first shank, the second blade root and the second shank, and a side opening of the common cavity thereinbetween. In a related embodiment, the turbine blade assembly includes the annular flat disc, one adhered on each side surface of the rim section of the rotor disc wherein each of the annular flat discs at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first blade root and the first shank section, the second blade root and the second shank section, and the side opening of the common cavity thereinbetween. Thus, the first blade root and the first shank section, the second blade root and the second shank section, and the side opening of the common cavity thereinbetween i.e. side of the sealing element are protected from the aforementioned harmful effects of the hot working gas.
In another embodiment of the turbine blade assembly, the annular flat disc is aerodynamically shaped. This smoothens out the surface of the rim section and the blade roots inserted into the rim section of the rotor disc and results in reduction of drag on the rotor disc when the rotor disc is rotating.
In another embodiment of the turbine blade assembly, the adhesive material is a ceramic or a graphite compound. Such composite adhesive materials have a high thermal stability i.e. such adhesive materials do not melt or liquefy at high temperatures that exist within the gas turbine, especially at the non-aerofoil sections of the turbine blades and the rim section of the rotor disc. The ceramic composite adhesive material is thermally stable between 400°C (degrees Celsius) and 1200°C, and more particularly between 900°C and 1100°C. Furthermore, the adhesive material is such that it adheres or sticks to the inner surface of the common cavity i.e. to the alloy material of the turbine blade non-aerofoil sections and/or to the alloy material of the rim section of the rotor disc.
In a second aspect of the present technique, a gas turbine is presented. The gas turbine includes a turbine blade assembly. The turbine blade assembly is same as the turbine blade assembly according to the aforementioned first aspect of the present technique.
In a third aspect of the present technique, a method for manufacturing a turbine blade assembly is presented. In the method, a first blade root and a second blade root of a first turbine blade and a second turbine blade, respectively, are inserted adjoining to each other, into a rim section of a rotor disc thereby defining a common cavity between a first non-aerofoil section of the first turbine blade and a second non-aerofoil section of the second turbine blade. The first and the second turbine blades and the rotor disc are same as the turbine blade assembly according to the aforementioned first aspect of the present technique. Thereafter in the method, an adhesive material is poured into the common cavity such that the adhesive material adheres to an inner surface of the common cavity. The adhesive material may be poured through a gap between the first and second platforms or through one or more holes in the first and/or the second platform. The adhesive material cures to form a sealing element located within the common cavity. The adhesive material is poured into the common cavity such that the sealing element so formed by the pouring, and subsequent curing, of the adhesive material extends between the first and the second non-aerofoil sections and thereby at least partially fills the common cavity. The adhesive material may be, but not limited to, a ceramic (such as a metallic ceramic) or graphite compound same as aforementioned.
In the turbine blade assembly manufactured by the aforementioned method, the sealing element seals the gap between the platforms and thereby obviates, completely or partially, leakage of the hot working gas from the mainstream flow into the space beneath the platform i.e. onto the non-aerofoil sections of the turbine blades, consequently protecting the turbine blade non-aerofoil sections, and the rotor disc on which the turbine blades are mounted, especially the rim section of the rotor disc, from the abovementioned harmful effects of the hot working gas.
In an embodiment of the method for manufacturing the turbine blade assembly, in the method before pouring the adhesive material into the common cavity, the rotor disc with the first and the second turbine blades so inserted, is laid on a flat surface such that a second side of the rotor disc is positioned on the flat surface and such that a second side opening of the common cavity, i.e. the opening of the common cavity at the second side of the rotor disc is closed by the flat surface. The second side opening of the common cavity is closed by the flat surface such that the adhesive is sealed off from flowing out of the common cavity via the second side opening of the common cavity. In this embodiment of the method, the adhesive material is poured into the common cavity from a first side, i.e. the face opposite to the second side or face of the rotor disc, and through a first side opening of the common cavity. In the method, the rotor disc is removed from the flat surface after the adhesive material is cured to form the sealing element. In other words the rotor disc and flat surface are removed from each other such that the flat surface does not close the second side opening of the common cavity any further.
In another embodiment of the method for manufacturing the turbine blade assembly, the adhesive material is spread in form of an annular flat disc on a side surface of the rim section of the rotor disc, such that the adhesive material at least partially covers, by adhesion, the first and the second blade roots, and/or the first and the second shanks and/or the rim section of the rotor disc thereinbetween. The annular disc of adhesive material during curing or solidification adheres onto the first and the second blade roots and the rim section of the rotor disc thereinbetween, and consequently protects the first and the second blade roots and the rim section of the rotor disc thereinbetween from the aforementioned harmful effects of hot working gas. In a related embodiment of the method, one annular flat disc adhered on each side surface of the rim section of the rotor disc is formed.
In another embodiment of the method for manufacturing the turbine blade assembly, the adhesive material so spread in form of the annular flat disc on the side surface of the rim section of the rotor disc at least partially covers, by adhesion, the first non-aerofoil section, the second non-aerofoil section, and a side opening of the common cavity thereinbetween. In a related embodiment of the method, the annular flat disc, one adhered on each side surface of the rim section of the rotor disc is formed on either sides or faces of the rotor disc. Each of the annular flat discs at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first non-aerofoil section, the second non-aerofoil section, and the side opening of the common cavity thereinbetween. Thus, the first non-aerofoil section, the second non-aerofoil section, and the side opening of the common cavity thereinbetween i.e. side of the sealing element are protected from the aforementioned harmful effects of the hot working gas.
In another embodiment of the method for manufacturing the turbine blade assembly, the adhesive material so spread is shaped into an aerodynamic form. This may be achieved for example by flattening a surface of the adhesive material so poured by pressing onto the surface a flat sheet or a shaped sheet. The aerodynamic form of the annular flat disc may include a flat surface in the annular part of the annular flat disc and bent surfaces at an outer and inner circumference of the annular flat disc and thus avoiding any sharp edges or angles within the annular flat surface.
The sealing element in the present technique is present in the turbine blade assembly as adhered to or glued to the inner surface of the common cavity. This provides enhanced damping and ensures that the sealing element is not displaced from its spatial position during operation of the gas turbine. Furthermore, since the sealing element has a fixed position, the sealing element of the present technique does not contribute to any imbalance or dynamic issues of the rotor disc, and of the turbine blade assembly, during operation of the gas turbine that may have resulted from a conventional seal strip which is, albeit slightly, displaceable within the grooves into which it is inserted. The above mentioned attributes and other features and advantages of the present technique and the manner of attaining them will become more apparent and the present technique itself will be better understood by reference to the following description of embodiments of the present technique taken in conjunction with the accompanying drawings, wherein:

FIG 1: shows part of a gas turbine in a sectional view and in which a turbine blade assembly of the present technique is incorporated;
FIG 2: schematically illustrates a perspective view of an exemplary embodiment of a section of the turbine blade assembly without depicting the sealing element;
FIG 3: schematically illustrates a cross-sectional view, from a circumferential direction, of a part of an exemplary embodiment of a turbine blade of the turbine blade assembly of the present technique;
FIG 4: schematically illustrates a cross-section, as viewed from an axial direction, of a part of an exemplary embodiment of the turbine blade assembly of the present technique depicting a common cavity;
FIG 5: schematically illustrates a perspective view of a part of an exemplary embodiment of the turbine blade of the turbine blade assembly of the present technique;
FIG 6: schematically illustrates a perspective view of a part of another exemplary embodiment of the turbine blade of the turbine blade assembly of the present technique;
FIG 7: schematically illustrates a perspective view of a part of yet another exemplary embodiment of the turbine blade of the turbine blade assembly of the present technique;
FIG 8: schematically illustrates a part of an exemplary embodiment of the turbine blade assembly of the present technique with the sealing element of the present technique;
FIG 9: schematically illustrates a part of another exemplary embodiment of the turbine blade assembly of the present technique with the sealing element of the present technique;
FIG 10: schematically illustrates a part of yet another exemplary embodiment of the turbine blade assembly of the present technique with the sealing element of the present technique;
FIG 11: schematically illustrates a part of an exemplary embodiment of the turbine blade assembly of the present technique with an exemplary embodiment of an annular flat disc;
FIG 12: schematically illustrates an exemplary embodiment of a rotor disc of the turbine blade assembly of the present technique depicting a positioning of the annular flat disc; and
FIG 13: presents a flow chart of an exemplary embodiment of a method for manufacturing the turbine blade assembly of the present technique; in accordance with aspects of the present technique.

Hereinafter, above-mentioned and other features of the present technique are described in details. Various embodiments are described with reference to the drawing, wherein like reference numerals are used to refer to like elements throughout. In the following description, for the purpose of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It may be noted that the illustrated embodiments are intended to explain, and not to limit the invention. It may be evident that such embodiments may be practiced without these specific details.
It may be noted that in the present disclosure, the terms "first", "second", "third" etc. are used herein only to facilitate discussion, and carry no particular temporal or chronological significance unless otherwise indicated.
FIG. 1 shows an example of a gas turbine 10 or a gas turbine engine 10 in a sectional view. The gas turbine 10 comprises, in flow series, an inlet 12, a compressor or compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20. The gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine 10. The shaft 22 drivingly connects the turbine section 18 to the compressor section 14.
In operation of the gas turbine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a longitudinal axis 35 of the burner, a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air 24 enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine section 18.
The turbine section 18 comprises a number of blade carrying discs 36, also referred to as the rotor discs 38, attached to the shaft 22. In the present example, two rotor discs 36 each carry an annular array of turbine blades 38 are depicted. However, the number of blade carrying discs 36 could be different, i.e. only one disc 36 or more than two discs 36. In addition, guiding vanes 40, which are fixed to a stator 42 of the gas turbine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and that turn the flow of working gas onto the turbine blades 38.
The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and compressor blade stages 48. The compressor blade stages 48 comprise a disc supporting an annular array of compressor blades. The compressor section 14 also comprises a casing 50 that surrounds the compressor blade stages 48 and supports the vane stages 48. The guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present air flow at an optimal angle for the compressor blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operational conditions.
The casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of compressor blades 48.
The present technique is described with reference to the above exemplary gas turbine 10 having a single shaft or spool connecting a single, multi-stage compressor section 14 and a single, one or more stage turbine section 18. However, it should be appreciated that the present technique is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications. Furthermore, the cannular combustor section arrangement 16 is also used for exemplary purposes and it should be appreciated that the present technique is equally applicable to annular type and can type combustion chambers.
The terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine, unless otherwise stated.
FIG 2 schematically shows a turbine blade assembly 100 that may be incorporated into the turbine section 18 of the gas turbine 10 as shown in FIG 1. The turbine blade assembly 100, hereinafter also referred to as the blade assembly 100 or simply to as the assembly 100, includes a plurality of turbine blades 110,210,310 mounted on a rotor disc 70 shown in FIG 12. The rotor disc 70 is not depicted in FIG 2 for sake of simplicity. The turbine blades 110,210,310 of FIG 2 and the rotor disc 70 of FIG 12 are exemplary embodiments of the turbine blades 38 and the rotor discs 36 explained hereinabove in relation to FIG 1.
Each turbine blade, for example a first turbine blade 110, a second turbine blade 210, and a third turbine blade 310 of the assembly 100 has an aerofoil section for example a first aerofoil section 117 or a first aerofoil 117, a second aerofoil section 217 or a second aerofoil 217, and a third aerofoil section 317 or a third aerofoil 317 and a blade root for example a first blade root 120, a second blade root 220, and a third blade root 230, respectively. The first aerofoil 117, the second aerofoil 217, and the third aerofoil 317 emanate from a first surface 116 of a first platform 114, a second surface 216 of a second platform 214, and a third surface 316 of a third platform 314 of the first turbine blade 110, the second turbine blade 210 and the third turbine blade 310, respectively. FIG 2 also depicts an arrow marked with reference numeral 102 showing an axial direction, an arrow marked with reference numeral 103 showing a circumferential direction, and an arrow marked with reference numeral 104 showing a radial direction - all with respect to the rotational axis 20 of FIG 1.
As depicted in FIG 12, the rotor disc 70, hereinafter also referred to as the disc 70, includes a hub 75 using which the disc 70 is mounted on a turbine shaft (not shown in FIG 12), a body 76 and a rim section 73. The rim section 73 is the peripheral region of the disc 70 and includes a plurality of slots 77 arranged circumferentially on the outer edge of the disc 70 forming the rim section 73. FIG 12 depicts an orientation of the disc 70 with respect to the axial direction 102, which is same as an orientation of the assembly 100, when installed within the gas turbine 10.
The blade roots 120,220,320 of the turbine blades 110,210,310 are inserted into the slots 77 of the rim section 73 of the disc 70 thereby forming a disc 70 having an array of arranged or mounted, generally circumferentially, turbine blades 110,210,310. The blade roots 120,220,320 may have variety of shapes for example fir tree design, dove-tail design and so on and so forth, and the slots 77 have corresponding shape for receiving the blade roots 120,220,320 of the turbine blades 110,210,310.
The turbine blade assembly 100 includes the disc 70 and a plurality of turbine blades 110,210,310 mounted on the disc 70. FIG 3 schematically illustrates a cross-sectional view, from the circumferential direction 103, of a part of an exemplary embodiment of the first turbine blade 110, which is same for the second turbine blade 210 and the third turbine blade 310, of the turbine blade assembly 1, and for any further turbine blades (not shown). FIG 4 schematically illustrates a cross-section, as viewed from the axial direction 102, of a part of an exemplary embodiment of the turbine blade assembly 100 of the present technique depicting an arrangement of the first turbine blade 110 and the second turbine blade 210 mounted on the disc 70.
As shown in FIGs 2 to 6, in combination with FIG 12, the first turbine blade 110 has two sections or parts: the first aerofoil section 117 and a first non-aerofoil section 112. The first non-aerofoil section 112 of the first turbine blade 110 is a section or part of the first turbine blade 110 below the first aerofoil 117, and that extends in a direction radially opposite to the emanating first aerofoil 117. The first non-aerofoil section 112 has the first platform 114, a first front face 115, a first back face 113, a first shank 401 and the first blade root 120. The first aerofoil 117 has a leading edge and a trailing edge. The first front face 115 is the part of the first non-aerofoil section 112, particularly side or surface of the shank 401, at the leading edge side of the first aerofoil 117 when viewed along the axial direction 102. In other words, the first front face 115 is the part of the first non-aerofoil section 112, particularly side or surface of the shank 401, radially beneath the leading edge of the first aerofoil 117. The first back face 113 is the part of the first non-aerofoil section 112, particularly side or surface of the shank 401, at the trailing edge side of the first aerofoil 117 when viewed along the axial direction 102. In other words, the first back face 113 is the part of the first non-aerofoil section 112, particularly side or surface of the shank 401, radially beneath the trailing edge of the first aerofoil 117. As mentioned above the first turbine blade 110 is mounted on the rotor disc 70 by the first blade root 120 being inserted into one of the slots 77 of the rim section 73 of the rotor disc 70. The first shank 401 is limited by the first platform 114, the first blade root 120, the first back face 113 and the first front face 115 of the first turbine blade 110.
Similarly, the second turbine blade 210 has two sections or parts: the second aerofoil section 217 and a second non-aerofoil section 212. The second non-aerofoil section 212 of the second turbine blade 210 is a section or part of the second turbine blade 210 below the second aerofoil 217, and that extends in a direction radially opposite to the emanating second aerofoil 217. The second non-aerofoil section 212 has the second platform 214, a second front face 215, a second back face 213, a second shank 402 and the second blade root 220. The second aerofoil 217 has a leading edge and a trailing edge. The second front face 215 is the part of the second non-aerofoil section 212, particularly side or surface of the shank 402, at the leading edge side of the second aerofoil 217 when viewed along the axial direction 102. In other words, the second front face 115 is the part of the second non-aerofoil section 212, particularly side or surface of the shank 402, radially beneath the leading edge of the second aerofoil 217. The second back face 213 is the part of the second non-aerofoil section 212, particularly side or surface of the shank 402, at the trailing edge side of the second aerofoil 217 when viewed along the axial direction 102. In other words, the second back face 213 is the part of the second non-aerofoil section 212, particularly side or surface of the shank 402, radially beneath the trailing edge of the second aerofoil 217. As mentioned above the second turbine blade 210 is mounted on the rotor disc 70 by the second blade root 220 being inserted into one of the slots 77 of the rim section 73 of the rotor disc 70. The second shank 402 is limited by the second platform 214, the second blade root 220, the second back face 213 and the second front face 215 of the second turbine blade 210.
As shown in FIG 5 the back faces 113,213 and the front faces 115,215 may be parts or regions of a continuous surface of the non-aerofoil sections 112,212, and particularly of the shank 401,402 of the non-aerofoil sections 112,212. Alternatively, as shown in FIGs 6 and 7, the back faces 113,213 and the front faces 115,215 may be disposed as projections in the circumferential direction and preferably substantially parallel to each other.
When mounted on the rotor disc 70, the blade roots 120, 220 are inserted within body of the rotor disc 70. The platforms 114, 214 and the shanks 401,402 are present outside the body of the rotor disc 70. The back faces 113,213 and the front faces 115, 215 are also outside the body of the rotor disc 70 and in-between the blade roots 120,220 and the platforms 114,214.
As a result of being mounted adjoining or adjacent, and as depicted in FIG 4, the first and the second turbine blades 110,210 are disposed such that the first non-aerofoil section 112 adjoins the second non-aerofoil section 212. As a result of the adjacent mounting a common cavity 99 is created between the first and the second non-aerofoil sections 112,212, or in other words the first and the second non-aerofoil sections 112, 212 define the common cavity 99 thereinbetween, i.e. in between each other. Particularly, the common cavity 99 is formed between the shanks of the adjoining turbine blades arranged on the disc 70, i.e. between the first shank 401 of the first turbine blade 110 and the second shank 402 of the second turbine blade 210.
As shown in FIGs 6 and 7, the first non-aerofoil section 112 may include a first cavity 111, particularly the first shank 401 includes the first cavity 111. The first cavity 111 may be partially surrounded by at least the first platform 114, the first front face 115 and/or the first back face 113. Optionally, as shown in FIG 7 the first cavity 111 may be surrounded by the rim surface 74 of the rotor disc 70 and/or as shown in FIG 6 by a surface 119 of a first lower platform 118 of the first turbine blade 110. The first lower platform 118 is an additional platform, besides the first platform 114, and is located radially beneath the first platform 114.
Similarly, as shown in FIGs 6 and 7, the second non-aerofoil section 212 may include a second cavity 211, particularly the second shank 402 includes the second cavity 211. The second cavity 211 may be partially surrounded by at least the second platform 214, the second front face 215 and/or the second back face 213. Optionally, as shown in FIG 7 the second cavity 211 may be surrounded by the rim surface 74 of the rotor disc 70 and/or as shown in FIG 6 by a surface 219 of a second lower platform 218 of the second turbine blade 210. The second lower platform 218 is an additional platform, besides the second platform 214, and is located radially beneath the second platform 114.
The common cavity 99 may include the first cavity 111 and/or the second cavity 211, when present in the turbine blades 110,210. Furthermore, as shown in FIGs 2 and 4, the first and second platforms 114,214 when in a mounted state on the disc 70 may have a gap 3 in-between them i.e. in-between the first and second platforms 114,214. The gap 3 may also form a part of the common cavity 99.
Thus, as aforementioned and as depicted in the exemplary embodiments of the turbine assembly 100 in FIGs 2 to 7, the common cavity 99 is formed in-between the non-aerofoil sections 112,212 and may have the inner surface 98 formed by the platforms 114,214, the front faces 115,215, and/or the back faces 113,213. Furthermore, the inner surface 98 of the common cavity 99 is formed by the rim surface 74 and/or the surface 119,219 of the lower platforms 118,218. The common cavity 99 may optionally includes the first and/or the second cavities 111,211. The common cavity 99 may further optionally include the gap 3.
According to the present technique, as shown in FIGs 8, 9 and 10, the turbine assembly 100 further includes a sealing element 60. The sealing element 60 is located inside the common cavity 99. The sealing element 60 is formed in situ i.e. within the common cavity 99, as opposed to conventionally known seals that are prefabricated before being assembled into the conventional turbine blade assemblies. The sealing element 60 is formed of adhesive material. The adhesive material in its fluid form, for example as a liquid or a solution, is poured into the common cavity 99 and is cured within the common cavity 99, consequently resulting into formation of the sealing element 60 within the common cavity 99 resulting from the curing or solidification of the adhesive material. As a result, the sealing element 60 is adhered to the inner surface 98 of the common cavity 99. The sealing element 60 extends between the first and the second non-aerofoil sections 112,212 to at least partially fill the common cavity 99. The sealing element 60 includes an adhesive material, i.e. a material that acts as an adhesive to materials such as alloys from which the turbine blades 110,210 and the rim section 73 of the disc 70 are formed. The sealing element 60 may include the adhesive material along with some non-adhesive material, or may be formed solely from the adhesive material.
An example of the adhesive material is a ceramic or a graphite composite, such as a metallic ceramic composite, and particularly those composites that have a high thermal stability i.e. such adhesive materials do not melt or liquefy at high temperatures that exist within the gas turbine 10, especially at the temperatures that exist at the non-aerofoil sections 112,212 of the turbine blades 110,210 and the rim section 73 of the rotor disc 70, when the gas turbine 10 is being operated, particularly when the gas turbine 10 is being operated at its maximum. The ceramic or graphite composite adhesive material is thermally stable between 400°C (degrees Celsius) and 1200°C, and more particularly between 900°C and 1100°C. Furthermore, the adhesive material is such that it adheres or sticks to the inner surface 98 of the common cavity 99 i.e. to the alloy material of the turbine blade non-aerofoil sections 112,212 and/or to the alloy material of the rim section 73 of the rotor disc 70 before the adhesive material has been cured and remains stuck or adhered to the inner surface 98 of the common cavity 99 i.e. to the alloy material of the turbine blade non-aerofoil sections 112,212 and/or to the alloy material of the rim section 73 of the rotor disc 70 after the adhesive material has been cured or solidified to form the sealing element 60.
The sealing element 60 acts as a sealant with respect to the hot working gas headed towards the rim section 73 of the disc 70. The sealing element 60 also acts as a sealant with respect to the hot working gas for parts of the non-aerofoil sections 112,212 that are embedded within the adhesive material of the sealing element 60 or for parts of the non-aerofoil sections 112,212 to which the adhesive material of the sealing element 60 is adhered or stuck. The sealing element 60 may preferably fill in the common cavity 99 completely, including or excluding the gap 3.
FIG 8 shows the assembly 100, as opposed to FIG 4, depicting the sealing element 60 within the common cavity 99. The sealing element 60 in FIG 8 is adhered to the non-aerofoil sections 112,212 other than the blade roots 120,220, and to the rim surface 74 between the non-aerofoil sections 112,212. FIG 9 also shows assembly 100 depicting the sealing element 60 in the central cavity 99. The sealing element 60 in FIG 9, in combination with FIG 6, is adhered to the non-aerofoil sections 112,212 other than the blade roots 120,220, and to the rim surface 74 between the non-aerofoil sections 112,212, and to the lower platforms 118,218 including the surfaces 119, 219 of the lower platforms 118,218.
The turbine blade assembly 100 may have other components such as a lock strip 4, a seal wire 6 and a seal strip 7 as shown in FIG 6. The structure and functions of the lock strips, the seal wires and the seal strips within turbine blade assemblies in conventionally known and therefore not explained herein in further details for sake of brevity. FIG 10 shows the assembly 100 depicting the sealing element 60. The assembly 100 of FIG 10 is same as the assembly 100 of FIG 8 but further comprises the seal wire 6. The other components such as the lock strip 4 and the seal strip 7 may also be present in the assembly 100.
Referring to FIGs 11 and 12 further exemplary embodiments of the assembly 100 have been explained hereinafter. As shown in FIG 12, the rotor disc 70 has a first side 78 and a second side 79 on flip side of the fist side 78. A surface of the rim section 73 on the first side 78 is referred to as a first side surface 71 of the rim section 73 of the disc 70 and a surface of the rim section 73 on the second side 79 is referred to as a second side surface 72 of the rim section 73 of the disc 70. In an embodiment of the assembly 100 as shown in FIG 11, the assembly further includes at least one annular flat disc 80 adhered on one of the side surfaces 71,72 of the rim section 73 of the rotor disc 74. The annular flat disc 80 at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first and the second blade roots 120,220, the first and the second shanks 401,402 particularly at the front faces 115,215 and/or the back faces 113,213 forming the sides of the shanks 401,402 and the rim section 73 of the rotor disc 70 thereinbetween. The annular flat disc 80 includes the adhesive material, or is solely formed from the adhesive material. The adhesive material used for the annular flat disc 80 may be same as the adhesive material used for making the sealing element 60. The assembly 100 may include one such annular flat disc 80 as depicted in FIGs 11 and 12 or may includes in a related embodiment (not shown) two such annular flat discs 80 - one annular flat disc 80 adhered on each side surface 71,72 of the rim section 73 of the rotor disc 70. It may be noted that FIGs 11 and 12 show only part of the annular flat disc 80. The flat surfaces of the annular flat disc 80 are radially aligned whereas the annular flat disc 80 is disposed circumferentially around the rotational axis 20 or around the hub 75 of the rotor disc, but preferably limited within the rim section 73 of the rotor disc 70.
In another embodiment of the assembly 100, the annular flat disc 80 at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first non-aerofoil section 112, the second non-aerofoil section 212, and a side opening of the common cavity 99 thereinbetween. The common cavity 99 may have two side openings - one on each side 78,79 of the rotor disc 70. In a related embodiment, the assembly 100 includes two such annular flat discs 80 - one such annular flat disc 80 adhered on each side surface 71,72 of the rim section 73 of the rotor disc 70. Each of the two annular flat discs 80 at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first non-aerofoil section 112, the second non-aerofoil section 212, and the side opening of the common cavity 99 thereinbetween on either side 78,79 of the disc 70. The annular flat disc 80 may be aerodynamically shaped for example by smoothening or flattening out a surface 82 of the annular flat disc 80. The annular flat disc 80 may have curved parts of the surface corresponding to surface contours of the rim section 73 to maintain the aerodynamic shape.
The present technique also presents a method 500 for manufacturing a turbine blade assembly 100, as depicted by a flow chart of FIG 13. The method 500 for manufacturing the turbine blade assembly 100, hereinafter also referred to as the method 500, of the present technique has been explained hereinafter in combination with FIGs 1 to 12. In the method 500, in a step 510 the first blade root 120 and the second blade root 220 of the first and the second turbine blades 110,120, respectively, are inserted adjoining to each other 110,120 into the rim section 73 of the rotor disc 70. As a result of the step 510 the common cavity 99 between the first non-aerofoil section 112 of the first turbine blade 110 and the second non-aerofoil section 212 of the second turbine blade 210 is defined. Thereafter in the method 500, in a step 520 the adhesive material is poured into the common cavity 99 such that the adhesive material adheres to the inner surface 98 of the common cavity 99. The adhesive material may be poured through the gap 3 or through one or more holes (not shown) in the first and/or the second platforms 114,214. The adhesive material cures to form the sealing element 60 located within the common cavity 99. The adhesive material is poured into the common cavity 99 such that the sealing element 60 so formed by the pouring 520, and subsequent curing, of the adhesive material extends between the first and the second non-aerofoil sections 112,212 and thereby at least partially fills the common cavity 99. The adhesive material may be, but not limited to, ceramic or graphite composite same as aforementioned.
In an embodiment of the method 500, before the step 520, in a step 515 the rotor disc 70 having the first and the second turbine blades 110,210 inserted therewithin, is laid on a flat surface such that the second side 79 of the rotor disc 70 is positioned on the flat surface and such that the side opening of the common cavity 99 on the second side 79 of the rotor disc 70 is closed or sealed off by the flat surface. The second side opening of the common cavity 99 is closed by the flat surface such that the adhesive is sealed off from flowing out of the common cavity 99 via the second side opening of the common cavity 99. In this embodiment of the method 500, the adhesive material is poured 520 into the common cavity 99 from the first side 78 of the rotor disc 70, and through the first side opening of the common cavity 99. Furthermore, in the method 500, after the step 520, in a step 525 the rotor disc 70 is removed from the flat surface after the adhesive material is cured or solidified to form the sealing element 60.
In another embodiment of the method 500, in a step 530 the adhesive material is spread in form of the annular flat disc 80 on one of the side surfaces 71,72 of the rim section 73 of the rotor disc 70, such that the adhesive material at least partially covers, by adhesion, the first and the second blade roots 120,220 and the rim section 73 of the rotor disc 70 thereinbetween. The annular flat disc 80 of adhesive material during curing or solidification adheres onto the first and the second blade roots 120,220 and the rim section 73 of the rotor disc 70 thereinbetween. In a related embodiment of the method 500, the step 530 is performed twice to form one annular flat disc 80 adhered on each side surface 71,72 of the rim section 73 of the rotor disc 70.
In the method 500, in the step 530 the adhesive material is spread such that the adhesive material at least partially covers, by adhesion, the first non-aerofoil section 112, the second non-aerofoil section 212, and one of the side openings of the common cavity 99 thereinbetween. In another embodiment of the method 500, the adhesive material is spread on both sides sequentially in shape of the annular flat disc 80, such that the adhesive material adheres or sticks on each side surface 71,72 of the rim section 73 of the rotor disc 70 and the annular flat discs 80 are formed on either sides or faces 78,79 of the rotor disc 70. Each of the annular flat discs 80 at least partially, and preferably completely, covers or superimposes on, by adhesion onto, the first non-aerofoil section 112, the second non-aerofoil section 212, and the side opening of the common cavity 99 thereinbetween.
In another embodiment of the method 500, after the step 530, in a step 535 the adhesive material so spread is shaped into an aerodynamic form. This may be achieved for example by flattening of or smoothening of the surface 82 of the adhesive material so spread by pressing onto the surface 82 a flat sheet or a shaped sheet. The aerodynamic form of the annular flat disc 80 may be understood as has been explained hereinabove with respect to FIGs 11 and 12.
In the method 500, optionally before the step 520, a surface preparation is performed for the inner surface 98 or the surfaces that would form the inner surface 98 of the common cavity 99, for example by cleaning the inner surface 98 by chemicals or by sand blasting the inner surface 98. This helps in adhesion of the adhesive material onto the inner surface 98 of the common cavity 99.
While the present technique has been described in detail with reference to certain embodiments, it should be appreciated that the present technique is not limited to those precise embodiments. Rather, in view of the present disclosure which describes exemplary modes for practicing the invention, many modifications and variations would present themselves, to those skilled in the art without departing from the scope and spirit of this invention. The scope of the invention is, therefore, indicated by the following claims rather than by the foregoing description. All changes, modifications, and variations coming within the meaning and range of equivalency of the claims are to be considered within their scope.

Claims

A turbine blade assembly (100) for a gas turbine (10), the turbine blade assembly (100) comprising:
- a rotor disc (70) comprising a rim section (73);

- a first turbine blade (110) having a first aerofoil section (117) and a first non-aerofoil section (112), the first non-aerofoil section (112) comprising a first platform (114), a first front face (115), a first back face (113), a first shank (401) and a first blade root (120), wherein the first turbine blade (110) is mounted on the rotor disc (70) by the first blade root (120) inserted into the rim section (73) of the rotor disc (70);

- a second turbine blade (210) having a second aerofoil section (217) and a second non-aerofoil section (212), the second non-aerofoil section (212) comprising a second platform (214), a second front face (215), a second back face (213), a second shank (402) and a second blade root (220), wherein the second turbine blade (210) is mounted on the rotor disc (70) by the second blade root (220) inserted into the rim section (73) of the rotor disc (70);

- wherein the first and the second turbine blades (110,210) are mounted on the rotor disc (70) such that the first non-aerofoil section (112) adjoins the second non-aerofoil section (212) and wherein the first and the second non-aerofoil sections (112, 212) define a common cavity (99) thereinbetween; and

- a sealing element (60) located inside the common cavity (99), wherein the sealing element (60) is adhered to an inner surface (98) of the common cavity (99) and extends between the first and the second non-aerofoil sections (112, 212) to at least partially fill the common cavity (99), and wherein the sealing element (60) comprises an adhesive material.
The turbine blade assembly (100) according to claim 1, wherein the first shank (401) of the first turbine blade (110) and the second shank (402) of the second turbine blade (210) define the common cavity (99).
The turbine blade assembly (100) according to claim 1 or 2, wherein the first shank (401) comprises a first cavity (111) and/or the second shank (402) comprises a second cavity (211), and wherein the common cavity (99) comprises the first cavity (111) and/or the second cavity (211).
The turbine blade assembly (100) according to any of claims 1 to 3, wherein the inner surface (98) of the common cavity (99) comprises a rim surface (74) of the rotor disc (70) .
The turbine blade assembly (100) according to any of claims 1 to 4, further comprising at least an annular flat disc (80) adhered on a side surface (71,72) of the rim section (73) of the rotor disc (70), wherein the annular flat disc (80) at least partially covers the first and the second non-aerofoil sections (112, 212) and the rim section (73) of the rotor disc (70) thereinbetween, and wherein the annular flat disc (80) comprises the adhesive material.
The turbine blade assembly (100) according to claim 5, wherein the annular flat disc (80) is adhered onto the first non-aerofoil section (112), the second non-aerofoil section (212), and a side opening of the common cavity (99) thereinbetween.
The turbine blade assembly (100) according to claim 5 or 6, wherein the annular flat disc (80) is aerodynamically shaped.
The turbine blade assembly (100) according to any of claims 1 to 7, wherein the adhesive material is a ceramic or a graphite composite.
A gas turbine (10) comprising a turbine blade assembly (100), wherein the turbine blade assembly (100) is according to any of the claims 1 to 8.
A method (500) for manufacturing a turbine blade assembly (100), the method (500) comprising:
- inserting (510), adjoining to each other, a first and a second blade roots (120,220) of a first and a second turbine blade (110, 210), respectively, into a rim section (73) of a rotor disc (70) to define a common cavity (99) between a first non-aerofoil section (112) of the first turbine blade (110) and a second non-aerofoil section (212) of the second turbine blade (210),
wherein the first non-aerofoil section (112) comprises a first platform (114), a first front face (115), a first back face (113), a first shank (401) and the first blade root (120), and wherein the second non-aerofoil section (212) comprises a second platform (214), a second front face (215), a second back face (213), a second shank (402) and the second blade root (220); and

- pouring (520) an adhesive material into the common cavity (99) such that the adhesive material adheres to an inner surface (98) of the common cavity (99) and cures to form a sealing element (60) located within the common cavity (99) such that the sealing element (60) extends between the first and the second non-aerofoil sections (112, 212) to at least partially fill the common cavity (99).
The method (500) according to claim 10, comprising:
- laying (515) the rotor disc (70), with the first and the second turbine blades (110,120) so inserted, on a flat surface such that a second side (79) of the rotor disc (70) is positioned on the flat surface and such that a second side opening of the common cavity (99) is closed by the flat surface,
wherein the adhesive material is poured (520) into the common cavity (99) from a first side (78) of the rotor disc (70) and through a first side opening of the common cavity (99); and wherein the method (500) further comprises:
- removing (525) the rotor disc (70) from the flat surface after the adhesive material is cured.
The method (500) according to claim 10 or 11, comprising:
- spreading (530) the adhesive material in form of an annular flat disc (80) on a side surface (71,72) of the rim section (73) of the rotor disc (70), such that the adhesive material at least partially adheres onto the first and the second non-aerofoil sections (112, 212) and the rim section (73) of the rotor disc (70) thereinbetween.
The method (500) according to claim 12, wherein the adhesive material is spread (530) in form of the annular flat disc (80) such that the adhesive material adheres onto the first non-aerofoil section (112), the second non-aerofoil section (212), and a side opening (91, 92) of the common cavity (99) thereinbetween.
The method (500) according to claim 12 or 13, further comprising:
- shaping (535) the adhesive material so spread (530) into an aerodynamic form.
The method (500) according to any of claims 10 to 14, wherein the adhesive material is a ceramic or a graphite composite.