Blade ring assembly and gas turbine

Abstract

A blade ring assembly includes a turbine blade ring extending in a circumferential direction around an axis, a cooling target member disposed on an inner peripheral side of the turbine blade ring, and an outer peripheral side member disposed on an outer peripheral side of the turbine blade ring. The turbine blade ring has a cooling medium inlet port that opens from an outer peripheral surface of the turbine blade ring to an inner peripheral surface. The outer peripheral side member has a first wall portion that covers at least a portion of the cooling medium inlet port from the outer peripheral side of the turbine blade ring and extends to a downstream side of the axis in both sides of the axis further than the cooling medium inlet port, and a second wall portion that extends from an end portion of the downstream side of the axis of the first wall portion toward the outer peripheral surface of the turbine blade ring and covers at least a portion of a space between the first wall portion and the cooling medium inlet port from the downstream side of the axis.

Description

Technical Field

[0001] This invention relates to a blade ring assembly, a gas turbine, and a method for modifying a gas turbine.

[0002] This application claims priority based on Japanese Patent Application No. 2021-070571, filed on April 19, 2021, the contents of which are incorporated herein by reference. Background Technology

[0003] Patent Document 1 discloses a gas turbine that includes a mechanism for separating particles contained in cooling air. The gas turbine has a particle separation protection element disposed on the inner circumferential side (rotor side) of the stationary blades, which prevents floating particles from flowing into the intake opening.

[0004] Previous technical documents

[0005] Patent documents

[0006] Patent Document 1: Japanese Patent Publication No. 2010-501764 Summary of the Invention

[0007] The technical problem to be solved by the invention

[0008] However, the protective element for particle separation described in Patent Document 1 does not prevent cooling air supplied from the compressor to the housing from flowing directly into the air passage leading to the interior of the stationary blades. Therefore, cooling air containing foreign matter may flow into the interior of the stationary blades, causing blockage in the cooling path inside the stationary blades.

[0009] The present invention was made to solve the above-mentioned problems, and its purpose is to provide a blade ring assembly, a gas turbine, and a method for modifying a gas turbine that can suppress the inflow of foreign matter into components such as stationary blades that are being cooled.

[0010] means for solving technical problems

[0011] To address the aforementioned issues, the blade ring assembly of the present invention comprises: a turbine blade ring extending circumferentially about an axis; a cooling target component disposed on the inner circumferential side of the turbine blade ring; and an outer circumferential side component disposed on the outer circumferential side of the turbine blade ring. The turbine blade ring has a cooling medium inlet extending from the outer circumferential surface of the turbine blade ring to the inner circumferential surface. The outer circumferential side component comprises: a first wall portion covering at least a portion of the cooling medium inlet from the outer circumferential side of the turbine blade ring and extending to the downstream side of the central axis on both sides in an axial direction extending further along the axis than the cooling medium inlet; and a second wall portion extending from the downstream end of the first wall portion toward the outer circumferential surface of the turbine blade ring, covering at least a portion of the space between the first wall portion and the cooling medium inlet from the downstream side of the axis.

[0012] To address the aforementioned issues, the gas turbine of the present invention comprises: a blade ring assembly; a rotor rotatable about an axis; a housing covering the outer periphery of the rotor; and a canister burner that generates combustion gas through fuel combustion and delivers the combustion gas into the housing. The blade ring assembly comprises: a turbine blade ring extending circumferentially about the axis; a cooling target component disposed on the inner periphery of the turbine blade ring; and an outer periphery component disposed on the outer periphery of the turbine blade ring. The turbine blade ring has a cooling medium inlet extending from the outer periphery of the turbine blade ring to the inner periphery. The outer periphery component comprises: a first wall portion covering at least a portion of the cooling medium inlet from the outer periphery of the turbine blade ring and extending downstream of the axis on both sides in an axial direction extending further along the axis than the cooling medium inlet; and a second wall portion extending from the downstream end of the first wall portion toward the outer periphery of the turbine blade ring, covering at least a portion of the space between the first wall portion and the cooling medium inlet from the downstream side of the axis. The blade ring assembly is disposed on the inner circumferential side of the housing. The housing has a partition that, as part of the wall of a receiving chamber defining the exposure of the blade ring assembly and the flow of the cooling medium, is disposed downstream of the axis and extends radially along the turbine blade ring.

[0013] To address the aforementioned issues, the present invention relates to a method for modifying a gas turbine, wherein the gas turbine comprises: a turbine blade ring extending circumferentially about an axis; and a stationary blade disposed on the inner circumferential side of the turbine blade ring. The method for modifying the gas turbine includes the following steps: removing the stationary blade from the turbine blade ring; installing a shield on the stationary blade, wherein the stationary blade has a blade body disposed within a combustion gas flow path and having a blade shape, and an inner shield disposed at the end of the inner circumferential side of the blade body and having a space, wherein by installing the shield on the inner shield, the shield covers at least a portion of the space of the inner shield from the inner circumferential side; and installing the stationary blade with the shield installed on it onto the turbine blade ring with a foreign matter separation cover installed thereon, wherein the turbine blade ring has a cooling medium inlet extending from the outer circumferential surface of the turbine blade ring to the inner circumferential surface, and the foreign matter separation cover is disposed on the outer circumferential side of the turbine blade ring, covering at least a portion of the cooling medium inlet.

[0014] Invention Effects

[0015] The blade ring assembly, gas turbine, and gas turbine modification method of the present invention can suppress the inflow of foreign matter into components such as stationary blades that are subject to cooling. Attached Figure Description

[0016] Figure 1 This is a schematic cross-sectional view of the gas turbine according to the first embodiment of the present invention.

[0017] Figure 2 This is an enlarged cross-sectional view showing a portion of the gas turbine according to the first embodiment of the present invention.

[0018] Figure 3 This is a perspective view schematically showing a portion of the blade ring assembly according to the first embodiment of the present invention.

[0019] Figure 4 This is an exploded perspective view showing a portion of the blade ring assembly according to the first embodiment of the present invention.

[0020] Figure 5 yes Figure 3 The cross-sectional view of the blade ring assembly shown is taken along line F5-F5.

[0021] Figure 6 This is a top view showing the dust separator according to the first embodiment of the present invention.

[0022] Figure 7 This is a front view showing the dust separator according to the first embodiment of the present invention.

[0023] Figure 8 This is a cross-sectional view showing the shielding cover according to the first embodiment of the present invention.

[0024] Figure 9 This is a flowchart illustrating the steps of the method for modifying a gas turbine according to the first embodiment of the present invention.

[0025] Figure 10 This is a cross-sectional view illustrating the function of the dust separator according to the first embodiment of the present invention.

[0026] Figure 11 This is a perspective view schematically showing a portion of the blade ring assembly according to the second embodiment of the present invention.

[0027] Figure 12 This is an enlarged and schematic cross-sectional view showing a portion of the blade ring assembly according to the second embodiment of the present invention.

[0028] Figure 13 This is an exploded perspective view of the L-shaped tube according to the second embodiment of the present invention.

[0029] Figure 14 This is a cross-sectional view illustrating the function of the L-shaped tube according to the second embodiment of the present invention. Detailed Implementation

[0030] Hereinafter, with reference to the accompanying drawings, the blade ring assembly, gas turbine, and method for modifying the gas turbine according to the first embodiment of the present invention will be described. In the following description, structures having the same or similar functions will be labeled with the same reference numerals. Furthermore, repeated descriptions of these structures will sometimes be omitted.

[0031] <First Embodiment>

[0032] (Structure of a gas turbine)

[0033] Figure 1 This is a schematic cross-sectional view showing the gas turbine 10 of the first embodiment. The gas turbine 10 includes: a compressor 20 that compresses air A; a combustor 30 that burns fuel F in the air A compressed by the compressor 20 to generate combustion gas G; and a turbine 40 that is driven by the combustion gas G.

[0034] The compressor 20 has a compressor rotor 21 that rotates around an axis Ar, a compressor housing 25 covering the outer periphery of the compressor rotor 21, and multiple layers of stationary blades 26. The turbine 40 has a turbine rotor 41 that rotates around an axis Ar, a turbine housing 45 covering the outer periphery of the turbine rotor 41, and multiple layers of stationary blades 46.

[0035] The compressor rotor 21 and the turbine rotor 41 are located on the same axis Ar and are connected to each other to form the gas turbine rotor 11. For example, a generator rotor GEN is connected to the gas turbine rotor 11. The compressor housing 25 and the turbine housing 45 are connected to each other to form the gas turbine housing 15. The gas turbine housing 15 is an example of a "casing". The burner 30 is, for example, a canister burner.

[0036] In the following description, the direction in which the axis Ar extends is defined as the axial direction Da, the circumferential direction centered on the axis Ar is defined as the circumferential direction Dc, and the direction perpendicular to the axis Ar is defined as the radial direction Dr. On both sides of the axial direction Da, with the turbine 40 as a reference, the compressor 20 side is defined as the upstream side Dau, and its opposite side is defined as the downstream side Dad. Furthermore, on both sides of the radial direction Dr, the side closer to the axis Ar is defined as the radially inner side Dri, and its opposite side is defined as the radially outer side Dro. Furthermore, on both sides of the circumferential direction Dc, the direction from the positive pressure surface of the blade body 110 of the stationary blade 46a (described later) towards the negative pressure surface is defined as the circumferential direction Dcn, and the direction from the negative pressure surface of the blade body 110 towards the positive pressure surface is defined as the circumferential direction Dcp. Hereinafter, the upstream side Dau of the axial direction Da is sometimes referred to as the front side, and the downstream side Dad of the axial direction Da is sometimes referred to as the rear side. The axial direction Da is the flow direction of the combustion gas G.

[0037] The compressor rotor 21 has a rotor shaft 22 extending along the axial direction Da with axis Ar as the center, and multiple moving blade layers 23 mounted on the rotor shaft 22. The multiple moving blade layers 23 are arranged along the axial direction Da. Each moving blade layer 23 is composed of multiple moving blades 23a arranged circumferentially Dc. A stationary blade layer 26 is disposed on the downstream side Dad of each moving blade layer 23. Each stationary blade layer 26 is disposed inside the compressor housing 25. Each stationary blade layer 26 is composed of multiple stationary blades 26a arranged circumferentially Dc.

[0038] The turbine rotor 41 has a rotor shaft 42 extending along the axial direction Da with axis Ar as its center, and multiple moving blade layers 43 mounted on the rotor shaft 42. The multiple moving blade layers 43 are arranged along the axial direction Da. Each moving blade layer 43 is composed of multiple moving blades 43a arranged circumferentially Dc. A stationary blade layer 46 is disposed upstream of each moving blade layer 43 on its axial direction Dau. Each stationary blade layer 46 is located inside the turbine housing 45. Each stationary blade layer 46 is composed of multiple gas turbine stationary blades 46a arranged circumferentially Dc. In the following description, the gas turbine stationary blades are simply referred to as stationary blades. The gas turbine stationary blades (stationary blades) 46a are an example of a "cooled component".

[0039] Figure 2This is an enlarged cross-sectional view showing a portion of the gas turbine 10 according to the first embodiment. Furthermore, for ease of explanation, Figure 2 schematic representation Figure 3 The cross-section of the blade ring assembly WS shown is taken along line F2-F2. The turbine housing 45 has a cylindrical outer housing 45a constituting the outer shell of the turbine housing 45, an inner housing 45b fixed to the inner side of the outer housing 45a, and a plurality of segmented rings 45c fixed to the inner side of the inner housing 45b. The plurality of segmented rings 45c are all positioned between two adjacent stationary blade layers 46 in the plurality of stationary blade layers 46. A moving blade layer 43 is disposed radially inward (Dri) on the inner side of each segmented ring 45c.

[0040] The space between the rotor shaft 42 on the radial direction Dr and the turbine housing 45, where stationary blades 46a and moving blades 43a are arranged, constitutes a combustion gas flow path 49 for the combustion gas G from the combustor 30 to flow. The combustion gas flow path 49 is annular with the axis Ar as its center and is longer in the axial direction Da.

[0041] The gas turbine 10 of this embodiment includes a cooling device 50 that supplies cooling air to the stationary blades 46a and the dividing ring 45c in the axial direction Da from the second layer onwards. The cooling device 50 includes a foreign object trap 51, a cooler 52, a booster compressor 53, and a cooling air line 54. The foreign object trap 51, for example, is a filter with multiple fine pores that separates foreign objects contained in the cooling air flowing through the cooling air line 54. The cooler 52 cools the cooling air flowing through the cooling air line 54. The booster compressor 53 pressurizes the cooling air flowing through the cooling air line 54. The cooling air line 54 draws compressed air Ac from the gas turbine housing 15 as cooling air and supplies the drawn cooling air to the inside of the outer housing 45a via the foreign object trap 51, the cooler 52, and the booster compressor 53.

[0042] A cooling air passage 45p is formed in the inner chamber 45b of the turbine housing 45, extending radially outward (Dro) to radially inward (Dri). Cooling air supplied from the cooling air duct 54 to the inner side of the outer chamber 45a is introduced through the cooling air passage 45p of the inner chamber 45b into the stationary blades 46a and the dividing ring 45c beyond the second layer, for cooling these stationary blades 46a and the dividing ring 45c. Here, since a foreign object trap 51 is provided in the cooling air duct 54, foreign objects are less likely to reach the stationary blades 46a and the dividing ring 45c beyond the second layer. In addition, the path for supplying cooling air to the stationary blades 46a is not limited to the above.

[0043] (The operation of a gas turbine)

[0044] Return to Figure 1The operation of the gas turbine 10 is explained below. Compressor 20 compresses air A to generate compressed air Ac. The compressed air Ac generated by compressor 20 flows into combustor 30. Fuel F is supplied to combustor 30. Inside combustor 30, fuel F burns in the compressed air to generate high-temperature, high-pressure combustion gas G. Combustion gas G generated by combustor 30 is transported from combustor 30 to combustion gas flow path 49 within turbine 40. As combustion gas G flows downstream along the axis in combustion gas flow path 49, it rotates turbine rotor 41. The rotation of turbine rotor 41 causes the rotor of generator GEN, connected to gas turbine rotor 11, to rotate. As a result, generator GEN generates electricity.

[0045] (Blade ring assembly)

[0046] like Figure 2 As shown, the gas turbine 10 has a blade ring assembly WS. The blade ring assembly WS is disposed on the inner circumferential side of the gas turbine housing 15. In this embodiment, the blade ring assembly WS is disposed on the upstream side of the turbine 40 in the axial direction Da, and has a turbine blade ring 70 and a plurality of stationary blades 46a1 that respectively constitute the first layer of stationary blades 46a in the axial direction Da. In this embodiment, the compressed air Ac supplied from the compressor 20 to the gas turbine housing 15 is directly supplied as cooling air to the first layer of stationary blades 46a1 without passing through a foreign matter trap. Therefore, in this embodiment, a dust separator 90 is provided on the turbine blade ring 70, which prevents foreign matter contained in the compressed air Ac from reaching the stationary blades 46a1. Hereinafter, this will be described in detail. Hereinafter, for ease of explanation, the compressed air Ac will sometimes be referred to as cooling air Ac.

[0047] (Structure surrounding the blade ring assembly)

[0048] The gas turbine housing 15 has a front wall 61, a peripheral wall 62, and a rear wall 63 as walls defining a housing C for accommodating the combustor 30. The housing C is the space within the gas turbine housing 15 where the blade ring assembly WS is exposed and compressed air Ac flows.

[0049] The front wall 61 is located on the upstream side of the axis relative to the receiving chamber C. The front wall 61 has a cylindrical portion 61a with an opening 61h and a cylindrical cover portion 61b mounted on the cylindrical portion 61a and covering the opening 61h. A portion of the burner 30 is disposed inside the cylindrical portion 61a and the cylindrical cover portion 61b. For example, the intake portion 31 of the burner 30 is disposed inside the cylindrical cover portion 61b.

[0050] The peripheral wall 62 is located radially outward (Dro) relative to the receiving chamber C. The peripheral wall 62 extends between and connects the front wall 61 and the rear wall 63. The peripheral wall 62 has a first portion 62a and a second portion 62b. The first portion 62a is the portion connected to the front wall 61. The first portion 62a extends along the axial direction Da. The second portion 62b is located further downstream of the axial direction (Dad) than the first portion 62a and is the portion connected to the rear wall 63. The second portion 62b is, for example, an inclined portion (reduced diameter portion) that slopes radially inward (Dai) as it moves downstream of the axial direction (Dad). The second portion 62b includes, for example, an arcuate portion whose slope steepens relative to the axial direction (Dad) as it moves downstream of the axial direction (Dad).

[0051] The rear wall 63 is located downstream of the housing C on the Dad side of the axis. The rear wall 63 extends radially Dr. The rear wall 63 is a partition that closes the downstream Dad side of the housing C on the axis. The blade ring fixing part 71 of the turbine blade ring 70, described later, is fixed to the rear wall 63. The rear wall 63 is located further downstream of the axis on the Dad side than the air intake 72 of the turbine blade ring 70, described later.

[0052] like Figure 2 As shown, when viewed from the compressor 20, a guide section 64 for compressed air Ac is provided at the inlet of the housing C. The guide section 64 is, for example, a guide vane inclined relative to the axial direction Da, causing the flow direction of the compressed air Ac flowing from the compressor 20 to change toward the peripheral wall 62 of the gas turbine housing 15. However, if a portion of the compressed air Ac flowing from the compressor 20 is directed toward the peripheral wall 62 or the rear wall 63, the guide section 64 can be omitted.

[0053] (Structure of the blade ring assembly)

[0054] Figure 3 This is a perspective view showing a portion of the blade ring assembly WS. The blade ring assembly WS includes a turbine blade ring 70, multiple combustor connection components 80, multiple sealing components 85, and multiple stationary blades 46a1. Figure 5 Only one is shown in the image), multiple dust separators 90 and multiple shields 150 (only one is shown in the image). Figure 5 Only one is shown in the image.

[0055] (Turbine blade ring)

[0056] The turbine blade ring 70 extends circumferentially Dc around the axis Ar, forming a ring shape. The turbine blade ring 70 has an outer circumferential surface 70o, an inner circumferential surface 70i, an upstream end face 70u, and a downstream end face 70d. The outer circumferential surface 70o faces radially outward (Dro). The outer circumferential surface 70o is exposed in the housing C of the gas turbine housing 15. The inner circumferential surface 70i is located on the opposite side of the outer circumferential surface 70o, facing radially inward (Dri). The inner circumferential surface 70i faces multiple stationary blades 46a1. The upstream end face 70u faces upstream (Dau) of the axis. The downstream end face 70d faces downstream (Dad) of the axis.

[0057] The turbine blade ring 70 has a blade ring fixing portion 71 fixed to the gas turbine housing 15. The blade ring fixing portion 71 is located at the downstream end of the turbine blade ring 70's axis, on the Dad side. The blade ring fixing portion 71 is, for example, a flange protruding radially outward on the Dro side. The blade ring fixing portion 71 is located radially inward on the Dri side of the rear wall 63 of the gas turbine housing 15 (see reference). Figure 2 The blade ring fixing part 71 is fixedly supported on the rear wall 63 of the gas turbine housing 15.

[0058] like Figure 3 As shown, a plurality of air intakes 72 are provided on the turbine blade ring 70. The plurality of air intakes 72 are spaced apart at predetermined intervals on the circumference of the turbine blade ring 70. The air intakes 72 extend from the outer circumferential surface 70o of the turbine blade ring 70 to the inner circumferential surface 70i. The air intakes 72 are located upstream of the blade ring fixing portion 71 on the axial direction, at a distance Dau.

[0059] Figure 4 This is an exploded perspective view showing a portion of the blade ring assembly WS. In this embodiment, the air intake 72 is a cutout (groove) provided at the end of the turbine blade ring 70 on the upstream side of the axial direction, Dau. That is, the air intake 72 extends from the outer peripheral surface 70o of the turbine blade ring 70 to the inner peripheral surface 70i and opens towards the upstream side of the axial direction, Dau. Alternatively, the air intake 72 may also be a through hole in the central portion of the turbine blade ring 70 in the axial direction Da, extending from the outer peripheral surface 70o to the inner peripheral surface 70i. The air intake 72 guides a portion of the cooling air Ac flowing through the receiving chamber C of the gas turbine housing 15 from the outer peripheral side to the inner peripheral side of the turbine blade ring 70. The air intake 72 is an example of a "cooling medium intake". In this embodiment, the compressed air Ac compressed by the compressor 20 is an example of a "cooling medium".

[0060] In this embodiment, the upstream end of the turbine blade ring 70 on the axial direction has a plurality of protrusions 73. The plurality of protrusions 73 are disposed at different positions on the circumferential direction Dc and protrude toward the upstream end of the axial direction Dau, respectively. The end faces of the upstream end of the plurality of protrusions 73 on the axial direction Dau form the upstream end face 70u of the turbine blade ring 70. The plurality of protrusions 73 and the plurality of air intakes 72 are alternately arranged on the circumferential direction Dc (see reference). Figure 3 Therefore, considering a single air intake 72, the plurality of protrusions 73 have a first protrusion 73a located on the circumferential Dcn side of the air intake 72 and a second protrusion 73b located on the circumferential Dcp side of the air intake 72. In other words, the air intake 72 is located between the first protrusion 73a and the second protrusion 73b in the circumferential Dc direction.

[0061] (Burner connection components)

[0062] Next, the burner connection component 80 will be described. The burner connection component 80 is a fixing component that fixes the tailstock 32 of the burner 30. Multiple burner connection components 80 are provided on the circumferential direction Dc at positions corresponding to multiple protrusions 73 of the turbine blade ring 70.

[0063] like Figure 4 As shown, the burner connecting component 80 has a frame portion 81 and a flange 82. The frame portion 81 is the part that connects to the tail tube 32 of the burner 30. The tail tube 32 of the burner 30 is fixed to the frame portion 81, for example, by welding. The flange 82 is the part that extends radially outward from the frame portion 81. The flange 82 has a planar shape along the circumferential direction Dc and the radial direction Dr. The flange 82 faces the end face 73u of the upstream side of the protrusion 73 in the axial direction Da. A connecting hole 73h is provided on the end face 73u of the upstream side of the protrusion 73, which can be engaged by a connecting tool 83 such as a bolt.

[0064] A through hole 82h is provided on the flange 82 for the connecting tool 83 to pass through. The flange 82 is fixed to the protrusion 73 by engaging the connecting tool 83, which passes through the through hole 82h of the flange 82, with the connecting hole 73h of the protrusion 73. The end face 73u of the protrusion 73, which includes the connecting hole 73h, is an example of a "connecting portion" capable of being connected to the burner connecting member 80. In this embodiment, the end faces 73u of the first protrusion 73a and the second protrusion 73b respectively have connecting holes 73h and are fitted with burner connecting members 80.

[0065] (Sealing components)

[0066] Next, the sealing component 85 will be described. The sealing component 85 is disposed between the tailstock 32 of two adjacent burners 30 in the circumferential direction Dc, and hermetically seals the gap between the tailstock 32 of the two burners 30. The sealing component 85 is disposed on the inner circumferential side of the turbine blade ring 70 (see reference). Figure 5 The sealing component 85 is positioned on the radial Dr at a location overlapping the air intake 72 (see reference). Figure 5 and Figure 6 With the burner 30's tailstock 32 detached from the burner connection assembly 80, the sealing component 85 can be removed from the outer periphery of the turbine blade ring 70 via the air intake 72. The sealing component 85 is an example of a "removable component".

[0067] (Stationary blade)

[0068] Next, the stationary blade 46a1 will be described.

[0069] Figure 5 yes Figure 3 The cross-sectional view of the blade ring assembly WS shown is taken along line F5-F5. Multiple stationary blades 46a1 ( Figure 5 (Only one is shown in the image) It is disposed on the inner circumferential side of the turbine blade ring 70 and arranged circumferentially Dc. The stationary blade 46a1 is held by a retaining member 74 provided on the turbine blade ring 70. The stationary blade 46a1 has a blade body 110, an outer shroud 120, an inner shroud 130 and an air passage 140.

[0070] The blade body 110 is blade-shaped and extends radially along Dr. That is, the blade height direction of the blade body 110 is radially Dr. The blade body 110 is disposed within the combustion gas flow path 49 through which the combustion gas G passes. On the surface of the blade body 110, in the face facing the circumferential direction Dc, the convex surface forms the back surface (= negative pressure surface), and the concave surface forms the ventral surface (= positive pressure surface). Multiple exhaust holes 110h are provided at the end of the blade body 110 upstream of the central axis (Dau) and at the end of the blade body 110 downstream of the central axis (Dad).

[0071] The outer shroud 120 is disposed at the radially outer end of the blade body 110 (Dro), defining the outer peripheral position of the annular combustion gas flow path 49. The outer shroud 120 has a shroud body 121, a peripheral wall 122, and a collision plate 123.

[0072] The shield body 121 is formed as a plate extending along the axial direction Da and the circumferential direction Dc. The shield body 121 has a gas path surface 121a and an outer inner surface 121b. The gas path surface 121a is the surface in contact with the combustion gas G (the surface facing the combustion gas flow path 49), facing radially inward Dri. The outer inner surface 121b is the surface facing the side opposite to the gas path surface 121a.

[0073] The peripheral wall 122 protrudes radially outward from the outer periphery of the protective cover body 121. In this embodiment, the peripheral wall 122 is formed over the entire circumference of the outer periphery of the protective cover body 121. The peripheral wall 122 has a front wall facing upstream of the axis (Dau), a rear wall facing downstream of the axis (Dad), a dorsal side wall facing circumferentially (Dcn), and a ventral side wall facing circumferentially (Dcp) (not shown). The outer protective cover 120 has a space, namely the first space S1, surrounded by the peripheral wall 122 in four directions.

[0074] A collision plate 123 is disposed within the first space S1 of the outer shield 120, dividing the first space S1 into a radially outer region Dro and a radially inner region Dri, i.e., the cavity CA. Multiple air holes 123h are formed on the collision plate 123, extending radially through Dr. A portion of the cooling air Ac present on the radially outer Dro of the stationary blade 46a1 flows into the cavity CA through the air holes 123h of the collision plate 123. After cooling the outer shield 120, a portion of the air flowing into the cavity CA is discharged into the combustion gas flow path 49 from an exhaust port (not shown) disposed on the outer shield 120.

[0075] The inner shield 130 is disposed at the radially inner end of the blade body 110 (Dri), defining the inner circumferential position of the annular combustion gas flow path 49. The inner shield 130 has a shield body 131, a peripheral wall 132, and a collision plate 133.

[0076] The shield body 131 is formed as a plate extending along the axial direction Da and the circumferential direction Dc. The shield body 131 has a gas path surface 131a and an inner surface 131b. The gas path surface 131a is the surface in contact with the combustion gas G (the surface facing the combustion gas flow path 49), facing radially outward Dro. The inner surface 131b is the surface facing the side opposite to the gas path surface 131a. The shield body 131 has an exhaust port 131h, which connects the second space S2 (described later) to the combustion gas flow path 49.

[0077] The peripheral wall 132 protrudes radially inward from the outer periphery of the protective cover body 131. In this embodiment, the peripheral wall 132 is formed over the entire circumference of the outer periphery of the protective cover body 131. The peripheral wall 132 has a front wall facing upstream of the axis (Dau), a rear wall facing downstream of the axis (Dad), a dorsal side wall facing circumferentially (Dcn), and a ventral side wall facing circumferentially (Dcp) (not shown). The inner protective cover 130 has a space, namely the second space S2, surrounded by the peripheral wall 132 in four directions.

[0078] The collision plate 133 is disposed within the second space S2 of the inner shield 130, dividing the second space S2 into a radially inner Dri region and a radially outer Dro region, i.e., the cavity CA. Multiple air holes 133h are formed on the collision plate 133, extending radially through Dr. Alternatively, the collision plate 133 may be omitted.

[0079] Multiple air passages 140 extend from the outer shroud 120 through the blade body 110 to the inner shroud 130. A portion of adjacent air passages 140 may communicate with each other in either a radially outer (Dro) or radially inner (Dri) portion. Any one of the multiple air passages 140 communicates with a first space S1 of the outer shroud 120. Any one of the multiple air passages 140 communicates with a second space S2 of the inner shroud 130. The air passages 140 communicate with multiple exhaust ports 110h of the blade body 110. The air passage 140 is an example of a "cooling medium passage".

[0080] A portion of the cooling air Ac flowing into the inner circumference of the turbine blade ring 70 through the air intake 72 flows into the air hole 123h of the collision plate 123 of the outer shroud 120, cooling the outer shroud 120 by flowing through the cavity CA of the outer shroud 120. Another portion of the cooling air Ac flowing into the inner circumference of the turbine blade ring 70 flows into the air passage 140, cooling the blade body 110 as it passes through the air passage 140. A portion of the cooling air Ac flowing into the air passage 140 is discharged into the combustion gas flow path 49 from multiple exhaust holes 110h provided on the blade body 110. Another portion of the cooling air Ac flowing into the air passage 140 flows into the second space S2 of the inner shroud 130, cooling the inner shroud 130. The cooling air Ac flowing through the second space S2 of the inner shroud 130 is discharged into the combustion gas flow path 49 from the exhaust hole 131h of the inner shroud 130.

[0081] (Dust separator)

[0082] Next, the dust separator 90 will be described.

[0083] like Figure 3 As shown, multiple dust separators 90 are disposed on the outer peripheral side of the turbine blade ring 70. The multiple dust separators 90 are arranged circumferentially Dc apart, located at positions corresponding to multiple air intakes 72. The dust separators 90 and the turbine blade ring 70 are formed of the same material (e.g., stainless steel) and have the same coefficient of thermal expansion. The dust separator 90 is an example of an "outer peripheral component." The dust separator 90 is an example of an "outer peripheral shroud," and an example of a "foreign matter separation shroud."

[0084] like Figure 4As shown, the dust separator 90 is configured to span the area between the first protrusion 73a and the second protrusion 73b, which are separated on both sides of the corresponding air intake 72. The dust separator 90 has a top wall 91, a rear wall 92, a first side wall 93, and a second side wall 94.

[0085] Figure 6 This is a top view showing the dust separator 90.

[0086] The top wall 91 is spaced apart from the outer peripheral surface 70o of the turbine blade ring 70 and extends parallel to the outer peripheral surface 70o of the turbine blade ring 70. That is, the top wall 91 has a gentle arc shape along the outer peripheral surface 70o of the turbine blade ring 70. The top wall 91 covers at least a portion of the air intake 72 from the outer peripheral side of the turbine blade ring 70. In this embodiment, the circumferential width W1 of the top wall 91 is greater than the maximum circumferential width W2 of the air intake 72. The top wall 91 extends from Dau, which is more upstream of the air intake 72, to Dad, which is more downstream of the air intake 72. The top wall 91 is an example of a "first wall portion".

[0087] In this embodiment, the top wall 91 does not cover the entire air intake 72, but only a portion of it. Specifically, the top wall 91 is positioned in a radially Dr region that does not overlap with the sealing member 85. For example, the top wall 91 has a cutout 95 in the radially Dr region where it overlaps with the sealing member 85. When removing the sealing member 85, it can be removed radially Dr towards the outer periphery of the turbine blade ring 70 through the inner side of the air intake 72 and the cutout 95.

[0088] More specifically, the top wall 91 has a first region 91a, a second region 91b, and a third region 91c. The first region 91a is the region covering the air intake 72. The first region 91a is located downstream of the axial direction Da in the axial direction Da, beyond the end face 73u of the protrusion 73. The second region 91b is disposed relative to the first region 91a on the circumferential direction Dcn. The third region 91c is disposed relative to the first region 91a on the circumferential direction Dcp. Compared to the first region 91a, the second region 91b and the third region 91c protrude upstream of the axial direction Dau. Thus, the region surrounded by the first region 91a, the second region 91b, and the third region 91c forms the cutout 95.

[0089] The first region 91a of the top wall 91 has an inclined portion 96. The inclined portion 96 is provided on the upstream side Dau of the first region 91a of the top wall 91. The inclined portion 96 is inclined away from the outer peripheral surface 70° of the turbine blade ring 70 as it advances toward the upstream side Dau (see reference). Figure 5 ).

[0090] The rear wall 92 extends from the downstream end of the top wall 91 (Dad) toward the outer circumferential surface 70° of the turbine blade ring 70. The rear wall 92 extends circumferentially (Dc). The rear wall 92, from the downstream end of the top wall (Dad), covers the space S between the top wall 91 and the air intake 72 (reference). Figure 5 The rear wall 92 (hereinafter referred to as "the internal space S") is at least a portion of the turbine blade ring 70. In this embodiment, the rear wall 92 contacts the turbine blade ring 70. That is, there is no gap between the rear wall 92 and the outer peripheral surface 70o of the turbine blade ring 70. Alternatively, there may be a gap between the rear wall 92 and the outer peripheral surface 70o of the turbine blade ring 70 (for example, a gap smaller than the plate thickness of the top wall 91). The rear wall 92 is an example of a "second wall portion".

[0091] Figure 7 This is the front view of the dust separator 90.

[0092] The first sidewall 93 extends from the end of the top wall 91 on the circumferential Dcn side on the circumferential Dc towards the outer peripheral surface 70o of the turbine blade ring 70. The first sidewall 93 extends along the axial direction Da. For example, the first sidewall 93 extends further upstream on the axial side Dau than a portion of the cutout 95. Furthermore, the first sidewall 93 extends further downstream on the axial side Dad than the air intake 72 and connects to the rear wall 92. The first sidewall 93 covers at least a portion of the internal space S of the dust separator 90 from the circumferential Dcn side. In this embodiment, the first sidewall 93 contacts the turbine blade ring 70. That is, there is no gap between the first sidewall 93 and the outer peripheral surface 70o of the turbine blade ring 70. Instead, there may be a gap between the first sidewall 93 and the outer peripheral surface 70o of the turbine blade ring 70 (e.g., a gap smaller than the plate thickness of the top wall 91). The first sidewall 93 is an example of a "third wall portion".

[0093] The second sidewall 94 extends from the end of the top wall 91 on the circumferential Dcp side toward the outer peripheral surface 70o of the turbine blade ring 70 on the circumferential Dc. The second sidewall 94 extends along the axial direction Da. For example, the second sidewall 94 extends further upstream on the axial side Dau than a portion of the cutout 95. Furthermore, the second sidewall 94 extends further downstream on the axial side Dad than the air intake 72 and connects to the rear wall 92. The second sidewall 94 covers at least a portion of the internal space S of the dust separator 90 from the circumferential Dcp side. In this embodiment, the second sidewall 94 is in contact with the turbine blade ring 70. That is, there is no gap between the second sidewall 94 and the outer peripheral surface 70o of the turbine blade ring 70. Instead, there may be a gap between the second sidewall 94 and the outer peripheral surface 70o of the turbine blade ring 70 (e.g., a gap smaller than the plate thickness of the top wall 91). The second sidewall 94 is an example of a "fourth wall portion".

[0094] With the structure described above, the internal space S of the dust separator 90 is surrounded by the top wall 91, the rear wall 92, the first side wall 93, and the second side wall 94. On the other hand, the internal space S of the dust separator 90 is open to the upstream side Dau of the axis.

[0095] Next, the connection structure CS of the fixed dust separator 90 will be explained.

[0096] like Figure 4 As shown, the connecting structure CS includes multiple (e.g., four) connecting portions 97. Each connecting portion 97 has: a support portion 97a, fixed to the outer peripheral surface 70o of the turbine blade ring 70; and a connecting tool 97b, detachably mounted on the support portion 97a. The support portion 97a is, for example, a boss provided on the outer peripheral surface 70o of the turbine blade ring 70, having a locking hole 97h opening radially outward. The top wall 91 of the dust separator 90 is supported on the support portion 97a. The top wall 91 has an insertion hole 91h at a position corresponding to the locking hole 97h through which the connecting tool 97b can be inserted. The connecting tool 97b is, for example, a bolt, which passes through the insertion hole 91h of the top wall 91 of the dust separator 90 and engages with the locking hole 97h of the support portion 97a. By engaging the connecting tool 97b with the locking hole 97h, the dust separator 90 is detachably connected to the outer peripheral surface 70o of the turbine blade ring 70.

[0097] (Shielding cover)

[0098] Figure 8 This is a cross-sectional view showing the shield 150. The shield 150 is mounted on the inner shield 130 and fixed to the inner shield 130 by welding or the like. The shield 150 covers at least a portion of the second space S2 of the inner shield 130 from the radially inner side Dri. In this embodiment, the shield 150 is mounted on the peripheral wall 132 of the inner shield 130 and covers the entire second space S2 of the inner shield 130. By mounting the shield 150 on the inner shield 130, it is not easy to supply compressed air Ac to the second space S2 of the inner shield 130 from the radially inner side Dri. Compressed air Ac flows in from the air intake 72 of the turbine blade ring 70 and flows through the air passage 140 of the outer shield 120 and the blade body 110 to the second space S2 of the inner shield 130. The shield 150 is an example of an "inner peripheral shield".

[0099] (Methods for modifying gas turbines)

[0100] Next, an example of a modification method for the gas turbine 10 will be described. Here, a modification method for adding a dust separator 90 and a shield 150 to a gas turbine 10 that does not already have a dust separator 90 and a shield 150 will be described. In addition, in the modification method for adding a shield 150 to a gas turbine 10 that already has a dust separator 90, the dust separator installation step (S12) described later is omitted.

[0101] Figure 9 This is a flowchart illustrating the steps of a method for modifying a gas turbine 10. The modification method of this embodiment includes, for example, a disassembly step (S11), a dust separator installation step (S12), a shield installation step (S13), and an assembly step (S14).

[0102] In the disassembly process (S11), necessary parts of the gas turbine 10 are disassembled. For example, the gas turbine housing 15 is disassembled, and the blade ring assembly WS is removed. Furthermore, in the blade ring assembly WS, multiple stationary blades 46a1 are removed from the turbine blade ring 70.

[0103] In the dust separator installation process (S12), the dust separator 90 is installed on the turbine blade ring 70. Specifically, a hole 70h for fixing the support portion 97a is provided on the outer peripheral surface 70o of the turbine blade ring 70 (see reference). Figure 7 The support portion 97a is fixed to the outer peripheral surface 70o of the turbine blade ring 70. Next, the dust separator 90 is placed on the support portion 97a. Then, the engaging tool 97b is passed through the insertion hole 91h of the top wall 91 of the dust separator 90, engaging it with the engagement hole 97h of the support portion 97a. Thus, the dust separator 90 is installed on the outer peripheral side of the turbine blade ring 70.

[0104] In the shielding installation process (S13), the shielding cover 150 is installed on the inner cover 130 of each stationary blade 46a1. For example, the shielding cover 150 is installed on the inner cover 130 by welding it to the peripheral wall 62 of the inner cover 130. As a result, the second space S2 of the inner cover 130 is closed by the shielding cover 150.

[0105] In the assembly process (S14), the gas turbine 10 is assembled. Specifically, multiple stationary blades 46a1, each equipped with a shield 150, are installed on the inner circumferential side of the turbine blade ring 70. Next, the gas turbine housing 15 is assembled onto the outer circumferential side of the turbine blade ring 70. Thus, the modification of the gas turbine 10 is completed.

[0106] (Effects)

[0107] Next, the function and effect of the dust separator 90 will be explained.

[0108] like Figure 2As shown, compressed air Ac, compressed by compressor 20, is supplied from compressor 20 to the housing C where burner 30 is located. The angle of the flow direction of the compressed air Ac supplied to housing C is changed by guide 64, flowing towards the peripheral wall 62 of gas turbine housing 15. Furthermore, the compressed air Ac that collides with the peripheral wall 62 is split in two. That is, a portion (e.g., most) of the compressed air Ac changes its flow direction towards the intake section 31 of burner 30, and flows towards the intake section 31 of burner 30 (see reference). Figure 2 (See arrow A1 in the image). On the other hand, after another portion of the compressed air Ac collides with the peripheral wall 62 of the gas turbine chamber 15, its flow direction changes in a manner that flows from the peripheral wall 62 along the rear wall 63 (see reference). Figure 2 (See arrow A2 in the diagram). As a result, inside the housing C, compressed air Ac is generated flowing from the downstream side Dad towards the upstream side Dau through a counter vortex V in the space located radially outward relative to the turbine blade ring 70. Furthermore, a portion of the compressed air Ac flowing in this vortex V enters the inner circumference of the turbine blade ring 70 from the air intake 72.

[0109] Figure 10 This is a cross-sectional view showing the function of the dust separator 90.

[0110] In this embodiment, a dust separator 90 is provided on the outer periphery of the turbine blade ring 70. The dust separator 90 has: a top wall 91 that covers at least a portion of the air intake 72 from the outer periphery of the turbine blade ring 70; and a rear wall 92 that covers the internal space S of the dust separator 90 from the downstream side Dad of the axis. Therefore, the compressed air Ac flowing along the rear wall 63 of the gas turbine housing 15 does not flow directly into the air intake 72, but flows from the downstream side Dad toward the upstream side Dau of the axis through the space located radially outward relative to the dust separator 90 Dro, and then turns at an angle greater than 90 degrees (e.g., about 180 degrees) to flow into the air intake 72 (see reference). Figure 10 (See arrow A3 in the diagram). During this process, foreign matter M (e.g., rust or dust) contained in the compressed air Ac flows to the upstream side Dau of the axis, which is further upstream than the air intake 72, due to the inertial force of the foreign matter M. Therefore, foreign matter M contained in the compressed air Ac is less likely to enter the air intake 72. As a result, the blockage of foreign matter M in the cooling structure of the stationary blade 46a1 (e.g., the exhaust port 110h of the blade body 110 or the exhaust port 131h of the inner shroud 130) is prevented.

[0111] Foreign matter M, flowing by inertial force to a position Dau further upstream on the axis than the air intake 72, enters the combustor 30 from the intake 31, is contained in the combustion gas G, and is discharged to the outside of the gas turbine 10. Since the combustion gas flow path 49 has fewer or no orifices blocked by foreign matter compared to the stationary blades 46a1, adverse conditions caused by foreign matter M are less likely to occur.

[0112] <Second Implementation>

[0113] Next, the second embodiment will be described. The difference between the second embodiment and the first embodiment is that the blade ring assembly WS replaces the dust separator 90 and includes the L-shaped tube 290. Otherwise, the structure is the same as the first embodiment, except for the following description.

[0114] (Turbine blade ring)

[0115] Figure 11 This is a perspective view schematically showing a portion of the blade ring assembly WS according to the second embodiment of the present invention.

[0116] Figure 12 This is an enlarged and schematic cross-sectional view showing a portion of the blade ring assembly WS according to the second embodiment of the present invention.

[0117] like Figure 11 , Figure 12 As shown, an air intake 72 is formed on the outer peripheral surface 70o of Dad, downstream of the axis of the blade ring fixing portion 71 near the turbine blade ring 70. The air intake 72 is located in the radial direction Dr, overlapping with the dividing ring 45c between the first and second layer stationary blades 46a located in the axial direction Da.

[0118] The second impact plate 223 is disposed in a third space S3 between the dividing ring 45c, which overlaps with the air intake 72 in the radial direction (Dr), and the inner circumferential surface 70i of the turbine blade ring 70. The second impact plate 223 divides the third space S3 into a region of radially outer (Dro) and a region of radially inner (Dri). A plurality of second air holes 223h are formed on the second impact plate 223, which extends radially through the turbine blade ring 70. The second impact plate 223 and the dividing ring 45c, which overlaps with the second impact plate 223 in the radial direction (Dr), are disposed on the inner circumferential side of the turbine blade ring 70, and are an example of a "cooling target component".

[0119] (L-shaped tube)

[0120] An L-shaped tube 290 is disposed on the outer periphery of the turbine blade ring 70. More specifically, the L-shaped tube 290 is disposed at a position on the outer peripheral surface 70o of the turbine blade ring 70 that overlaps with the air intake 72. The L-shaped tube 290 is formed, for example, of the same material as the turbine blade ring 70. The L-shaped tube 290 and the turbine blade ring 70 are formed, for example, of a metallic material such as stainless steel. When viewed from the circumferential direction Dc, the L-shaped tube 290 is formed in an L-shape. The L-shaped tube 290 has a first tube portion 290a and a second tube portion 290b.

[0121] The first tube section 290a extends radially Dr from the cooling medium inlet 72 along the turbine blade ring 70.

[0122] The second tube portion 290b extends from the end of the first tube portion 290a on the side opposite to the turbine blade ring 70 toward the side opposite to the downstream side of the axis, Dad. In this embodiment, the second tube portion 290b extends from the first tube portion 290a toward the upstream side of the axis on both sides in the axial direction Da, Dau. Furthermore, the second tube portion 290b may be inclined relative to the axis Ar at an angle within ±10 degrees, for example. The second tube portion 290b communicates with the first tube portion 290a and has an opening 290c on the side opposite to the first tube portion 290a. The opening 290c opens toward the upstream side of the axis, Dau.

[0123] On the other hand, the L-shaped tube 290 described above has a top wall 291, a rear wall 292, a first side wall 293, and a second side wall 294.

[0124] The top wall 291 extends along the outer peripheral surface 70° of the turbine blade ring 70, with a gap between it and the outer peripheral surface 70°. The top wall 291 covers at least a portion of the air intake 72 from the outer peripheral side of the turbine blade ring 70. In this embodiment, the top wall 291 covers the entire air intake 72 from the outer peripheral side of the turbine blade ring 70. The top wall 291 extends to the upstream side Dau of the centerline on both sides in the axial direction Da, further than the air intake 72. The top wall 291 is an example of a "first wall portion".

[0125] The rear wall 292 extends from the end of the Dad downstream of the axis of the top wall 291 toward the outer peripheral surface 70o of the turbine blade ring 70. The rear wall 292 extends from the end of the Dad downstream of the axis of the top wall 291 toward the outer peripheral surface 70o of the turbine blade ring 70. The rear wall 292 covers at least a portion of the space S between the top wall 291 and the air intake 72 from the Dad downstream of the axis. In this embodiment, the rear wall 292 covers the entire space S between the top wall 291 and the air intake 72 from the Dad downstream of the axis. The rear wall 292 is an example of a "second wall portion".

[0126] The first sidewall 293 extends from the circumferentially Dcn-side end of the top wall 291 on the circumferential Dc toward the outer circumferential surface 70o of the turbine blade ring 70. The first sidewall 293 covers at least a portion of the space S between the top wall 291 and the air medium inlet 72 from the circumferential Dcn. When viewed from the circumferential Dc, the first sidewall 293 is L-shaped. The first sidewall 293 is an example of a "third wall portion".

[0127] The second sidewall 294 extends from the circumferentially Dcp-side end of the top wall 291 on the circumferential Dc toward the outer circumferential surface 70o of the turbine blade ring 70. The second sidewall 294 covers at least a portion of the space S between the top wall 291 and the air medium inlet 72 from the circumferential Dcp. When viewed from the circumferential Dc, the second sidewall 294 is L-shaped. The second sidewall 294 is another example of a "third wall portion".

[0128] Figure 13 This is an exploded perspective view of the L-shaped tube 290 according to the second embodiment of the present invention.

[0129] On the other hand, such as Figure 13 As shown, the L-shaped tube 290 has a bend 295 and a loading / unloading tube 296.

[0130] (Bend)

[0131] The bend 295 has a cylindrical portion 295a and a connecting portion 295b. The cylindrical portion 295a is a tube that extends radially Dr.

[0132] The inner radial end of the cylindrical portion 295a (Dri) is open. The outer radial end of the cylindrical portion 295a (Dro) is closed. The cylindrical portion 295a is inserted into the air intake 72. The cylindrical portion 295a is welded to the entire circumference of the turbine blade ring 70, for example, with the cylindrical portion 295a screwed into the air intake 72. A connecting portion 295b is provided at the outer radial end of the cylindrical portion 295a on the side of the upstream side (Dau) of the axis.

[0133] The connecting portion 295b is formed as a cylinder extending along the axial direction Da. The connecting portion 295b has a quadrilateral cylindrical shape, and when viewed from the axial direction Da, its inner circumferential surface is circular. The connecting portion 295b communicates with the cylindrical portion 295a. The connecting portion 295b opens towards the upstream side Dau of the axial direction. This opening on the upstream side Dau of the connecting portion 295b is designated as the bend opening 295c.

[0134] The loading / unloading pipe 296 is connected to the bend opening 295c of the bend pipe 295. The loading / unloading pipe 296 has a large-diameter pipe 296a and a throttling orifice 296b.

[0135] The large-diameter pipe 296a is positioned upstream of the bend 295 on the axial side, at a position called Dau. The large-diameter pipe 296a is formed into a cylindrical shape extending along the axial direction Da. Both ends of the large-diameter pipe 296a are open in the axial direction Da. The opening upstream of the axial side Dau in the opening of the large-diameter pipe 296a is the opening 290c of the L-shaped pipe 290 mentioned above.

[0136] A throttling orifice 296b is located at the downstream end of Dad on the axial side of the large-diameter pipe 296a. The throttling orifice 296b is formed as a cylinder extending along the axial direction Da. The central axis of the throttling orifice 296b is aligned with the central axis of the large-diameter pipe 296a.

[0137] The outer diameter of the throttling orifice 296b is smaller than that of the large-diameter pipe 296a, and the inner diameter of the throttling orifice 296b is also smaller than that of the large-diameter pipe 296a. That is, the inner diameter of the throttling orifice 296b is smaller than the inner diameter of the opening 290c of the upstream side of the L-shaped pipe 290's axis. The throttling orifice 296b is formed as if inserted into the large-diameter pipe 296a. The throttling orifice 296b and the large-diameter pipe 296a are integrally formed.

[0138] The throttling orifice 296b is inserted into the bend opening 295c of the bend 295. A parallel thread (not shown) is formed on the outer circumferential surface of the throttling orifice 296b. A parallel thread groove (not shown) is formed on the inner circumferential surface of the connecting portion 295b of the bend 295. The throttling orifice 296b is spot-welded to the bend 295, for example, while screwed into the bend opening 295c.

[0139] Furthermore, the orifice portion 296b is formed such that its flow path area is smaller than that of other portions of the L-shaped tube 290. In this embodiment, the orifice portion 296b is formed such that its flow path area is smaller than that of all portions of the L-shaped tube 290 except for the orifice portion 296b.

[0140] Furthermore, the orifice portion 296b is detachably provided on other parts of the L-shaped tube. In this embodiment, the orifice portion 296b is spot-welded to the bend 295. Therefore, compared to the case where the orifice portion 296b is welded to the entire circumference of the bend 295, the orifice portion 296b can be easily installed and removed relative to the bend 295. Thus, the orifice portion 296b can be fully installed and removed relative to the bend 295.

[0141] (Effects)

[0142] Next, the function and effect of the L-shaped tube 290 will be explained.

[0143] Figure 14 This is a cross-sectional view showing the function of the L-shaped tube 290.

[0144] In this embodiment, an L-shaped tube 290 is provided on the outer periphery of the turbine blade ring 70. The L-shaped tube 290 has: a top wall 291 that covers at least a portion of the air intake 72 from the outer periphery of the turbine blade ring 70; and a rear wall 292 that covers the space S between the top wall 291 and the air intake 72 from the downstream side of the axis (Dad). Therefore, the compressed air Ac flowing along the rear wall 63 of the gas turbine housing 15 does not flow directly into the air intake 72, but flows from the downstream side (Dad) toward the upstream side (Dau) through the space located radially outward (Dro) relative to the L-shaped tube 290, and then turns at an angle greater than 90 degrees (e.g., approximately 180 degrees) to flow into the air intake 72 (see reference). Figure 14 (See arrow A4 in the diagram). During this process, foreign matter M (e.g., rust or dust) contained in the compressed air Ac flows to the upstream side Dau of the axis, which is further upstream than the air intake 72, due to the inertial force of the foreign matter M. Therefore, foreign matter M contained in the compressed air Ac is less likely to enter the air intake 72. As a result, the blockage of foreign matter M in the second air hole 223h of the second impact plate 223 or in the structure formed by the second impact plate 223 and the dividing ring 45c is prevented.

[0145] (Other implementation methods)

[0146] The embodiments of the present invention have been described in detail above with reference to the accompanying drawings. However, the specific structure is not limited to this embodiment, and design changes that do not depart from the scope of the present invention are also included. For example, the dust separator 90 is not limited to a component mounted on the turbine blade ring 70 via the connecting structure CS, but can also be integrally cast with the turbine blade ring 70. The first sidewall 93 and the second sidewall 94 of the dust separator 90 may also be omitted. However, if the dust separator 90 has the first sidewall 93 and the second sidewall 94, it is possible to further suppress foreign matter from entering the stationary blade 46a. The dust separator 90 may not have the inclined portion 96. The blade ring assembly WS may not have the shield 150.

[0147] Furthermore, for example, in the L-shaped pipe 290, it is not limited to the case where a throttling orifice 296b is provided in the loading and unloading pipe 296; the bend pipe 295 can also have the function of the throttling orifice 296b.

[0148] <Note>

[0149] The blade ring assembly WS, gas turbine 10, and modification method of gas turbine 10 described in the embodiments can be understood as follows, for example.

[0150] (1) The blade ring assembly WS according to the first embodiment includes: a turbine blade ring 70 extending circumferentially in the direction Dc centered on the axis Ar; cooling target components (stationary blade 46a1, second collision plate 223, dividing ring 45c) disposed on the inner circumferential side of the turbine blade ring 70; and outer circumferential components (dust separator 90, L-shaped tube 290) disposed on the outer circumferential side of the turbine blade ring 70. The turbine blade ring 70 has a cooling medium inlet (air inlet 72) extending from the outer circumferential surface 70o of the turbine blade ring 70 to the inner circumferential surface 70i. The outer peripheral cover has: a first wall portion (top wall 91, top wall 291) that covers at least a portion of the cooling medium inlet from the outer peripheral side of the turbine blade ring 70 and extends to the downstream side Dad on both sides of the central axis in the axial direction Da that extends further along the axis Ar than the cooling medium inlet; and a second wall portion (rear wall 92, rear wall 292) that extends from the end of the downstream side Dad of the first wall portion toward the outer peripheral surface 70o of the turbine blade ring 70 and covers at least a portion of the space S between the first wall portion and the cooling medium inlet from the downstream side Dad.

[0151] According to this structure, by utilizing the flow of cooling medium within the gas turbine chamber 15, in which the blade ring assembly WS is configured, the foreign object and the cooling medium can be separated by the inertial force acting on the foreign object. This prevents the foreign object from flowing into the component to be cooled. Specifically, the inventors analyzed the flow of compressed air Ac within the containment chamber C and found that around the cooling medium inlet of the turbine blade ring 70, there exists a counter vortex of compressed air Ac flowing from the downstream side Dad towards the upstream side Dau of the axis. The structure of the first embodiment described above utilizes this counter vortex to efficiently separate the foreign object and the cooling medium using a relatively simple-shaped outer peripheral component.

[0152] (2) The blade ring assembly WS involved in the second method is the blade ring assembly WS described in (1), and the outer peripheral component also has a third wall portion (first side wall 93, first side wall 293), which extends from the end of the first wall portion on the circumferential direction Dc toward the outer peripheral surface 70o of the turbine blade ring 70 and covers at least a portion of the space S from the circumferential direction Dc.

[0153] This structure suppresses the flow of cooling medium from the circumferential direction (Dc) into the inner side of the outer peripheral cover. Consequently, foreign matter and cooling medium can be separated with a higher probability, further preventing foreign matter from entering the component being cooled.

[0154] (3) The blade ring assembly WS involved in the third method is the blade ring assembly WS described in (1) or (2), and the outer peripheral component can be an outer peripheral cover (dust separator 90) disposed on the outer peripheral side of the turbine blade ring 70.

[0155] Based on this structure, foreign objects and cooling media can be efficiently separated through the relatively simple-shaped outer peripheral shield.

[0156] (4) The blade ring assembly WS involved in the fourth method is the blade ring assembly WS described in (3). The turbine blade ring 70 further has: a first protrusion 73a, protruding towards the upstream side Dau of the central axis on both sides in the axial direction Da; and a second protrusion 73b, disposed at a different position from the first protrusion 73a in the circumferential direction Dc, and protruding towards the upstream side Dau of the axis. The cooling medium inlet is located in the circumferential direction Dc between the first protrusion 73a and the second protrusion 73b. The outer peripheral shroud is configured to span the area between the first protrusion 73a and the second protrusion 73b.

[0157] With this structure, a first wall portion extending relatively far upstream on the axis can be provided across the region between the first protrusion 73a and the second protrusion 73b. This allows for a higher probability of separating foreign matter and the cooling medium. Consequently, it further suppresses the inflow of foreign matter into the component being cooled.

[0158] (5) The blade ring assembly WS involved in the fifth method is the blade ring assembly WS described in (4), wherein the first protrusion 73a and the second protrusion 73b have a connecting part (an end face 73u including a connecting hole 73h), which can be connected to the fixing part (burner connecting part 80) of the tail tube 32 of the fixed canister burner (burner 30).

[0159] With this structure, by utilizing the first protrusion 73a and the second protrusion 73b, which are provided with a connecting portion for connection to the fixed component, a first wall portion that extends relatively far upstream on the axis can be provided. Therefore, without the need for additional special protrusions, foreign matter and cooling medium can be separated with a higher probability, further suppressing the inflow of foreign matter into the component to be cooled.

[0160] (6) The blade ring assembly WS involved in the sixth embodiment is the blade ring assembly WS described in any one of (3) to (6), wherein the turbine blade ring 70 and the outer peripheral shroud are formed of the same material. With this structure, the effect of thermal expansion of the turbine blade ring 70 acting between the turbine blade ring 70 and the outer peripheral shroud can be reduced. Therefore, adverse conditions at the fixing portion between the turbine blade ring 70 and the outer peripheral shroud are suppressed, and the lifespan of the blade ring assembly WS can be extended.

[0161] (7) The blade ring assembly WS involved in the seventh method is the blade ring assembly WS described in any one of (3) to (6), and may also include a connecting structure CS that detachably connects the outer peripheral shroud to the turbine blade ring 70. Here, for example, if the outer peripheral shroud is directly fixed to the turbine blade ring 70 by welding, cracks may occur at the weld between the turbine blade ring 70 and the outer peripheral shroud due to the thermal expansion of the turbine blade ring 70. On the other hand, according to the fixing method using the connecting structure CS as described above, cracks can be suppressed at the weld between the turbine blade ring 70 and the outer peripheral shroud. As a result, the service life of the blade ring assembly WS can be extended.

[0162] (8) The blade ring assembly WS involved in the eighth embodiment is the blade ring assembly WS of any one of (3) to (7), and further includes a removable component (sealing component 85) disposed on the inner circumferential side of the turbine blade ring 70, overlapping with the cooling medium inlet on the radial Dr of the turbine blade ring 70, and capable of being removed to the outer circumferential side of the turbine blade ring 70 through the cooling medium inlet. The outer circumferential side cover is disposed on the radial Dr in a region that does not overlap with the removable component.

[0163] With this structure, even with an outer peripheral cover, removable components can be disassembled to the outer peripheral side via the cooling medium inlet. This maintains or improves the workability of the assembly work related to the blade ring assembly (WS) and prevents foreign matter from entering the component being cooled.

[0164] (9) The blade ring assembly WS according to the ninth embodiment is the blade ring assembly WS of any one of (3) to (7), and further includes a removable component (sealing component 85) disposed on the inner circumferential side of the turbine blade ring 70, overlapping with the cooling medium inlet on the radial Dr of the turbine blade ring 70, and capable of being detached to the outer circumferential side of the turbine blade ring 70 through the cooling medium inlet. The outer circumferential side cover has a cutout 95 in the area on the radial Dr where it overlaps with the removable component.

[0165] With this structure, even with an outer peripheral cover, removable components can be disassembled to the outer peripheral side via the cooling medium inlet. This maintains or improves the workability of the assembly work related to the blade ring assembly (WS) and prevents foreign matter from entering the component being cooled.

[0166] (10) The blade ring assembly WS involved in the 10th method is any one of (3) to (9) and at least a portion of the first wall has an inclined portion 96 that is inclined in a direction away from the outer peripheral surface 70o of the turbine blade ring 70 as it moves toward the upstream side of the axis Dau.

[0167] With this structure, the flow direction of compressed air Ac near the first wall can be changed to a direction away from the cooling medium inlet. As a result, foreign objects subjected to inertial forces are less likely to enter the cooling medium inlet. Consequently, the inflow of foreign objects into the component to be cooled can be further suppressed.

[0168] (11) The blade ring assembly WS involved in the 11th method is the blade ring assembly WS of any one of (1) to (10), and the cooling target component is the stationary blade 46a1 disposed on the inner circumferential side of the turbine blade ring.

[0169] Here, the stationary blade 46a1 has a certain volume and has a greater impact on the performance of the gas turbine 10 compared to other parts. According to the structure of this method, foreign matter can be prevented from entering the stationary blade 46a1, thus enabling efficient cooling of the stationary blade 46a1 and further suppressing the performance degradation of the gas turbine 10.

[0170] (12) The blade ring assembly WS involved in the 12th method is the blade ring assembly WS described in (11), wherein the stationary blade 46a1 has: a blade body 110 disposed in the combustion gas flow path 49 and in the shape of a blade; an outer shield 120 disposed at the end of the outer peripheral side of the blade body 110 and having a first space S1; an inner shield 130 disposed at the end of the inner peripheral side of the blade body 110 and having a second space S2; a cooling medium passage (air passage 140) extending from the outer shield 120 through the blade body 110 to the inner shield 130; and an inner peripheral side cover (shield 150) mounted on the inner shield 130 and covering at least a portion of the second space S2 from the inner peripheral side.

[0171] This structure prevents the cooling medium from flowing into the interior of the stationary blade 46a1 from the inner circumferential side. A large amount of cooling medium, separated from foreign matter by the outer circumferential side cover, is supplied to the interior of the stationary blade 46a1. This further suppresses the inflow of foreign matter into the stationary blade 46a1.

[0172] (13) The blade ring assembly WS involved in the 13th embodiment is the blade ring assembly WS described in (1) or (2), and the outer peripheral component is an L-shaped tube 290. The L-shaped tube 290 has: a first tube portion 290a extending radially Dr from the cooling medium inlet along the turbine blade ring 70; and a second tube portion 290b extending from the end of the first tube portion 290a on the side opposite to the turbine blade ring 70 toward the side opposite to the downstream side Dad of the axis. The second tube portion 290b communicates with the first tube portion 290a and has an opening 290c on the side opposite to the first tube portion 290a.

[0173] Based on this structure, foreign matter and cooling medium can be efficiently separated through the relatively simple-shaped L-shaped tube 290.

[0174] (14) The blade ring assembly WS involved in the 14th method is the blade ring assembly WS described in (13), wherein the second tube 290b extends from the first tube 290a toward the upstream side of the central axis on both sides of the axis direction Da, and the opening 290c opens toward the upstream side of the axis Da.

[0175] This structure further suppresses the inflow of foreign matter into the L-shaped tube 290. Consequently, it enables more efficient separation of foreign matter and cooling medium.

[0176] (15) The blade ring assembly WS involved in the 15th method is the blade ring assembly WS described in (13) or (14), and the L-shaped tube 290 has a throttling orifice 296b, the flow path area of ​​which is smaller than the other parts of the L-shaped tube 290.

[0177] With this structure, the flow rate of the cooling medium flowing into the component to be cooled can be appropriately adjusted by using the throttling orifice 296b. As a result, the component to be cooled can be cooled efficiently.

[0178] (16) The blade ring assembly WS involved in the 16th method is the blade ring assembly WS described in (15), with the throttling orifice portion 296b detachably provided in other parts of the L-shaped tube 290.

[0179] With this structure, the throttling orifice 296b can be easily replaced. Therefore, the inner diameter of the throttling orifice 296b can be easily changed.

[0180] (17) The blade ring assembly WS involved in the 17th method is the blade ring assembly WS described in (16), wherein the inner diameter of the throttling orifice 296b is smaller than the inner diameter of the opening 290c.

[0181] With this structure, the opening 290c, located upstream of the throttling orifice 296b, is larger than the inner diameter of the orifice 296b. Therefore, when the cooling medium flows into the L-shaped tube 290 from the opening 290c, the increase in the flow rate of the cooling medium can be suppressed. Thus, it is possible to prevent foreign objects from being drawn into the L-shaped tube 290.

[0182] (18) The blade ring assembly WS involved in the 18th embodiment is the blade ring assembly WS described in any one of (13) to (17), wherein the turbine blade ring 70 and the L-shaped tube 290 are formed of the same material. With this structure, the effect of thermal expansion of the turbine blade ring 70 acting between the turbine blade ring 70 and the L-shaped tube 290 can be reduced. Therefore, adverse conditions at the fixing portion of the turbine blade ring 70 and the L-shaped tube 290 are suppressed, and the lifespan of the blade ring assembly WS can be extended.

[0183] (19) The gas turbine 10 according to the 19th embodiment includes: a blade ring assembly WS as described in any one of (1) to (18); a rotor (gas turbine rotor 11) rotatable about an axis Ar; a housing (gas turbine chamber 15) covering the outer periphery of the rotor; and a canister burner (burner 30) that generates the combustion gas by combustion of fuel and delivers the combustion gas into the housing. The blade ring assembly WS is disposed on the inner periphery of the housing. The housing has a partition that, as part of the wall of a receiving chamber C that defines the exposure of the blade ring assembly WS and the flow of the cooling medium, is disposed on a downstream side Dad of the axis and extends radially Dr along the turbine blade ring 70. According to this structure, in the gas turbine 10, foreign matter and the cooling medium can be separated by inertial force, and foreign matter can be prevented from flowing into the cooling target component.

[0184] (20) The gas turbine 10 involved in the 20th method is the gas turbine 10 described in (19), wherein the cooling target component included in the blade ring assembly WS constitutes the first layer of turbine stator blades in the axial direction Da. Here, compared with the turbine stator blades of the second layer and beyond, the combustion gas G is more likely to rise in temperature in the first layer of turbine stator blades, which has a greater impact on the performance of the gas turbine 10. According to the structure of this method, foreign matter can be prevented from flowing into the first layer of turbine stator blades, thus the first layer of turbine stator blades can be cooled efficiently, and the performance degradation of the gas turbine 10 can be further suppressed.

[0185] (21) In the modification method of the gas turbine 10 according to the 21st method, the gas turbine 10 has a turbine blade ring 70 and a stationary blade 46a1 disposed on the inner circumferential side of the turbine blade ring 70. The modification method of the gas turbine 10 includes the following steps: removing the stationary blade 46a1 from the turbine blade ring 70; installing a shield 150 on the turbine stationary blade; and installing the stationary blade 46a1 with the shield 150 installed on the turbine blade ring 70 with a foreign matter separation cover (dust separator 90) installed. The turbine stationary blade has: a blade body 110 disposed in the combustion gas flow path 49 and in the shape of a blade; and an inner shield 130 disposed at the end of the inner circumferential side of the blade body 110 and having a space (second space S2). By installing the shield 150 on the inner shield 130, the shield 150 covers at least a portion of the space of the inner shield 130 from the inner circumferential side. The turbine blade ring 70 has a cooling medium inlet (air inlet 72) extending from the outer peripheral surface 70o of the turbine blade ring 70 to the inner peripheral surface 70i. A foreign matter separation cover is disposed on the outer peripheral side of the turbine blade ring 70, covering at least a portion of the cooling medium inlet.

[0186] This structure prevents the cooling medium from flowing into the interior of the stationary blade 46a1 from the inner circumferential side. A large amount of cooling medium, separated from foreign matter by the outer circumferential side cover, is supplied to the interior of the stationary blade 46a1. This further suppresses the inflow of foreign matter into the stationary blade 46a1.

[0187] Industrial availability

[0188] According to the blade ring assembly, gas turbine, and gas turbine modification method of the present invention, it is possible to suppress the inflow of foreign matter into components such as stationary blades that are subject to cooling.

[0189] 10-Gas turbine, 11-Gas turbine rotor, 15-Gas turbine housing, 20-Compressor, 30-Burner, 41-Turbine rotor, 45-Turbine housing, 46a1-Stationary blade, 61-Front wall, 62-Peripheral wall, 63-Rear wall, 70-Turbine blade ring, 70o-Outer peripheral surface, 70i-Inner peripheral surface, 72-Air intake, 73a-First protrusion, 73b-Second protrusion, 73h-Matching hole, 80-Burner connection component, 81-Frame, 82-Flange, 90-Dust separator, 91-Top wall, 92-Rear wall, 93-First side wall, 94-Second side wall, 95-Slit, 96-Inclined Part, 110-blade body, 120-outer shroud, 130-inner shroud, 140-air passage, 150-shield, WS-blade ring assembly, 223-second collision plate, 223h-second air hole, 290-L-shaped tube, 290a-first tube section, 290b-second tube section, 290c-opening section, 291-top wall, 292-rear wall, 293-first side wall, 294-second side wall, 295-bend, 295a-cylindrical section, 295b-connecting section, 296-loading and unloading pipe, 296a-large diameter pipe, 296b-throttle orifice section, S1-first space, S2-second space, S3-third space.

Claims

1. A blade ring assembly, comprising: Turbine blade rings extend circumferentially around the axis. The cooling component is disposed on the inner circumferential side of the turbine blade ring; and The outer peripheral component is disposed on the outer peripheral side of the turbine blade ring. The turbine blade ring has a cooling medium inlet that extends from the outer circumferential surface of the turbine blade ring to the inner circumferential surface. The outer peripheral component has: The first wall portion covers at least a portion of the cooling medium inlet from the outer peripheral side of the turbine blade ring and extends to the downstream side of the central axis on both sides in the axial direction extending further along the axis than the cooling medium inlet; and The second wall portion extends from the downstream end of the first wall portion toward the outer peripheral surface of the turbine blade ring, covering at least a portion of the space between the first wall portion and the cooling medium inlet from the downstream side of the axis. The outer peripheral component is an outer peripheral cover disposed on the outer peripheral side of the turbine blade ring. The turbine blade ring further comprises: a first protrusion protruding towards the upstream side of the central axis on both sides in the axial direction; and a second protrusion disposed at a different position from the first protrusion in the circumferential direction and protruding towards the upstream side of the axis. The cooling medium inlet is located in the circumferential direction between the first protrusion and the second protrusion. The outer peripheral cover is configured to span the area between the first protrusion and the second protrusion.

2. The blade ring assembly according to claim 1, wherein, The outer peripheral component also has a third wall portion that extends from the circumferential end of the first wall portion toward the outer peripheral surface of the turbine blade ring and covers at least a portion of the space in the circumferential direction.

3. The blade ring assembly according to claim 1, wherein, The first protrusion and the second protrusion have connecting parts that can be connected to the fixing parts of the tail tube of the fixed canister burner.

4. The blade ring assembly according to claim 1, wherein, The turbine blade ring and the outer peripheral shroud are formed of the same material.

5. The blade ring assembly according to claim 1, further comprising a connecting structure that detachably connects the outer peripheral shroud to the turbine blade ring.

6. A blade ring assembly, comprising: Turbine blade rings extend circumferentially around the axis. The cooling component is disposed on the inner circumferential side of the turbine blade ring; and The outer peripheral component is disposed on the outer peripheral side of the turbine blade ring. The turbine blade ring has a cooling medium inlet that extends from the outer circumferential surface of the turbine blade ring to the inner circumferential surface. The outer peripheral component has: The first wall portion covers at least a portion of the cooling medium inlet from the outer peripheral side of the turbine blade ring and extends to the downstream side of the central axis on both sides in the axial direction extending further along the axis than the cooling medium inlet; and The second wall portion extends from the downstream end of the first wall portion toward the outer peripheral surface of the turbine blade ring, covering at least a portion of the space between the first wall portion and the cooling medium inlet from the downstream side of the axis. The outer peripheral component is an outer peripheral cover disposed on the outer peripheral side of the turbine blade ring. The blade ring assembly also includes a detachable component disposed on the inner circumferential side of the turbine blade ring, overlapping the cooling medium inlet in the radial direction of the turbine blade ring, and capable of being detached from the outer circumferential side of the turbine blade ring through the cooling medium inlet. The outer peripheral cover is positioned radially in a region that does not overlap with the removable component.

7. A blade ring assembly, comprising: Turbine blade rings extend circumferentially around the axis. The cooling component is disposed on the inner circumferential side of the turbine blade ring; and The outer peripheral component is disposed on the outer peripheral side of the turbine blade ring. The turbine blade ring has a cooling medium inlet that extends from the outer circumferential surface of the turbine blade ring to the inner circumferential surface. The outer peripheral component has: The first wall portion covers at least a portion of the cooling medium inlet from the outer peripheral side of the turbine blade ring and extends to the downstream side of the central axis on both sides in the axial direction extending further along the axis than the cooling medium inlet; and The second wall portion extends from the downstream end of the first wall portion toward the outer peripheral surface of the turbine blade ring, covering at least a portion of the space between the first wall portion and the cooling medium inlet from the downstream side of the axis. The outer peripheral component is an outer peripheral cover disposed on the outer peripheral side of the turbine blade ring. The blade ring assembly also includes a detachable component disposed on the inner circumferential side of the turbine blade ring, overlapping the cooling medium inlet in the radial direction of the turbine blade ring, and capable of being detached from the outer circumferential side of the turbine blade ring through the cooling medium inlet. The outer peripheral cover has a cutout in the region where it overlaps with the removable component in the radial direction.

8. The blade ring assembly according to any one of claims 1 to 7, wherein, At least a portion of the first wall has an inclined portion that tilts toward the outer peripheral surface of the turbine blade ring as it advances toward the upstream side of the central axis on both sides in the axial direction.

9. The blade ring assembly according to any one of claims 1 to 7, wherein, The cooling target component is a stationary blade disposed on the inner circumferential side of the turbine blade ring.

10. The blade ring assembly according to claim 9, further comprising an inner circumferential side cover, the inner circumferential side cover being mounted on the stationary blade. The stationary blade has: The blade body is positioned within the combustion gas flow path and is blade-shaped. An outer protective cover is provided at the end of the outer peripheral side of the blade body and has a first space; An inner protective cover is disposed at the end of the inner circumferential side of the blade body and has a second space; and Cooling medium channels extend from the outer shroud, through the blade body, to the inner shroud. The inner circumferential side cover is mounted on the inner side cover and covers at least a portion of the second space from the inner circumferential side of the blade body.

11. A gas turbine comprising: The blade ring assembly according to any one of claims 1 to 10; The rotor is capable of rotating about the axis described above. Housing, covering the outer periphery of the rotor; and A canister burner generates combustion gases through the combustion of fuel and then delivers these combustion gases into the housing. The blade ring assembly is disposed on the inner circumferential side of the housing. The housing has a partition that serves as part of the wall of a accommodating chamber that defines the exposed blade ring assembly and the flow of cooling medium. The partition is located downstream of the axis and extends radially along the turbine blade ring.

12. The gas turbine according to claim 11, wherein, The cooling object component included in the blade ring assembly constitutes the first layer of turbine stationary blades in the axial direction.

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