Techniques for cooling inner shrouds of gas turbine vanes

Abstract

The present technology proposes a bucket for a gas turbine. The bucket has an inner shroud with a pressure wall side, a suction wall side, a leading edge side, and a trailing edge side. A seal unit is disposed at a lower surface of the inner shroud and defines a first region and a second region at the lower surface of the inner shroud. A first impingement unit is arranged in the first region and has a first impingement plate facing the inner shroud defining a first impingement chamber therebetween, wherein the first impingement plate receives cooling air and emits an impingement jet into the first impingement chamber. A second impingement unit is arranged in the second region and includes a second impingement plate facing the inner shroud defining a second impingement chamber therebetween. At least one connector flow passage directs cooling air from the first impingement chamber to the second region. The second impingement plate receives cooling air from the connector flow passage and forms an impingement jet directed into the second impingement chamber.

Description

TECHNICAL FIELD

[0001] The present invention relates to a gas turbine, in particular to cooling of a gas turbine bucket, and more particularly to a technology for cooling an inner shroud of a gas turbine bucket. BACKGROUND

[0002] A gas turbine bucket, also referred to as nozzle section, comprises an airfoil extending radially between an inner shroud and an outer shroud, also referred to as inner platform and outer platform. The inner shroud and the outer shroud define a part of the hot gas flow path through the turbine section and are immersed in hot gas during turbine operation. Cooling of the shrouds is therefore necessary.

[0003] Figure 2 A conventional cooling scheme for an inner shroud 100 of a conventional turbine bucket 44 is schematically illustrated. The turbine bucket 44, also referred to as bucket 44 in the following, is positioned in an annular hot gas path 55 downstream of a transition duct 17 extending from a combustion chamber of a gas turbine, the hot gas path 55 being defined in part by an inner shroud 100 and an outer shroud 90 of the bucket 44. An airfoil 60 of the bucket 44 extending between the shrouds 90, 100 is arranged in the hot gas path 55. The outer shroud 90 is arranged towards an outer casing of the stator 42 and defines a radially outer surface 52 of the hot gas path 55, while the inner shroud 100 is arranged radially inwards towards a central axis, i.e. a rotational axis, of the gas turbine and defines a radially inner surface 54 of the hot gas path 55.

[0004] Hot gas or combustion products 34 flow from the combustion chamber via the transition duct 17 into the hot gas path 55, so that the airfoil 60 as well as surfaces of the shrouds 90, 100, e.g. an upper surface 100a of the inner shroud 100, are immersed in the hot gas 34. Cooling air 5 is guided from a compressor into the bucket 44 to cool the bucket 44.

[0005] Figure 2A first flow passage 9a is schematically depicted through which a portion 5a of the cooling air 5 is directed towards the inner shroud 100, in particular towards the lower surface 100b of the inner shroud 100. At the lower surface 100b of the inner shroud 100 a sealing unit 70, also referred to as axial seal 70, is provided to maintain a proper pressure of the cooling air 5 in the first region R1 which is axially upstream of the second region R2 at the lower surface 100b of the inner shroud 100. The sealing unit 70 can be formed as an annular plate extending around the central axis of the gas turbine and sealing or blocking or limiting the flow of cooling air 5a towards the second region R2. The sealing unit 70 is referred to as axial seal because the sealing unit 70 is configured to block the flow of cooling air 5 in axial direction. The sealing unit 70 is required to maintain a proper pressure in the first region R1 so that a sufficient flow of cooling air can be maintained to different portions of the vane 44.

[0006] Figure 2 A second flow passage 9b is also schematically depicted through which a portion 5b of the cooling air 5 is directed towards the turbine blade 38 downstream of the vane 44 via a cooling air passage 36c formed through the blade carrier disc 36. An inter-stage seal 80, also referred to as radial seal, such as a labyrinth seal 80, is provided between the vane 44 and the disc 36 and is configured to block the flow of cooling air 5 in radial direction between the vane stage comprising the vane 44 and the blade stage comprising the blade 38 and the blade carrier disc 36.

[0007] From Figure 2 It can be seen that a portion of the inner shroud 100 located in the second region R2 is not properly cooled.

[0008] It is therefore an object of the present application to provide a mechanism or technique for effectively cooling an inner shroud of a gas turbine vane. Preferably, it is desired that the technique for cooling an inner shroud of a gas turbine vane enhances the cooling without increasing the amount of cooling air extracted from the compressor for such cooling. SUMMARY

[0009] One or more of the above objects are achieved by a turbine vane for a gas turbine, a turbine vane assembly or turbine vane stage for a gas turbine and a gas turbine according to the independent claims appended hereto. Advantageous embodiments of the present technology are provided in the dependent claims. The features of the independent claims can be combined with the features of the claims dependent thereon, and the features of the dependent claims can be combined with each other.

[0010] In a first aspect of the present technology, a turbine vane for a gas turbine is presented. The turbine vane comprises an inner shroud, a sealing unit, a first impingement unit, a second impingement unit, and at least one connector flow passage.

[0011] The inner shroud has an upper surface and a lower surface. The turbine vane can comprise an airfoil extending from the upper surface. The turbine vane can further comprise an outer shroud, and the airfoil can extend radially between the inner shroud and the outer shroud.

[0012] A "radial" direction as used in the present disclosure can be understood as a radial direction with respect to a rotational axis or central axis of the gas turbine. Similarly, terms such as "radially inner" or "radially inward" as used in the present disclosure can be understood as in a radial direction towards the rotational axis or central axis of the gas turbine; wherein terms such as "radially outer" or "radially outward" as used in the present disclosure can be understood as in a radial direction away from the rotational axis or central axis of the gas turbine. Further, a radial distance can be understood as a distance measured along the radial direction.

[0013] The inner shroud can be understood as a radially inner shroud or radially inner platform with respect to the rotational axis of the gas turbine or simply as an inner platform or lower shroud or lower platform. The inner shroud forms a radially inner surface of an annular hot gas flow path of the turbine. The outer shroud can be understood as a radially outer shroud or radially outer platform with respect to the rotational axis of the gas turbine or simply as an outer platform or upper platform or upper shroud. The outer shroud forms a radially outer surface of the annular hot gas flow path of the turbine. The inner shroud and the outer shroud define or limit the hot gas flow path, i.e. the annular shape of the gas flow path through which the combustion products flow in the turbine section of the gas turbine.

[0014] The airfoil has a pressure wall and a suction wall meeting at a leading edge and a trailing edge. The inner shroud can comprise a pressure wall side, a suction wall side, a leading edge side, and a trailing edge side.

[0015] The pressure wall side, the suction wall side, the leading edge side, and the trailing edge side of the inner shroud can correspond to the pressure wall, the suction wall, the leading edge, and the trailing edge of the airfoil.

[0016] The sealing unit is provided at the lower surface of the inner shroud and defines a first region and a second region at the lower surface of the turbine vane. The sealing unit seals the first region with respect to the second region, i.e. blocks or limits or prevents a flow of cooling air from the first region to the second region.

[0017] The sealing unit can be an axial seal and can seal or block or limit or prevent a flow of cooling air from a first region to a second region in an axial direction of the gas turbine. The first region can be axially upstream of the second region, and the first region and the second region can be separated from each other by the sealing unit.

[0018] The sealing unit can extend in a circumferential direction around a central axis of the gas turbine.

[0019] The sealing unit can be arranged between the leading edge side and the trailing edge side, i.e. spaced apart from the leading edge side and the trailing edge side and extending from the pressure wall side to the suction wall side of the inner shroud. Thus, the sealing unit defines a first region between the leading edge side and the sealing unit and a second region between the sealing unit and the trailing edge side at the lower surface of the inner shroud.

[0020] The first impingement unit is arranged in the first region and comprises a first impingement plate facing the lower surface of the inner shroud. The first impingement plate is radially spaced apart from the lower surface of the inner shroud to define a first impingement chamber between the lower surface of the inner shroud and the first impingement plate. The first impingement unit can be arranged in radial alignment with the first portion of the inner shroud. The first impingement unit can face the lower surface of the first portion of the inner shroud and be spaced apart from the lower surface of the first portion of the inner shroud in a radial direction.

[0021] The first impingement plate can comprise a plurality of impingement holes or first impingement holes, i.e. through holes for generating impingement jets. The first impingement plate receives cooling air, preferably from the compressor, and subsequently forms impingement jets due to the cooling air passing through the impingement holes, which are directed into the first impingement chamber.

[0022] The second impingement unit is arranged in the second region and comprises a second impingement plate facing the lower surface of the inner shroud. The second impingement plate is radially spaced apart from the lower surface of the inner shroud to define a second impingement chamber between the lower surface of the inner shroud and the second impingement plate. The second impingement unit can be arranged in radial alignment with the second portion of the inner shroud. The second impingement unit can face the lower surface of the second portion of the inner shroud and be spaced apart from the lower surface of the second portion of the inner shroud in a radial direction.

[0023] The first portion and the second portion of the inner shroud can be understood as parts or portions of the inner shroud which are divided or distinguished or limited by the sealing unit when viewed in a radial direction. In other words, if the sealing unit were to extend hypothetically in a radial direction, such a hypothetical extension would divide the inner shroud into the first portion and the second portion. The first portion is arranged upstream of the second portion with respect to the hot gas flow path.

[0024] A distance between the first impingement plate and the lower surface of the inner shroud, i.e. the target cooling surface, can be between 1 mm (millimeter) and 4 mm.

[0025] The distance between the second impingement plate and the lower surface of the inner shroud, i.e. the target cooling surface, can be between 1 mm (millimeter) and 4 mm.

[0026] The first and second impingement units can be separated from each other or can be spaced apart from each other by a sealing unit.

[0027] The at least one connector flow passage can have an inlet positioned at the first impingement chamber so as to receive cooling air from the first impingement chamber and an outlet positioned at the second region so as to direct cooling air from the first impingement chamber to the second region.

[0028] The connector flow passage can be formed as a tubular structure or pipe or hollow duct or tube or through hole.

[0029] The second impingement plate can comprise a plurality of impingement holes or second impingement holes, i.e. through holes for generating impingement jets. The second impingement plate receives cooling air from the connector flow passage, i.e. cooling air that enters the inlet of the connector flow passage at the first region or first impingement chamber and exits the outlet of the connector flow passage at the second region. The second impingement plate then forms impingement jets due to the cooling air passing through the impingement holes, which are directed into the second impingement chamber.

[0030] Thus, the cooling air in the first region for impingement onto the lower surface of the inner shroud is then directed into the second region and then flows in the form of impingement jets through the second impingement plate into the second impingement chamber. Thus, the same cooling air is used for two impingement coolings - first at the lower surface of the inner shroud in the first region, i.e. cooling a first portion of the inner shroud, and thereafter at the lower surface of the inner shroud in the second region, i.e. cooling a second portion of the inner shroud.

[0031] For further explanation, the cooling air flows in the following order: towards the first impingement plate, through the first impingement holes and thus generating impingement jets into the first impingement chamber towards the inner shroud, then flows from the first impingement chamber through the connector flow passage into the second region, then towards the second impingement plate, then through the second impingement holes and thus generating impingement jets into the second impingement chamber towards the inner shroud.

[0032] The lower surface of the inner shroud can be radially recessed in the first region and thus can comprise a first impingement cavity, and the first impingement chamber can comprise the first impingement cavity.

[0033] The first impingement plate can be positioned at or within the first impingement cavity, thus making the arrangement compact.

[0034] The lower surface of the inner shroud can be radially recessed in the second region and can thus comprise a second impact cavity, and the second impact chamber can comprise the second impact cavity.

[0035] The second impact plate can be positioned at or within the second impact cavity, thereby making the arrangement compact.

[0036] The first impact cavity and the second impact cavity can be separated by an intervening section or wall of the inner shroud. The intervening section or wall of the inner shroud can extend radially inwardly from the lower surface of the inner shroud. The connector flow passage extends through the intervening section of the inner shroud. In other words, the connector flow passage can be formed as a through-hole in the intervening section or wall of the inner shroud.

[0037] The second impact plate can be arranged flush with the opening of the second impact cavity. This makes the arrangement even more compact.

[0038] The second impact plate can be arranged within the second impact cavity. This makes the arrangement even more compact.

[0039] The connector flow passage can extend through the sealing unit.

[0040] The sealing unit can comprise at least one of a sealing support ledge or member and a sealing plate.

[0041] The sealing support ledge can extend radially inwardly from the lower surface of the inner shroud. The sealing support ledge can define the first region and the second region. The connector flow passage can extend through the sealing support ledge. The sealing plate can be supported at or fixed to the sealing support ledge. The connector flow passage can be formed as a through-hole through the sealing support ledge or can be formed as a separate tubular structure or tube that is inserted through the through-hole formed in the sealing support ledge.

[0042] The sealing plate can not have any connector flow passage formed through the sealing plate. Alternatively and optionally, a further connector flow passage can extend through the sealing plate.

[0043] The sealing plate can be arranged to extend radially inwardly from the lower surface of the inner shroud or from the sealing support ledge. The sealing plate can define the first region and the second region. The connector flow passage can extend through the sealing plate. The sealing plate can be supported at or fixed to the lower surface of the inner shroud or to the sealing support ledge. The connector flow passage can be formed as a through-hole through the sealing plate or can be formed as a separate tubular structure or tube that is inserted through the through-hole formed in the sealing plate.

[0044] The sealing support ledge can not have any connector flow passage formed through the sealing support ledge. Alternatively and optionally, a further connector flow passage can extend through the sealing support ledge.

[0045] The width of the connector flow channel is between 2% and 40% of the width of the sealing support lug or sealing plate measured along the circumferential direction of the inner cover, and preferably between 5% and 15%.

[0046] As used in this disclosure, the term "circumferential" can be understood as the circumferential direction relative to the axis of rotation or central axis of the gas turbine.

[0047] The second impact unit may include a cover plate disposed radially inside the second impact plate and facing the second impact plate, i.e., the cover plate may be disposed between the second impact plate and the rotation axis or central axis of the gas turbine. The cover plate may be spaced apart from the second impact plate in the radial direction, and thus may define a cooling air receiving chamber therebetween, i.e., between the second impact plate and the cover plate.

[0048] The outlet of the connector flow channel can be located at the cooling air receiving chamber.

[0049] The lower surface of the inner shield in the first region may include a base opening of the airfoil. The base opening can be understood as an opening into the internal space or cavity of the airfoil, or as an opening in fluid communication with the internal space or cavity of the airfoil. The base openings of the first impact chamber and the airfoil are not overlapping, discontinuous, or in fluid communication with each other. In other words, the first impact chamber and the base openings of the airfoil are not in fluid connection with each other, such that there is no flow connection between the first impact chamber and the base opening of the airfoil.

[0050] The radial distance between the second impact plate and the lower surface of the inner shield can be less than or equal to the radial distance between the first impact plate and the lower surface of the inner shield.

[0051] The diameter of the impact hole in the second impact plate can be smaller than the diameter of the impact hole in the first impact plate.

[0052] The inner shield may include at least one shield cooling hole or channel having an inlet located at the second impact chamber and an outlet located at the upper surface or side surface of the inner shield.

[0053] In a second aspect of this technology, a gas turbine including turbine blades is proposed. The turbine blades can be configured according to the first aspect of this technology as described above.

[0054] A gas turbine may include turbine blades located downstream of the turbine impeller relative to the direction of hot air flow. The gas turbine may also include interstage seals extending between the turbine blades and the turbine impeller. Interstage seals can be understood as seals extending between the turbine blades and the turbine impeller, such as labyrinth seals.

[0055] The second impingement unit can be positioned within a space defined by an inner shroud of the turbine vane, the sealing unit, and an inter-stage seal. The inter-stage seal is configured to seal the space at a radially inner side of the space.

[0056] The first impingement plate can be configured to receive cooling air from a last stage or final stage of the compressor. BRIEF DESCRIPTION OF DRAWINGS

[0057] The above attributes of the technology, as well as other features and advantages of the technology, and the manner of attaining them, will become more apparent, and the technology itself will be better understood by reference to the following description of embodiments of the technology taken together with the accompanying drawings, wherein:

[0058] Figure 1 A cross-sectional view of a portion of an exemplary embodiment of a gas turbine showing an exemplary embodiment of a turbine vane that can incorporate the technology is shown;

[0059] Figure 2 A conventional cooling scheme for an inner shroud of a conventional turbine vane is schematically illustrated;

[0060] Figure 3 A cooling scheme for an inner shroud of a turbine vane of the technology is schematically illustrated;

[0061] Figure 4 A bottom view of an exemplary embodiment of a turbine vane of the technology is schematically depicted;

[0062] Figure 5 A cross-sectional view at line I-I of a turbine vane of the technology is depicted; Figure 4

[0063] Figures 6A-6C A cross-sectional view of different exemplary embodiments of a turbine vane of the technology is schematically depicted;

[0064] Figure 7 An exemplary embodiment of a cooling air flow within an inner shroud of a turbine vane of the technology is schematically depicted;

[0065] Figure 8 A cross-sectional view of yet another exemplary embodiment of a turbine vane of the technology is schematically depicted, which depicts an exemplary arrangement of a first impingement unit and a second impingement unit;

[0066] Figures 9A-9B A cross-sectional view of an exemplary embodiment of a turbine vane of the technology is schematically depicted, which depicts a shroud cooling hole;

[0067] Figure 10 ​A bottom cross-sectional view of an exemplary embodiment of a turbine bucket of the present technology is schematically depicted, depicting a base opening of an airfoil of the turbine bucket and the relative positioning of a first impingement chamber and a second impingement chamber of the turbine bucket of the present technology; and

[0068] Figures 11A-11C A cross-sectional view of a further exemplary embodiment of a turbine bucket of the present technology is schematically depicted. DETAILED DESCRIPTION

[0069] Hereinafter, the above-described features and other features of the present technology are described in detail. Various embodiments are described with reference to the drawings, wherein like reference numerals are used throughout to designate like elements. In the following description, for the purposes of explanation, numerous specific details are set forth in order to provide a thorough understanding of one or more embodiments. It can be noted that the illustrated embodiments are intended not as limitation but rather a illustration of possible embodiments. It will be apparent to one skilled in the art that embodiments can be practiced without these specific details.

[0070] Figure 1 An example of a gas turbine or gas turbine engine 10 is shown in cross-section. The gas turbine 10 can include, in flow series, an inlet 12, a compressor or compressor section 14, a combustor section 16, and a turbine section 18, generally arranged in flow series and generally about and along the direction of a longitudinal or rotational axis 20. The gas turbine 10 can also include a shaft 22 that is rotatable about the rotational axis 20 and that extends longitudinally through the gas turbine 10. The shaft 22 can drivingly connect the turbine section 18 to the compressor section 14.

[0071] During operation of the gas turbine 10, air 24 drawn in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or combustor section 16. The combustor section 16 can include a combustor plenum 26, one or more combustion chambers 28, and at least one combustor 30 secured to each combustion chamber 28. The combustion chambers 28 and combustors 30 can be located inside the combustor plenum 26. Compressed air from the compressor 14 can enter a diffuser 32 and can be discharged from the diffuser 32 into the combustor plenum 26, from which a portion of the air can enter the combustor 30 and mix with a gaseous or liquid fuel. The air / fuel mixture then combusts, and the combustion gases 34 or working gases resulting from the combustion are directed through the combustion chamber 28 to the turbine section 18 via a transition duct 17.

[0072] The example gas turbine 10 can have an annular combustor section arrangement 16 comprised of an annular array of combustor cans 19 each having a combustor 30 and a combustion chamber 28, with a transition duct 17 having a generally circular inlet interfacing with the combustion chamber 28 and an outlet in the form of an annular segment. The annular array of transition duct outlets can form an annulus for directing combustion gases to the turbine 18.

[0073] The turbine section 18 can include a plurality of blade carrying disks 36 attached to the shaft 22. In this example, two disks 36 are depicted each carrying an annular array of turbine blades 38. However, the number of blade carrying disks can vary, i.e. only one disk or more than two disks. Further, between annular arrays of turbine blades 38 at various stages can be provided guide vanes 40 fixed to a stator 42 of the gas turbine 10. Between the outlet of the combustion chamber 28 and the forward turbine blades 38 can be provided inlet guide vanes 44, and the inlet guide vanes 44 divert the working gas flow onto the turbine blades 38.

[0074] The combustion gases from the combustion chamber 28 enter the turbine section 18 and drive the turbine blades 38, which in turn rotate the shaft 22. The guide vanes 40, 44 serve to optimize the angle of the combustion or working gases on the turbine blades 38.

[0075] The turbine section 18 drives the compressor section 14. The compressor section 14 can include an axial series of vane stages 46 and rotor blade stages 48. The rotor blade stages 48 can include rotor disks supporting blades in an annular array. The compressor section 14 can also include a casing 50 surrounding the rotor stages and supporting the vane stages 48. The guide vane stages can include radially extending vanes in an annular array mounted to the casing 50. These vanes are provided to give the blades the gas flow at the optimum angle at a given engine operating point. Some of the guide vane stages can have variable vanes, where the angle of the vanes about their own longitudinal axis can be adjusted according to the gas flow characteristics that can occur at different engine operating conditions. The casing 50 can define a radially outer surface 52 of a passageway 56 of the compressor 14. A radially inner surface 54 of the passageway 56 can be at least partially defined by a rotor drum 53 of the rotor, which can be partially defined by the blades 48 in an annular array.

[0076] The technology is described with reference to the above example gas turbine having a single shaft or spool connecting a single multi-stage compressor and a single stage or more turbine. However, it will be appreciated that the technology is equally applicable to engines having two or three shafts and can be used in industrial, aeronautical or marine applications.

[0077] Unless otherwise stated, the terms "upstream" and "downstream" refer to the direction of the air and / or working gas flow through the gas turbine. The terms "forward" and "aft" refer to the general flow of hot gas through the gas turbine. Unless otherwise stated, the terms "axial", "radial" and "circumferential" are defined with reference to the rotational or central axis 20 of the gas turbine.

[0078] In the present technology, a turbine vane 1 is presented, which comprises an inner shroud 100 and an outer shroud 90 and has an airfoil 60 extending between the inner shroud 100 and the outer shroud 90, for example as shown in Figure 3 、 Figure 4 and Figures 6A-6C . The turbine vane 1 of the present technology can be a vane 40, 44 of the gas turbine 10 described above.

[0079] Figure 3 A cooling scheme for the inner shroud of the turbine vane 1 of the present technology is schematically illustrated and can be understood by comparison with Figure 2 . The explanations provided in relation to Figure 3 apply to Figure 2 . Figure 3 .

[0080] Figure 3 The inner shroud 100 and the cooling mechanism or scheme of the turbine vane 1 of the present technology are schematically illustrated in connection with Figure 4 、 Figure 6A 、 Figure 6B and Figure 6C . It can be noted that Figure 6A 、 Figure 6B and Figure 6C and other figures such as Figure 8 、 Figure 9A and Figure 9B are schematic representations and depictions of the cooling mechanism, for example the size of the impingement units and the distance between the impingement units and the target surface of the impingement cooling are exaggerated compared to other parts of the vane in order to facilitate understanding.

[0081] The turbine vane 1 is also referred to as vane 1 or nozzle portion 1 or guide vane 1 or inlet nozzle portion 1 or inlet guide vane 1 hereinafter, which comprises an inner shroud 100, a sealing unit 70, a first impingement unit 110, a second impingement unit 120 and at least one connector flow passage 130.

[0082] The inner shroud 100 has an upper surface 100a and a lower surface 100b. The turbine vane 1 can comprise an airfoil 60 extending from the upper surface 100a. The turbine vane can further comprise an outer shroud 90 and the airfoil 60 can extend radially between the inner shroud 100 and the outer shroud 90.

[0083] The inner shroud 100 can be understood as a radially inner shroud 100 with respect to the rotational axis 20 of the gas turbine 10. The inner shroud 100 forms a radially inner surface 54 of an annular hot gas flow path 52 of the turbine 10. The outer shroud 90 can be understood as a radially outer shroud 90 with respect to the rotational axis 20 of the gas turbine 10. The outer shroud forms a radially outer surface 54 of an annular hot gas flow path 55 of the turbine 10. The inner shroud 100 and the outer shroud 90 define or limit the hot gas flow path 55, i.e. the annular shape of the gas flow path through which the combustion products 34 flow in the turbine section 18 of the gas turbine 10.

[0084] An upper surface 100a of the inner shroud 100 faces the hot gas path 55, while a lower surface 100b of the inner shroud 100 faces the rotational axis 20 of the gas turbine 10. The upper surface 100a and the lower surface 100b of the inner shroud 100 are radially spaced apart and face opposite directions.

[0085] As shown in Fig. 1, the airfoil 60 has a pressure wall 62 and a suction wall 64 meeting at a leading edge 66 and a trailing edge 68. The inner shroud 100 can comprise a pressure wall side 102, a suction wall side 104, a leading edge side 106 and a trailing edge side 108. Figure 4 The pressure wall side 102, the suction wall side 104, the leading edge side 106 and the trailing edge side 108 of the inner shroud 100 can correspond to the pressure wall 62, the suction wall 64, the leading edge 66 and the trailing edge 68 of the airfoil 60.

[0086] The sealing unit 70 is arranged at the lower surface 100b of the inner shroud 100 and defines a first region R1 and a second region R2 at the lower surface 100b of the turbine vane 1. The sealing unit 70 seals the first region R1 with respect to the second region R2 in terms of a flow of cooling air 5 from the first region R1 to the second region R2.

[0087] The turbine vane 1 and the sealing unit 70 can be referred to as a turbine vane arrangement. The turbine vane arrangement can further comprise an inter-stage seal which will be explained later.

[0088] The sealing unit 70 can be arranged between the leading edge side 106 and the trailing edge side 108, i.e. spaced apart from the leading edge side 106 and the trailing edge side 108 and extending from the pressure wall side 102 to the suction wall side 104 of the inner shroud 100. Thus, the sealing unit 70 defines at the lower surface 100b of the inner shroud 100 a first region R1 between the leading edge side 106 and the sealing unit 70 and a second region R2 between the sealing unit 70 and the trailing edge side 108.

[0089]

[0090] ​More specifically, the sealing unit 70 defines, at the lower surface 100b of the inner shroud 100, a first region R1 between the leading edge side 106, the pressure wall side 102, the suction wall side 104 and the sealing unit 70 and a second region R2 between the sealing unit 70, the trailing edge side 108, the pressure wall side 102 and the suction wall side 104.

[0091] The upper surface or limit of the first region R1 and of the second region R2 can be defined by the lower surface 100b of the inner shroud 100.

[0092] From Figures 6A-6C It can be seen that the first impingement unit 110 is arranged in the first region R1 and comprises a first impingement plate 112 facing the lower surface 100b of the inner shroud 100. The first impingement plate 112 is radially spaced apart from the lower surface 100b of the inner shroud 100. A first impingement chamber 110c is defined in the first region R1 between the lower surface 100b of the inner shroud 100 and the first impingement plate 112.

[0093] The first impingement plate 112 can comprise a plurality of impingement holes 112h or first impingement holes 112h, i.e. through holes for generating impingement jets. The first impingement plate 112 receives cooling air 5, preferably from the compressor section 14, and subsequently forms impingement jets due to the cooling air 5 passing through the first impingement holes 112h. The impingement jets thus formed are emitted or injected or directed into the first impingement chamber 110c. The impingement jets impinge on a surface of the inner shroud 100, preferably on the lower surface 100b of the inner shroud 100, thereby cooling the inner shroud 100, i.e. a first portion of the inner shroud 100 corresponding to the first region R1.

[0094] The second impingement unit 120 is arranged in the second region R2 and comprises a second impingement plate 122 facing the lower surface 100b of the inner shroud 100. The second impingement plate 122 is radially spaced apart from the lower surface 100b of the inner shroud 100. A second impingement chamber 120c is defined between the lower surface 100b of the inner shroud 100 and the second impingement plate 122.

[0095] The first impingement unit 110 and the second impingement unit 120 can be separate from each other or can be spaced apart from each other by the sealing unit 70 and / or by the intervening section 101 of the inner shroud 100.

[0096] The at least one connector flow passage 130 can have an inlet 132 (shown in Figure 7 ) positioned at the first impingement chamber 100c to receive cooling air 5 from the first impingement chamber 100c and an outlet 134 (shown in Figure 7 ) positioned at the second region R2 to direct cooling air 5 from the first impingement chamber 100c to the second region R2.

[0097] The connector flow channel 130 can be formed as a tubular structure, a tube, a through hole, or a hollow pipeline, i.e., a conduit.

[0098] The second impact plate 122 may include a plurality of impact holes 122h or a second impact hole 122h, i.e., a through hole for generating an impact jet. The second impact plate 122 receives cooling air 5 from the first impact chamber 110c via a connector flow channel 130. The second impact plate 122 then forms an impact jet as the cooling air 5 passes through the second impact holes 122h. The impact jet thus formed is launched, sprayed, or guided into the second impact chamber 120c. The impact jet impacts the surface of the inner shield 100, preferably the lower surface 100b of the inner shield 100, thereby cooling the inner shield 100, i.e., cooling a second portion of the inner shield 100 corresponding to the second region R2.

[0099] To further explain, such as Figure 7 As shown, the cooling air 5 flows in the following sequence: towards the first impact plate 112, then through the first impact hole 112h and thus generating an impact jet that is injected toward the inner shield 100 into the first impact chamber 110c, then flows into the inlet 132 of the connector flow channel 130 and through the connector flow channel 130 from the first impact chamber 110c into the second region R2, then flows out from the outlet 134 of the connector flow channel 130 located in the second region R2, then flows toward the second impact plate 122, and through the second impact hole 122h and thus generating an impact jet that is injected toward the inner shield 100 into the second impact chamber 120c.

[0100] Therefore, as Figure 3 As schematically shown, a portion 5a of the cooling air 5 is used to impact the lower surface 100b of the inner shield 100 at the first impact chamber 110c in the first region R1, and is then guided across the sealing unit 70, but not necessarily through the sealing unit 70, via the connector flow channel 130 into the second region R2, and then flows as an impact jet through the second impact plate 120 into the second impact chamber 120c. Thus, two impact coolings are performed using the same cooling air—first, impact cooling is performed at the lower surface 100b of the inner shield 100 in the first region R1, i.e., cooling the first part of the inner shield 100, and then impact cooling is performed at the lower surface 100b of the inner shield 100 in the second region R2, i.e., cooling the second part of the inner shield 100.

[0101] Since the lower surface 100b of the inner shroud 100 is cooled by impingement cooling at both the first region R1 and the second region R2, the cooling efficiency is improved. Moreover, the same portion or volume of cooling air is used for impingement cooling at both the first region R1 and the second region R2, so that no additional cooling air needs to be extracted from the compressor for separately cooling the portion of the inner shroud corresponding to the second region R2.

[0102] Figure 3 A second flow channel 9b is also schematically depicted, through which a portion 5b of the cooling air 5 is guided via a cooling air channel 36c formed through the vane carrier disc 36 towards a turbine vane 38 located downstream of the wheel vane 1. Between the wheel vane 1 and the vane 38, an inter-stage seal 80, also referred to as a radial seal, is provided, such as a labyrinth seal 80, and is configured to block a flow of cooling air 5 in a radial direction between a wheel vane stage comprising the wheel vane 1 and a vane stage comprising the vane 38 and the vane carrier disc 36.

[0103] The inter-stage seal 80 can be understood as a seal, such as a labyrinth seal 80, extending between the turbine vane carrier disc 36 and the turbine wheel vane 1.

[0104] The second impingement unit 120 can be positioned within a space 82 defined by the inner shroud 100 of the turbine wheel vane 1, the seal unit 70 and the inter-stage seal 80, and the platform of the vane 38 and the disc 36.

[0105] In the following, different exemplary embodiments of the wheel vane 1 are explained with reference to Figure 5 , Figure 6A , Figure 6B and Figure 6C .

[0106] As shown in Figure 5 , Figures 6A-6C , the first impingement unit 110 can be formed such that the first impingement chamber 110c, also referred to as first chamber 110c for short, has an inlet for cooling air via the first impingement hole 112h, preferably exclusively via the first impingement hole 112h, i.e. the first impingement chamber 110c has no further inlet for receiving cooling air than the first impingement hole 112h. However, the present technology is not limited to the above-described arrangement, and the first impingement chamber 110c can have an inlet in addition to the first impingement hole 112h.

[0107] The first impingement unit 110 can be formed such that the first impingement chamber 110c has an outlet for cooling air via the connector flow passage 130, preferably only via the inlet 132 of the connector flow passage 130, i.e. the first impingement chamber 110c has no other outlet for ejecting cooling air from the first chamber 110c than through the connector flow passage 130. However, the present technology is not limited to the above-described arrangement and the first impingement chamber 110c can have an outlet in addition to the inlet 132 of the connector flow passage 130, e.g. a cooling hole (not shown) extending from the first impingement chamber 110c into the hot gas path 55.

[0108] As shown in Figs. 1 and 2, the first impingement unit 110 can comprise a first impingement plate 112 arranged radially inside the first impingement chamber 110c and facing the first impingement chamber 110c, i.e. the first impingement plate 112 can be arranged between the first impingement chamber 110c and the rotational axis 20 or central axis 20 of the gas turbine 10. The first impingement plate 112 can be spaced apart from the first impingement chamber 110c with respect to the radial direction and can thus define a cooling air receiving chamber 112c therebetween, i.e. between the first impingement plate 112 and the first impingement chamber 110c. Figure 5 , Figures 6A-6C As shown in Figs. 1 and 2, the first impingement unit 110 can comprise a first impingement plate 112 arranged radially inside the first impingement chamber 110c and facing the first impingement chamber 110c, i.e. the first impingement plate 112 can be arranged between the first impingement chamber 110c and the rotational axis 20 or central axis 20 of the gas turbine 10. The first impingement plate 112 can be spaced apart from the first impingement chamber 110c with respect to the radial direction and can thus define a cooling air receiving chamber 112c therebetween, i.e. between the first impingement plate 112 and the first impingement chamber 110c.

[0109] As shown in Figs. 1 and 2, the first impingement unit 110 can comprise a first impingement plate 112 arranged radially inside the first impingement chamber 110c and facing the first impingement chamber 110c, i.e. the first impingement plate 112 can be arranged between the first impingement chamber 110c and the rotational axis 20 or central axis 20 of the gas turbine 10. The first impingement plate 112 can be spaced apart from the first impingement chamber 110c with respect to the radial direction and can thus define a cooling air receiving chamber 112c therebetween, i.e. between the first impingement plate 112 and the first impingement chamber 110c. Figure 5 , Figures 6A-6C As shown in Figs. 1 and 2, the first impingement unit 110 can comprise a first impingement plate 112 arranged radially inside the first impingement chamber 110c and facing the first impingement chamber 110c, i.e. the first impingement plate 112 can be arranged between the first impingement chamber 110c and the rotational axis 20 or central axis 20 of the gas turbine 10. The first impingement plate 112 can be spaced apart from the first impingement chamber 110c with respect to the radial direction and can thus define a cooling air receiving chamber 112c therebetween, i.e. between the first impingement plate 112 and the first impingement chamber 110c.

[0110] The second impingement chamber 120c can be arranged radially outside the cooling air receiving chamber 124c and radially aligned with the cooling air receiving chamber 124c. The second impingement chamber 120c and the cooling air receiving chamber 124c can be fluidly connected to each other only through the second impingement hole 122h.

[0111] The cooling air receiving chamber 124c can be configured such that it has an inlet for cooling air via the connector flow channel 130, preferably only via the connector flow channel 130; that is, the cooling air receiving chamber 124c has no other inlets for receiving cooling air except for the connector flow channel 130, and particularly the outlet 134 of the connector flow channel 130. However, the present invention is not limited to the above arrangement, and the cooling air receiving chamber 124c may have inlets other than the outlet 134 of the connector flow channel 130.

[0112] The cooling air receiving chamber 124c can be configured such that it has an outlet for cooling air via the second impact hole 122h, preferably only via the second impact hole 122h; that is, the cooling air receiving chamber 124c has no other outlet for injecting cooling air besides the second impact hole 122h. However, the present invention is not limited to the above arrangement, and the cooling air receiving chamber 124c may have an outlet other than the second impact hole 122h.

[0113] like Figure 7 As shown, the inlet 132 of the connector flow channel 130 can be located at the first impact chamber 110c, and the outlet 134 of the connector flow channel 130 can be located at the cooling air receiving chamber 124c.

[0114] like Figure 6B and Figure 6C as well as Figure 7 As schematically shown, the lower surface 100b of the inner shield 100 may include a first impact cavity C1 in the first region R1. The first impact chamber 110c may include the first impact cavity C1.

[0115] The first impact plate 112 can be positioned at the opening C11 of the first impact cavity C1, such as... Figure 7 As shown in the diagram. In other words, the first impact plate 112 can be formed as a flat sheet or plate or planar surface, and the first impact plate 112 can be flush with the opening C11 of the first impact cavity C1, that is, completely covering or closing or covering the opening C11 of the first impact cavity C1. Except for the cooling air flowing through the first impact hole 112h, the first impact plate 112 can completely seal the opening C11 of the first impact cavity C1. Alternatively, the first impact plate 112 can be positioned outside the first impact cavity C1, that is, radially spaced from the opening C11 of the first impact cavity C1, such as... Figure 6B and Figure 6CThe first impingement unit 110 can comprise a side plate member 112s or side part plate member 112s extending radially between the first impingement plate 112 and the lower surface 100b of the inner shroud 100, thereby enclosing the first impingement cavity C1 with the first impingement chamber 110c. The first impingement plate 112 and the side plate member 112s can completely seal the opening C11 of the first impingement cavity C1 except for cooling air flowing through the first impingement hole 112h.

[0116] In another embodiment (not shown), the first impingement plate 112 can be positioned inside or within the first impingement cavity C1, i.e. radially between the lower surface 100b of the inner shroud 100 and the opening C11 of the first impingement cavity C1. The first impingement plate 112 can completely seal the portion or volume or space of the first impingement cavity C1 arranged between the first impingement plate 112 and the lower surface 100b of the inner shroud 100 except for cooling air flowing through the first impingement hole 112h.

[0117] As shown in Figure 6B and Figure 6C and Figure 7 The lower surface 100b of the inner shroud 100 can comprise a second impingement cavity C2 in the second region R2, as schematically shown in

[0118] The second impingement plate 122 can be positioned at or within or outside the second impingement cavity C2.

[0119] As shown in Figure 6C The second impingement plate 122 can be positioned at the opening C21 of the second impingement cavity C2. In other words, the second impingement plate 122 can form a flat sheet or plate-like or planar and the second impingement plate 122 can be flush with the opening C21 of the second impingement cavity C2, i.e. completely cover or close or cover the opening C21 of the second impingement cavity C2. The second impingement plate 122 can completely seal the opening C21 of the second impingement cavity C2 except for cooling air flowing through the second impingement hole 122h.

[0120] Alternatively, as shown in Figure 6BAs shown in FIG. 1 1 1, the second impingement plate 122 can be positioned outside of the second impingement cavity C2, i.e. radially spaced apart from the opening C21 of the second impingement cavity C2. The second impingement unit 120 can comprise a side plate member 122s or side plate members 122s extending radially between the second impingement plate 122 and the lower surface 100b of the inner shroud 100, thereby enclosing the second impingement cavity C2 from the second impingement chamber 120c. The second impingement plate 122 and the side plate member 122s can completely seal the opening C21 of the second impingement cavity C2, except for cooling air flowing through the second impingement hole 122h.

[0121] In a further embodiment as shown in FIG. 1 1 1, the second impingement plate 122 can be positioned inside or within the second impingement cavity C2, i.e. radially between the lower surface 100b of the inner shroud 100 and the opening C21 of the second impingement cavity C2. The second impingement plate 122 can completely seal a portion or volume or space of the second impingement cavity C2 provided between the second impingement plate 122 and the lower surface 100b of the inner shroud 100, except for cooling air flowing through the second impingement hole 122h. Figure 7 The cover plate 124 can be positioned at the opening C21 of the second impingement cavity C2, as shown in FIG. 1 1 1. In other words, the cover plate 124 can be formed as a flat sheet or plate-like or planar, and the cover plate 124 can be flush with the opening C21 of the second impingement cavity C2, i.e. completely covering or closing or covering the opening C21 of the second impingement cavity C2.

[0122] Figure 7 The cover plate 124 can completely seal a portion or volume or space of the second impingement cavity C2 provided between the second impingement plate 122 and the cover plate 124 of the inner shroud 100, except for cooling air flowing through the second impingement hole 122h.

[0123] The cover plate 124 can completely seal a portion or volume or space of the second impingement cavity C2 provided between the second impingement plate 122 and the cover plate 124 of the inner shroud 100, except for cooling air flowing through the second impingement hole 122h.

[0124] As shown in FIG. 1 1 1, the connector flow passage 130 can extend through the sealing unit 70, preferably axially through the sealing unit 70. Figure 4 Figure 6A The sealing unit 70 can comprise at least one of a sealing support lug or member or wall 72 and a sealing plate 74.

[0125] The sealing unit 70 can comprise at least one of a sealing support lug or member or wall 72 and a sealing plate 74.

[0126] ​​The seal support ledge 72 can extend radially inward from the lower surface 100b of the inner shroud 100. The seal support ledge 72 can define the first region R1 and the second region R2. The connector flow passage 130 can extend through the seal support ledge 72. The seal plate 74 can be supported at or secured to the seal support ledge 72, preferably at a radially outer end or side of the seal plate 74. A radially inner end or side (not shown) of the seal plate 74 can be held or supported by a seal housing (not shown). The connector flow passage 130 can be formed as a through-hole through the seal support ledge 72, or can be formed as a separate tubular structure or tube that is inserted through a through-hole formed in the seal support ledge 72.

[0127] The seal plate 74 can not have any connector flow passages formed therethrough. Alternatively and optionally, additional connector flow passages (not shown) can extend through the seal plate 74.

[0128] The seal plate 74 can be arranged to extend radially inward from the lower surface 100b of the inner shroud 100 or from the seal support ledge 72. The seal plate 74 can define the first region R1 and the second region R2. The connector flow passage 130 can extend through the seal plate 74. The seal plate 74 can be supported at or secured to the lower surface 100b of the inner shroud 100 or to the seal support ledge 72, preferably at a radially outer end or side of the seal plate 74. A radially inner end or side (not shown) of the seal plate 74 can be held or supported by a seal housing (not shown). The connector flow passage 130 can be formed as a through-hole through the seal plate 74, or can be formed as a separate tubular structure or tube that is inserted through a through-hole formed in the seal plate 74.

[0129] The seal support ledge 72 can not have any connector flow passages formed therethrough. Alternatively and optionally, additional connector flow passages (not shown) can extend through the seal support ledge 72.

[0130] As Figure 5 And Figure 6BAs shown in The connector flow passage 130 can extend through the intervening section 101 of the inner shroud 100, preferably axially through the intervening section 101 of the inner shroud 100. In other words, the connector flow passage 130 can be formed as a through-hole in the intervening section 101 or the intervening wall 101 of the inner shroud 100. Alternatively, the connector flow passage 130 can be formed as a separate tubular structure or tube inserted through a through-hole formed in the intervening section 101 or the intervening wall 101 of the inner shroud 100.

[0131] The sealing unit 70, preferably at least one of the sealing support lug 72 and the sealing plate 74, and more preferably both the sealing support lug 72 and the sealing plate 74, can be aligned with the intervening section 101 or the intervening wall 101 of the inner shroud 100 in the radial direction.

[0132] With reference to Figure 4 exemplary dimensions of the connector flow passage 130 relative to the exemplary dimensions of the sealing support lug or the exemplary dimensions of the sealing plate or the separation distance between the pressure wall side 102 and the suction wall side 104. As Figure 4 As shown in

[0133] It can be noted that, although all the figures only show one connector flow passage 130, the vane 1 can comprise (not shown) a plurality of connector flow passages 130, preferably axially extending, spaced apart from each other in the circumferential direction of the inner shroud 100. Thus, the cooling air 5 is received at the second region R2 or the cooling air receiving chamber 124c in a distributed manner, whereby a more uniform impingement jet is achieved by the second impingement plate 122.

[0134] With reference to Figure 8 Further embodiments of the vane 1 of the present technology are explained.

[0135] As Figure 8As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed.

[0136] A diameter of the second impingement hole 122h of the second impingement plate 122 can be less than a diameter of the first impingement hole 112h of the first impingement plate 112.

[0137] As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed. Figure 9A and Figure 9B As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed.

[0138] The cooling air from the shroud cooling hole 100h can be injected into the hot gas path 55. The outlet 100hb of the shroud cooling hole 100h can be positioned at the hot gas path 55.

[0139] As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed. Figure 9A As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed.

[0140] Figure 9A As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed.

[0141] As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed. Figure 9A As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed.

[0142] Figure 9B As shown in the example of FIG. 1, the radial distance H2 of the second impingement plate 122 from the lower surface 100b of the inner shroud 100 can be less than or equal to the radial distance H1 of the first impingement plate 112 from the lower surface 100b of the inner shroud 100. Accordingly, an impingement jet having an increased force for impinging onto the lower surface 100b of the inner shroud 100 in the second region R2 can be formed. ​​

[0143] The shroud cooling hole 100h can be an inclined through hole with respect to the lower surface 100b and the side surface 100s of the inner shroud 100, i.e. the shroud cooling hole 100h can be inclined with respect to the radial direction.

[0144] With reference to Figure 10 , the vane 1 can be formed such that the lower surface 100b of the inner shroud 100 can comprise the base opening 61 of the airfoil 60 in the first region R1. The base opening 61 can be understood as the opening of the inner space or lumen of the airfoil 60. The lumen of the airfoil 60 is the cavity or space enclosed by the airfoil shape, i.e. the space or cavity defined by the pressure wall 62, the suction wall 64, the leading edge 66 and the trailing edge 68 of the airfoil 60.

[0145] The first impingement chamber 110c and the base opening 61 of the airfoil 60 can be non-overlapping or non-continuous with respect to each other. In other words, the first impingement chamber 110c and the base opening 61 of the airfoil 60 are fluidically disconnected from each other, i.e. cooling air 5 introduced into the first chamber 110c does not flow into the base opening 61 of the airfoil 60.

[0146] The above described gas turbine vane 1 can be incorporated in a gas turbine 10, e.g. in a gas turbine 10 of Figure 1 .

[0147] The vane 1 can be an inlet guide vane 44 or a vane 1 of a first vane stage of the turbine section 18. The vane 1 can be located directly downstream of the transition duct 17.

[0148] The first impingement plate 112 can be configured to receive cooling air 5 from the last stage or final stage of the compressor section 14.

[0149] Figures 11A-11C Further exemplary embodiments of the present technology are depicted. As can be seen in Figure 11A , Figure 11B and Figure 11C , the first impingement plate 112 and the second impingement plate 122 are arranged to face the lower surface 100b of the inner shroud 100 in the first region R1 and in the second region R2, respectively. The cover plate 124 of the second impingement unit is also arranged to face the second impingement plate 122 in the second region R2.

[0150] As shown in Figure 11A , the first impingement plate 112 and the cover plate 124 can be flush with each other.

[0151] In other words, the distance between the upper surface 100a of the inner shroud 100 and the first impingement plate 112 can be the same as the distance between the upper surface 100a of the inner shroud 100 and the cover plate 124.

[0152] The second impact plate 122 can be arranged between the lower surface 100b of the inner shroud 100 and the cover plate 124.

[0153] The distance between the lower surface 100b of the inner shroud 100 and the cover plate 124 in the second region R2 can be greater than the distance between the lower surface 100b of the inner shroud 100 and the first impact plate 112 in the first region R1.

[0154] As shown in Figure 11B The first impact plate 112 and the cover plate 124 can not be flush with each other, as shown in

[0155] In other words, the distance between the first impact plate 112 in the first region R1 and the upper surface 100a and / or the lower surface 100b of the inner shroud 100 can be the same as the distance between the second impact plate 122 in the second region R2 and the upper surface 100a and / or the lower surface 100b of the inner shroud 100.

[0156] The second impact plate 122 can be arranged between the lower surface 100b of the inner shroud 100 and the cover plate 124.

[0157] The connector flow passage 130 can be formed obliquely, i.e. can be formed so as to be inclined from the inlet 132 towards the rotational axis of the gas turbine towards the outlet 134. In other words, can be formed so as to be inclined from a radially outward position in the first region R1 towards a radially inward position in the second region R2. In other words, the connector flow passage 130 can be formed so as to be inclined such that the inlet 132 is arranged at a radially outward position in the first region R1 and the outlet 134 is arranged at a radially inward position in the second region R2.

[0158] As shown in Figure 11C The first impact plate 112 can be formed in a stepped manner, i.e. a first portion of the first impact plate 112 arranged adjacent to the inlet 132 of the connector flow passage 130 can be formed so as to be arranged at a radially inward position relative to a second portion of the first impact plate 112 arranged distal to the inlet 132 of the connector flow passage 130, as shown in

[0159] The first portion of the first impact plate 112 can be flush with the cover plate 124.

[0160] The second portion of the first impact plate 112 can be flush with the second impact plate 122.

[0161] List of reference signs

[0162] 1 turbine blade

[0163] 5 cooling air

[0164] 5a first portion of cooling air

[0165] 5b second portion of cooling air

[0166] 9a first flow passage

[0167] 9b second flow passage

[0168] 10 gas turbine

[0169] 12 inlet

[0170] 14 compressor section

[0171] 16 combustion chamber section

[0172] 17 transition duct

[0173] 18 turbine section

[0174] 19 combustion chamber pot

[0175] 20 longitudinal or rotational axis

[0176] 22 shaft

[0177] 24 air

[0178] 26 combustor plenum

[0179] 28 combustion chamber

[0180] 30 combustor

[0181] 32 diffuser

[0182] 34 combustion or working gas

[0183] 36 blade carrier disk

[0184] 36c cooling air passage

[0185] 38 turbine blade

[0186] 40 guide vane

[0187] 42 stator

[0188] 44 inlet guide vane

[0189] 46 vane stage

[0190] 48 rotor blade stage

[0191] 50 casing

[0192] 52 radially outer surface

[0193] 53 rotor drum

[0194] 54 radially inner surface

[0195] 55 hot gas path

[0196] 56 passage

[0197] 60 airfoil

[0198] 61 base opening of airfoil

[0199] 62 pressure wall

[0200] 64 suction wall

[0201] 66 leading edge

[0202] 68 trailing edge

[0203] 70 seal unit

[0204] 72 seal support lug

[0205] 74 seal plate

[0206] 80 interstage seal

[0207] 82 space

[0208] 90 outer shroud

[0209] 92 blade fixation means

[0210] 100 inner shroud

[0211] 100a upper surface

[0212] 100b lower surface

[0213] 100h shroud cooling hole

[0214] 100ha inlet of shroud cooling hole

[0215] 100hb outlet of shroud cooling hole

[0216] 100s side surface of inner shroud

[0217] 101 intervening section of inner shroud

[0218] 102 pressure wall side

[0219] 104 suction wall side

[0220] 106 leading edge side

[0221] 108 trailing edge side

[0222] 110 first impact unit

[0223] 110c first impact chamber

[0224] 112 first impact plate

[0225] 112h first impact cooling hole

[0226] 112s side plate

[0227] 120 second impact unit

[0228] 120c second impact chamber

[0229] 122 second impact plate

[0230] 122h second impact cooling hole

[0231] 122s side plate

[0232] 124 cover plate

[0233] 124c cooling air receiving chamber

[0234] 130 connector flow passage

[0235] 132 inlet

[0236] 134 outlet

[0237] C1 first impact cavity

[0238] C11 opening of first impact cavity

[0239] C2 second impact cavity

[0240] C21 opening of second impact cavity

[0241] H2 distance of second impact plate from inner shroud

[0242] H1 distance of first impact plate from inner shroud

[0243] R1 first region

[0244] R2 second region

[0245] W1 width of connector flow passage

[0246] W2 width of sealing support lug or width of sealing plate

Claims

1. A turbine blade for a gas turbine, the turbine blade comprising: – An inner protective cover having an upper surface and a lower surface; – A sealing unit disposed on the lower surface of the inner cover and defining a first region and a second region on the lower surface of the inner cover, wherein the inner cover includes a first impact cavity located in the first region on the lower surface of the inner cover and a second impact cavity located in the second region on the lower surface of the inner cover. – A first impact unit, the first impact unit being disposed in the first region and including a first impact plate facing the inner shroud, defining a first impact chamber between the inner shroud and the first impact plate, wherein the first impact chamber includes the first impact cavity, and wherein the first impact plate is configured to receive cooling air and form an impact jet directed into the first impact chamber. – A second impact unit, disposed in the second region and including a second impact plate facing the inner shield, defining a second impact chamber between the inner shield and the second impact plate, wherein the second impact chamber includes the second impact cavity; and – At least one connector flow channel configured to guide cooling air from the first impact chamber to the second region; The second impact plate is configured to receive cooling air from the connector flow channel and form an impact jet guided into the second impact chamber. Its features are: The first impact chamber and the second impact chamber are separated by the intervention section of the inner shield, and the connector flow channel extends through the intervention section of the inner shield.

2. The turbine blade according to claim 1, wherein, The second impact plate is arranged flush with the opening of the second impact cavity, or the second impact plate is arranged inside the second impact cavity.

3. The turbine blade according to claim 1, wherein, The connector flow channel extends through the sealing unit.

4. The turbine blade according to claim 1, wherein, The sealing unit includes at least one of the following: – A sealing support lug extending radially inward from the lower surface of the inner shroud; and – A sealing plate, which is supported at the inner cover and arranged radially inward from the inner cover; The connector flow channel extends through at least one of the sealing support lug and the sealing plate.

5. The turbine blade according to claim 4, wherein, The width of the connector flow channel is between 2% and 40% of the width of the sealing support lug or the sealing plate measured along the circumferential direction of the inner cover.

6. The turbine blade according to claim 5, wherein, The width of the connector flow channel is between 5% and 15% of the width of the sealing support lug or the sealing plate measured along the circumferential direction of the inner cover.

7. The turbine blade according to claim 1, wherein, The second impact unit includes a cover plate arranged radially inside and facing the second impact plate and defining a cooling air receiving chamber between the cover plate and the second impact plate, wherein the outlet of the connector flow channel is located at the cooling air receiving chamber.

8. The turbine blade according to claim 1, wherein, The lower surface of the inner shield includes a base opening of the airfoil of the turbine blade in the first region, wherein the first impact chamber and the base opening of the airfoil do not overlap.

9. The turbine blade according to claim 1, in, The radial distance between the second impact plate and the lower surface of the inner shield is less than or equal to the radial distance between the first impact plate and the lower surface of the inner shield; and / or The diameter of the second impact hole in the second impact plate is smaller than the diameter of the first impact hole in the first impact plate.

10. The turbine blade according to claim 1, wherein, The inner shield includes at least one shield cooling hole, the at least one shield cooling hole having an inlet located at the second impact chamber and an outlet located at the upper surface or the side surface of the inner shield.

11. A gas turbine including turbine blades, wherein, The turbine blades are as described in claim 1.

12. The gas turbine according to claim 11, wherein, The second impact plate is arranged flush with the opening of the second impact cavity, or the second impact plate is arranged inside the second impact cavity.

13. The gas turbine of claim 11, comprising: – Turbine blades, the turbine blades being located downstream of the turbine wheel blades and mounted on a blade support disk; as well as – An interstage seal, the interstage seal being axially disposed between the blade bearing disk and the turbine blade; wherein the second impact unit is positioned in a space defined by the inner shroud of the turbine blade, the sealing unit and the interstage seal, and wherein the interstage seal is configured to seal the space radially inward.

14. The gas turbine of claim 11, further comprising a compressor, wherein, The first impact plate is configured to receive cooling air from the last stage of the compressor.

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