Blade ring assembly, and gas turbine equipped therewith

Abstract

A blade ring assembly according to the present disclosure comprises a turbine blade ring extending in a circumferential direction about an axis, stationary blades disposed on the inner-circumferential side of the turbine blade ring, and a deflector disposed on the outer-circumferential side of at least part of the turbine blade ring. The turbine blade ring has: a blade ring body; and a plurality of protrusions that protrude from the blade ring body toward an axis-upstream side among two sides in an axial direction in which the axis extends, the plurality of protrusions being disposed at different positions around the circumferential direction at intervals from each other. Two protrusions that are adjacent in the circumferential direction among the plurality of protrusions, together with the blade ring body, form a cooling medium intake port that passes through the turbine blade ring from the outer-circumferential side to the inner-circumferential side, and the deflector is disposed so as to block at least part of the cooling medium intake port as seen from a radial direction centered on the axis.

Description

Technical Field

[0001] The present disclosure relates to a vane ring assembly and a gas turbine including the same. This application claims priority to Japanese Patent Application No. 2022-165186 filed on Oct. 14, 2022, the contents of which are incorporated herein by reference.

Background Art

[0002] Patent Document 1 discloses a gas turbine including a mechanism for separating particles contained in cooling air. This gas turbine includes a particle separation protection element disposed on the inner circumferential side (rotor side) rather than the stator vane, and this particle separation protection element makes it difficult for floating particles to flow into the intake opening.

Prior Art Documents

Patent Documents

[0003]

Patent Document 1

Summary of the Invention

Problems to be Solved by the Invention

[0004] However, the particle separation protection element described in Patent Document 1 does not suppress the cooling air supplied from the compressor into the casing from flowing directly into the air passage leading to the inside of the stator vane. Therefore, cooling air containing foreign matter may flow into the inside of the stator vane, and clogging may occur in the cooling passages inside the stator vane. <​​​​​​​​​To solve the above problems, the blade ring assembly according to the present disclosure comprises a turbine blade ring extending in the circumferential direction about an axis, a stationary blade disposed on the inner circumference side of the turbine blade ring, and a deflector disposed on the outer circumference side of at least a portion of the turbine blade ring, wherein the turbine blade ring has a blade ring body and a plurality of protrusions that project from the blade ring body toward the upstream side of the axis on both sides in the axial direction in which the axis extends, and are spaced apart from each other and located at different positions in the circumferential direction, wherein two adjacent protrusions in the circumferential direction, included in the plurality of protrusions, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference side to the inner circumference side of the turbine blade ring, and the deflector is arranged to block at least a portion of the cooling medium intake port when viewed from the radial direction about the axis. The deflector includes a plate portion having a main surface that extends in a direction intersecting the radial direction. . Other blade ring assemblies according to the present disclosure include a turbine blade ring extending circumferentially about an axis, a stator blade disposed on the inner circumference of the turbine blade ring, and a deflector disposed on the outer circumference of at least a portion of the turbine blade ring, wherein the turbine blade ring has a blade ring body and a plurality of protrusions projecting from the blade ring body toward the upstream side of the axis on both sides in the axial direction in which the axis extends, and being spaced apart from each other and located at different positions in the circumferential direction, wherein two adjacent protrusions included in the plurality of protrusions in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference to the inner circumference of the turbine blade ring, and the deflector is positioned to obstruct at least a portion of the cooling medium intake port when viewed from the radial direction about the axis, and each of the two protrusions has a connection portion that can be connected to a fixing component to which the tailpipe of the combustor is fixed. Other blade ring assemblies according to the present disclosure include a turbine blade ring extending in the circumferential direction about an axis, a stationary blade disposed on the inner circumference side of the turbine blade ring, and a deflector disposed on the outer circumference side of at least a portion of the turbine blade ring, wherein the turbine blade ring has a blade ring body and a plurality of protrusions projecting from the blade ring body toward the upstream side of the axis on both sides in the axial direction in which the axis extends, and being spaced apart from each other and located at different positions in the circumferential direction, wherein two adjacent protrusions included in the plurality of protrusions in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference side to the inner circumference side of the turbine blade ring, and the deflector is positioned to block at least a portion of the cooling medium intake port when viewed from the radial direction about the axis, and the proportion of the cooling medium intake port blocked by the deflector when viewed from the outside in the radial direction is higher than the proportion of the cooling medium intake port blocked by the deflector when viewed from the upstream side of the axis. Other blade ring assemblies according to the present disclosure include a turbine blade ring extending in the circumferential direction about an axis, a stationary blade disposed on the inner circumference side of the turbine blade ring, and a deflector disposed on the outer circumference side of at least a portion of the turbine blade ring, wherein the turbine blade ring has a blade ring body and a plurality of protrusions projecting from the blade ring body toward the upstream side of the axis on both sides in the axial direction in which the axis extends, and being spaced apart from each other and located at different positions in the circumferential direction, wherein two adjacent protrusions included in the plurality of protrusions in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference side to the inner circumference side of the turbine blade ring, the deflector is disposed such that it obstructs at least a portion of the cooling medium intake port when viewed from the radial direction about the axis, and a fixing portion for fixing the deflector to the turbine blade ring is disposed between the two protrusions in the circumferential direction. Other blade ring assemblies according to the present disclosure include a turbine blade ring extending in the circumferential direction about an axis, a stator blade disposed on the inner circumference side of the turbine blade ring, and a deflector disposed on the outer circumference side of at least a portion of the turbine blade ring, wherein the turbine blade ring has a blade ring body and a plurality of protrusions projecting from the blade ring body toward the upstream side of the axis on both sides in the axial direction in which the axis extends, and being spaced apart from each other and located at different positions in the circumferential direction, wherein two adjacent protrusions included in the plurality of protrusions in the circumferential direction together with the blade ring body form a cooling medium intake port that extends from the outer circumference side to the inner circumference side of the turbine blade ring, and the deflector is disposed such as to block at least a portion of the cooling medium intake port when viewed from the radial direction about the axis, and is provided on the outer circumference side of the turbine blade ring and is located at a position that overlaps with the cooling medium intake port when viewed from the radial direction, and further comprises a forced cooling component through which a cooling medium flows, and the deflector is disposed on the inner circumference side of the forced cooling component.

[0007] To solve the above problems, the gas turbine according to the present disclosure comprises a rotor rotatable about the axis, a casing covering the rotor from the outer circumference, a blade ring assembly supporting the casing and covering the rotor from the outer circumference, and a combustor that generates combustion gas by burning fuel and sends the combustion gas into the casing, wherein the stator blades included in the blade ring assembly are the first stage turbine stator blades in the axial direction. [Effects of the Invention]

[0008] The blade ring assembly of this disclosure, and a gas turbine equipped therewith, can suppress the inflow of foreign matter into the stator blades. [Brief explanation of the drawing]

[0009] [Figure 1] This is a schematic cross-sectional view showing the entirety of a gas turbine according to the first embodiment of this disclosure. [Figure 2]It is a cross-sectional view showing an enlarged part of a gas turbine according to a first embodiment of the present disclosure. [Figure 3] It is a cross-sectional view schematically showing a part of a vane ring assembly according to a first embodiment of the present disclosure. [Figure 4] It is a perspective view showing a part of a vane ring assembly according to a first embodiment of the present disclosure, disassembled. [Figure 5] It is a cross-sectional view taken along the line V-V of the vane ring assembly shown in FIG. 3. [Figure 6] It is a plan view showing a deflector according to a first embodiment of the present disclosure. [Figure 7] It is a front view showing a deflector according to a first embodiment of the present disclosure. [Figure 8] It is a cross-sectional view showing a shielding cover according to a first embodiment of the present disclosure. [Figure 9] It is a cross-sectional view showing the operation of a deflector according to a first embodiment of the present disclosure. [Figure 10] It is a perspective view showing a part of a vane ring assembly according to a second embodiment of the present disclosure. [Figure 11] It is a plan view showing a deflector according to a second embodiment of the present disclosure. [Figure 12] It is a cross-sectional view showing the operation of a deflector according to a second embodiment of the present disclosure. [Figure 13] It is a perspective view showing a deflector according to other embodiments of the present disclosure.

Mode for Carrying Out the Invention

[0010] <First Embodiment> Hereinafter, a vane ring assembly according to a first embodiment of the present disclosure and a gas turbine including the same will be described with reference to the drawings. In the following description, the same reference numerals are given to configurations having the same or similar functions. And duplicate descriptions of those configurations may be omitted.

[0011] (Configuration of Gas Turbine) FIG. 1 is a cross-sectional view schematically showing the entirety of a gas turbine 10 in the present embodiment. The gas turbine 10 includes a compressor 20 that compresses air A, a combustor 30 that burns fuel F in the compressed air A to generate combustion gas G, and a turbine 40 that is driven by the combustion gas G.

[0012] The compressor 20 has a compressor rotor 21 that rotates about an axis Ar, a compressor casing 25 that covers the outer peripheral side of the compressor rotor 21, and a plurality of stationary blade stages 26. The turbine 40 has a turbine rotor 41 that rotates about the axis Ar, a turbine casing 45 that covers the outer peripheral side of the turbine rotor 41, and a plurality of stationary blade stages 46.

[0013] The compressor rotor 21 and the turbine rotor 41 are located on the same axis Ar and are connected to each other to form a gas turbine rotor 11. For example, a rotor of a generator GEN is connected to the gas turbine rotor 11. The compressor casing 25 and the turbine casing 45 are connected to each other to form a gas turbine casing 15. The gas turbine casing 15 is an example of a "casing". The combustor 30 is, for example, a can-type combustor.

[0014] Here, in the following description, the direction in which the axis Ar extends is defined as the "axial direction Da", the circumferential direction about the axis Ar is defined as the "circumferential direction Dc", and the direction perpendicular to the axis Ar about the axis Ar is defined as the "radial direction Dr". Of both sides in the axial direction Da, the side of the compressor 20 with reference to the turbine 40 is defined as the "axial upstream side Dau", and the opposite side is defined as the "axial downstream side Dad". Also, of both sides in the radial direction Dr, the side approaching the axis Ar is defined as the "radial inner side Dri", and the opposite side is defined as the "radial outer side Dro". Also, of both sides in the circumferential direction Dc, the side from the positive pressure surface to the negative pressure surface of a blade body 110 of a stationary blade 46a described later is defined as the "circumferential direction Dcn", and the side from the negative pressure surface to the positive pressure surface of the blade body 110 is defined as the "circumferential direction Dcp". The axial direction Da is the flow direction of the combustion gas G.

[0015] The compressor rotor 21 has a rotor shaft 22 extending in the axial direction Da with respect to the axis Ar, and a plurality of rotor blade stages 23 attached to the rotor shaft 22. The plurality of rotor blade stages 23 are arranged at intervals in the axial direction Da. Each rotor blade stage 23 is composed of a plurality of rotor blades 23a arranged in the circumferential direction Dc. A stator blade stage 26 is positioned on the downstream side Da of each of the plurality of rotor blade stages 23. Each stator blade stage 26 is mounted inside the compressor casing 25. Each stator blade stage 26 is composed of a plurality of stator blades 26a arranged in the circumferential direction Dc.

[0016] The turbine rotor 41 has a rotor shaft 42 extending in the axial direction Da with respect to the axis Ar, and a plurality of rotor blade stages 43 attached to the rotor shaft 42. The plurality of rotor blade stages 43 are arranged in the axial direction Da. Each rotor blade stage 43 is composed of a plurality of rotor blades 43a arranged in the circumferential direction Dc. A stator blade stage 46 is positioned on the upstream side Dau of each of the plurality of rotor blade stages 43. Each stator blade stage 46 is mounted inside the turbine casing 45. Each stator blade stage 46 is composed of a plurality of gas turbine stator blades 46a arranged in the circumferential direction Dc. For convenience of explanation, the gas turbine stator blades 46a will be simply referred to as "stator blades 46a" below.

[0017] Figure 2 is an enlarged cross-sectional view of a portion of the gas turbine 10 of the embodiment. For the sake of explanation, Figure 2 schematically shows a cross-section along the line II-II of the blade ring assembly WS shown in Figure 3. The turbine casing 45 has a cylindrical outer casing 45a that constitutes the outer shell of the turbine casing 45, an inner casing 45b fixed inside the outer casing 45a, and a plurality of segmented rings 45c fixed inside the inner casing 45b. Each of the plurality of segmented rings 45c is positioned between two adjacent stator stages 46 in a plurality of stator stages 46. A rotor stage 43 is positioned radially inward Dri of each segmented ring 45c.

[0018] The space between the rotor shaft 42 and the turbine casing 45 in the radial direction Dr, where the stationary blades 46a and rotor blades 43a are located, is a passage for combustion gas G from the combustor 30. Hereinafter, the above passage through which the combustion gas G from the combustor 30 flows will be referred to as the "combustion gas passage 49". The combustion gas passage 49 is formed as an annular space centered on the axis Ar and is elongated in the axial direction Da.

[0019] The gas turbine 10 in this embodiment is equipped with a cooling system 50 that supplies cooling air to the stator blades 46a and segmented rings 45c from the second stage onward, counting from the upstream side Dau of the axis. The cooling system 50 includes a foreign matter collector 51, a cooler 52, a boost compressor 53, and a cooling air line 54.

[0020] The foreign matter collector 51 is, for example, a strainer having multiple pores, which separates foreign matter 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 boost compressor 53 pressurizes the cooling air flowing through the cooling air line 54. The cooling air line 54 extracts compressed air Ac from the gas turbine casing 15 as cooling air, and supplies the extracted cooling air to the inside of the outer casing 45a via the foreign matter collector 51, the cooler 52, and the boost compressor 53.

[0021] In the inner casing 45b of the turbine casing 45, a cooling air passage 45p is formed that penetrates from the radially outer Dro to the radially inner Dri. Cooling air supplied from the cooling air line 54 to the inside of the outer casing 45a is introduced into the stator vanes 46a and segmented rings 45c of the second stage and beyond via the cooling air passage 45p in the inner casing 45b, and is used to cool the stator vanes 46a and segmented rings 45c. Here, because a foreign matter collector 51 is placed in the cooling air line 54, it is difficult for foreign matter to reach the stator vanes 46a and segmented rings 45c of the second stage and beyond. Note that the path for supplying cooling air to the stator vanes 46a is not limited to the above.

[0022] (Gas turbine operation) Returning to Figure 1, let's explain the operation of the gas turbine 10. The compressor 20 compresses air A to produce compressed air Ac. A portion (most) of the compressed air Ac produced by the compressor 20 flows into the combustor 30. Fuel F is supplied to the combustor 30. Inside the combustor 30, the fuel F burns in the compressed air Ac to produce high-temperature, high-pressure combustion gas G. The combustion gas G produced by the combustor 30 is sent from the combustor 30 to the combustion gas passage 49 in the turbine 40. As the combustion gas G flows through the combustion gas passage 49 toward the downstream side Da of the axis, it rotates the turbine rotor 41. As the turbine rotor 41 rotates, the rotor of the generator GEN connected to the gas turbine rotor 11 rotates. As a result, the generator GEN generates electricity.

[0023] (wing ring assembly) As shown in Figure 2, the gas turbine 10 includes a blade ring assembly WS. The blade ring assembly WS is located on the inner circumference side of the gas turbine casing 15. In this embodiment, the blade ring assembly WS is located at the upstream end of the turbine 40 in the axial direction Da (the most upstream axial direction Dau).

[0024] The blade ring assembly WS comprises a turbine blade ring 70, a plurality of stator blades 46a1 each forming a plurality of first-stage stator blades 46a in the axial direction Da, and a deflector 90. In this embodiment, compressed air Ac supplied from the compressor 20 into the gas turbine casing 15 is supplied directly to the first-stage stator blades 46a1 as cooling air without passing through the foreign matter collector 51. Therefore, in this embodiment, the turbine blade ring 70 is fitted with a deflector 90 to suppress foreign matter contained in the compressed air Ac from reaching the stator blades 46a1. This will be explained in detail below. For convenience of explanation, compressed air Ac may be referred to as "cooling air Ac" below.

[0025] (Construction around the wing ring assembly) The gas turbine casing 15 has a front wall 61, a circumferential wall 62, and a rear wall 63 as wall sections that define the housing chamber R in which the combustor 30 is housed. The housing chamber R is a space within the gas turbine casing 15 through which compressed air Ac flows, and the blade ring assembly WS is exposed.

[0026] The front wall 61 is located axially upstream Dau with respect to the containment chamber R. The front wall 61 has a cylindrical portion 61a with an opening 61h and a cylindrical lid portion 61b attached to the cylindrical portion 61a and covering the opening 61h. Part of the combustor 30 is located inside the cylindrical portion 61a and the cylindrical lid portion 61b. For example, the intake portion 31 of the combustor 30 is located inside the cylindrical lid portion 61b.

[0027] The circumferential wall 62 is located radially outward Dro with respect to the containment chamber R. The circumferential wall 62 extends between the front wall 61 and the rear wall 63, connecting the front wall 61 and the rear wall 63 in the axial direction Da. The circumferential wall 62 has a first circumferential wall 62a and a second circumferential wall 62b. The first circumferential wall 62a is the portion that connects to the front wall 61 from the axial downstream side Da. The first circumferential wall 62a extends in the axial direction Da. The second circumferential wall 62b is located axially downstream Da further than the first circumferential wall 62a, in an integral state with the first circumferential wall 62a, and is the portion that connects to the rear wall 63. The second circumferential wall 62b is, for example, an inclined portion (reduced diameter portion) that is positioned radially inward Dri as it proceeds toward the axial downstream side Da. The second peripheral wall 62b includes, for example, an arc portion whose inclination with respect to the axial direction Da becomes steeper as it proceeds downstream of the axis to Da.

[0028] The rear wall 63 is located downstream of the axial direction Dad relative to the containment chamber R. The rear wall 63 is connected to the second peripheral wall 62b from downstream of the axial direction Dad. The rear wall 63 extends radially along Dr. The rear wall 63 is a partition wall that closes the downstream of the axial direction Dad of the containment chamber R. The blade ring fixing portion 71 of the turbine blade ring 70, which will be described later, is fixed to the rear wall 63. The rear wall 63 is located downstream of the axial direction Dad of the turbine blade ring 70, which will be described later.

[0029] A guide section 64 is positioned at the inlet of the containment chamber R as viewed from the compressor 20, to guide the compressed air Ac into the containment chamber R. The guide section 64 is, for example, a turning vane positioned at an angle with respect to the axis Ar. The guide section 64 changes the direction of flow (flow direction) of the compressed air Ac flowing in from the compressor 20 toward the peripheral wall 62 of the gas turbine casing 15. However, if a portion of the compressed air Ac flowing in from the compressor 20 is directed toward the peripheral wall 62 or the rear wall 63, the guide section 64 may be omitted.

[0030] (Wing ring assembly structure) Figure 3 is a perspective view showing a part of the wing ring assembly WS. The blade ring assembly WS comprises a turbine blade ring 70, a plurality of combustor connecting members 80, a plurality of stator blades 46a1 (not shown in Figure 3, only one stator blade 46a1 is shown in Figure 5), a plurality of deflectors 90, a plurality of sealing members 85, and a plurality of shielding covers 150 (not shown in Figure 3, only one shielding cover 150 is shown in Figure 5).

[0031] (Turbine blade ring) The turbine blade ring 70 extends in the circumferential direction Dc centered on the axis Ar and is formed in an annular 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 to the housing chamber R of the gas turbine casing 15. The inner circumferential surface 70i is located on the opposite side of the outer circumferential surface 70o and faces radially inward Dri. The inner circumferential surface 70i faces a plurality of stator blades 46a1. The upstream end face 70u faces upstream Dau along the axis. The downstream end face 70d is located on the opposite side of the upstream end face 70u and faces downstream Dad along the axis.

[0032] The turbine blade ring 70 has a blade ring body 7, a blade ring fixing portion 71, and a plurality of protrusions 73. The blade ring body 7 is formed, for example, in a cylindrical shape. The blade ring fixing portion 71 fixes the blade ring body 7 to the gas turbine casing 15. The blade ring fixing portion 71 is integrally positioned with the blade ring body 7 at the end of the axial downstream Dad of the blade ring body 7. The blade ring fixing portion 71 is, for example, a flange that protrudes radially outward Dro from the blade ring body 7. The blade ring fixing portion 71 is located radially inward Dri of the rear wall 63 of the gas turbine casing 15 (see Figure 2). The blade ring fixing portion 71 is supported while fixed to the rear wall 63 of the gas turbine casing 15.

[0033] As shown in Figure 3, the turbine blade ring 70 has a plurality of air intake ports 72. The plurality of air intake ports 72 are arranged around the entire circumference of the turbine blade ring 70 at predetermined intervals in the circumferential direction Dc. The air intake ports 72 penetrate from the outer circumferential surface 70o to the inner circumferential surface 70i of the turbine blade ring 70. The air intake ports 72 are located upstream of the blade ring fixing portion 71 along the axis Dau.

[0034] Figure 4 is a perspective view showing a partially disassembled portion of the blade ring assembly WS. In this embodiment, the air intake 72 is a notch (groove) formed at the end of the turbine blade ring 70 on the upstream side Dau of the axis. That is, the air intake 72 is open to the upstream side Dau of the axis, penetrating from the outer circumferential surface 70o to the inner circumferential surface 70i of the turbine blade ring 70. Alternatively, the air intake 72 may be, for example, a through hole penetrating from the outer circumferential surface 70o to the inner circumferential surface 70i in the central part of the turbine blade ring 70 in the axial direction Da. The air intake 72 guides a portion of the cooling air Ac flowing through the housing chamber R of the gas turbine casing 15 from the outer circumferential side to the inner circumferential side of the turbine blade ring 70. The air intake 72 is an example of a "cooling medium intake," and the compressed air Ac compressed by the compressor 20 is an example of a "cooling medium."

[0035] In this embodiment, a plurality of protrusions 73 are formed by the air intake 72, which is created by cutting out the end of the turbine blade ring 70 on the upstream side Dau of the axis, thereby projecting from the blade ring body 7 toward the upstream side Dau of the axis. The plurality of protrusions 73 are spaced apart from each other and are positioned at different locations in the circumferential direction Dc, and each protrudes toward the upstream side Dau of the axis from the end face of the blade ring body 7 that faces the upstream side Dau of the axis. Hereinafter, the end face of the blade ring body 7 that faces the upstream side Dau of the axis will be referred to as the "upstream surface 7s". The end faces of the plurality of protrusions 73 on the upstream side Dau of the axis constitute the upstream end face 70u of the turbine blade ring 70 described above. The plurality of protrusions 73 and the plurality of air intakes 72 are arranged alternately in the circumferential direction Dc (see Figure 3). Therefore, as shown in Figure 4, when focusing on one air intake port 72, the multiple protrusions 73 have a first protrusion 73a located on the circumferential Dcn side of the air intake port 72 and a second protrusion 73b located on the circumferential Dcp side of the air intake port 72. The first protrusion 73a and the second protrusion 73b, together with the blade ring body 7, form an air intake port 72 that extends from the outer circumference to the inner circumference of the turbine blade ring 70. In other words, the air intake port 72 is located between the first protrusion 73a and the second protrusion 73b in the circumferential Dc direction. In this embodiment, the air intake port 72 is formed to narrow as it moves towards the radially inward Dri (see Figure 7). Hereinafter, for the sake of convenience in explanation, the narrowest part of the air intake port 72 located at the radially inward Dri may be referred to as the "narrow portion 72a". The narrow portion 72a is formed, for example, between the end of the radially inner Dri of the first projection 73a and the end of the radially inner Dri of the second projection 73b. The narrow portion 72a is an example of a "minimum portion".

[0036] As shown in Figure 5, the wing ring body 7 in this embodiment has a deflector fixing portion 75 that protrudes from the upstream surface 7s between the first protrusion 73a and the second protrusion 73b. The deflector fixing portion 75 is positioned, for example, in the center between the first protrusion 73a and the second protrusion 73b (see Figure 4). The deflector fixing portion 75 is positioned within the air intake 72 with a gap in the circumferential direction Dc between the first protrusion 73a and the second protrusion 73b (see Figure 4). As shown in Figure 5, the deflector fixing portion 75 has a support surface 75b facing radially outward Dro and a fixing surface 75a that is perpendicular to the support surface 75b and faces axially upstream Dau. The deflector 90, which will be described later, is attached to these support surface 75b and fixing surface 75a. The fixing surface 75a of the deflector fixing portion 75 has a coupling hole 75h into which a coupling device 93 such as a bolt can be coupled. Multiple (for example, two) connecting holes 75h are arranged on the fixed surface 75a.

[0037] As shown in Figure 3, a cooling medium flow section 100 is provided in adjacent first protrusions 73a and second protrusions 73b. The cooling medium flow section 100 connects the first protrusions 73a and second protrusions 73b so as to span across them. The cooling medium flow section 100 circulates the cooling medium through a cooling channel (not shown) formed in the blade ring body 7. In this disclosure, steam, air, etc., are used as the cooling medium, but the cooling medium is not limited to these, and any medium that has fluidity and can be used to cool the blade ring body 7 is appropriately adopted. The cooling medium flow section 100 is provided on the radially outer side Dro of the turbine blade ring 70 and overlaps with the air intake 72 when viewed in the radial direction Dr. The cooling medium flow section 100 is an example of a "forced cooling component".

[0038] (Combustion unit connection component) Next, the combustor connecting member 80 will be described. The combustor connecting member 80 is a fixing component to which the tailpipe 32 of the combustor 30 is fixed. As shown in Figure 3, the multiple combustor connecting members 80 are provided at positions corresponding to the multiple protrusions 73 of the turbine blade ring 70 in the circumferential direction Dc. As shown in Figures 3 and 4, the combustor connecting member 80 has a frame portion 81 and a flange 82.

[0039] The frame portion 81 is the part to which the tailpipe 32 of the combustor 30 is connected. The tailpipe 32 of the combustor 30 is fixed to the frame portion 81, for example, by welding. The flange 82 is the portion that extends radially outward from the frame portion 81 in Dro. 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 on the axial upstream side Dau of the projection 73 in the axial direction Da. A coupling hole 73h is formed in the end face 73u on the axial upstream side Dau of the projection 73, to which a coupling device 83 such as a bolt can be coupled.

[0040] The flange 82 has a through hole 82h through which a connector 83 is passed. The connector 83, passed through the through hole 82h of the flange 82, is coupled to the coupling hole 73h of the projection 73, thereby fixing the flange 82 to the projection 73. The end face 73u of the projection 73, including the coupling hole 73h, is an example of a "connection part" that can be connected to the combustor connection member 80. In this embodiment, the end faces 73u of the first projection 73a and the second projection 73b described above each have coupling holes 73h, and the combustor connection member 80 is attached to them.

[0041] (Silent Wing) Next, I will explain the stationary wing 46a1. Figure 5 is a cross-sectional view of the blade ring assembly WS shown in Figure 3 along the VV line. Multiple stator blades 46a1 (only one is shown in Figure 5) are arranged on the inner circumference side of the turbine blade ring 70 and aligned in the circumferential direction Dc. The stator blades 46a1 are held by retaining members 74 provided on the turbine blade ring 70. Each stator blade 46a1 has a blade body 110, an outer shroud 120, an inner shroud 130, and an air passage 140.

[0042] The airfoil body 110 has an airfoil cross-section and extends in the radial direction Dr. In other words, the height direction of the airfoil body 110 is in the radial direction Dr. The airfoil body 110 is positioned within the combustion gas passage 49 through which the combustion gas G passes. On the surface of the airfoil body 110, among the surfaces facing the circumferential direction Dc, the convex surfaces form the dorsal surfaces (negative pressure surfaces), and the concave surfaces form the ventral surfaces (positive pressure surfaces). Multiple exhaust holes 110h are formed at the ends of the airfoil body 110 on the upstream side Dau and the downstream side Dad.

[0043] The outer shroud 120 is provided at the radially outer end of the Dro of the wing body 110 and defines the outer peripheral position of the combustion gas flow path 49, which forms an annular space. The outer shroud 120 has a shroud body 121, a peripheral wall 122, a retainer 124, and an impact plate 123.

[0044] The shroud body 121 is formed in a plate shape that extends in the axial direction Da and the circumferential direction Dc. The shroud body 121 has a gas path surface 121a and an outer inner surface 121b. The gas path surface 121a is the surface that contacts the combustion gas G (the surface that faces the combustion gas flow path 49) and faces radially inward Dri. The outer inner surface 121b is the surface that faces the opposite side from the gas path surface 121a.

[0045] The peripheral wall 122 protrudes radially outward from the shroud body 121 along the outer peripheral edge of the shroud body 121 in Dro. In this embodiment, the peripheral wall 122 is formed around the entire circumference of the outer peripheral edge of the shroud body 121. The peripheral wall 122 has a front wall portion 122a with a front wall facing the downstream side of the axis Dad, a rear wall portion 122b with a rear wall facing the upstream side of the axis Dau, a rear wall portion 122c with a back wall facing the circumferential direction Dcp, and a ventral wall portion (not shown) with a ventral wall facing the circumferential direction Dcn. Therefore, the outer shroud 120 has a first space S1 which is a space surrounded from four directions by the peripheral wall 122.

[0046] The retainer 124 is positioned radially outward Dro of the front wall portion 122a in the peripheral wall 122. The retainer 124 extends radially outward Dro from the front wall portion 122a integrally with the front wall portion 122a. The radially outward Dro end of the retainer 124 is fixed, for example, to the deflector fixing portion 75 of the airfoil ring body 7. The retainer 124 has a front surface 124a facing the upstream side Dau of the axis.

[0047] The impact plate 123 is positioned within the first space S1 of the outer shroud 120, dividing the first space S1 into a region radially outward (Dro) and a cavity CA region radially inward (Dri). The impact plate 123 has a plurality of air holes 123h that penetrate radially (Dr). A portion of the cooling air Ac present in the radially outward (Dro) of the stator vane 46a1 flows into the cavity CA through the plurality of air holes 123h formed in the impact plate 123. After cooling the outer shroud 120, a portion of the air that has flowed into the cavity CA is exhausted into the combustion gas passage 49, for example, through an exhaust hole (not shown) provided in the outer shroud 120.

[0048] The inner shroud 130 is provided at the radially inner Dri end of the wing body 110 and defines the inner circumferential position of the combustion gas passage 49, which forms an annular space. The inner shroud 130 has a shroud body 131, a peripheral wall 132, and an impact plate 133.

[0049] The shroud body 131 is formed in a plate shape that extends in the axial direction Da and the circumferential direction Dc. The shroud body 131 has a gas path surface 131a and an inner surface 131b. The gas path surface 131a is the surface that contacts the combustion gas G (the surface that faces the combustion gas passage 49) and faces radially outward Dro. The inner surface 131b is the surface that faces the opposite side from the gas path surface 131a. The shroud body 131 has an exhaust hole 131h that connects the second space S2, which will be described later, to the combustion gas passage 49.

[0050] The peripheral wall 132 protrudes radially inward Dri from the shroud body 131 along the outer peripheral edge of the shroud body 131. In this embodiment, the peripheral wall 132 is formed around the entire circumference of the outer peripheral edge of the shroud body 131. The peripheral wall 132 has a front wall portion 133a with a front wall facing the downstream side Dad of the axis, a rear wall portion 133b with a rear wall facing the upstream side Dau of the axis, a rear wall portion 133c with a dorsal wall facing the circumferential direction Dcp, and a ventral wall portion (not shown) with a ventral wall facing the circumferential direction Dcn. The inner shroud 130 has a second space S2 which is a space surrounded from four directions by the peripheral wall 132.

[0051] The impact plate 133 is provided within the second space S2 of the inner shroud 130, and divides the second space S2 into a cavity CA, which is the radially inner Dri region and the radially outer Dro region. The impact plate 133 has a plurality of air holes 133h that penetrate in the radial Dr direction. The impact plate 133 may be omitted.

[0052] Multiple air passages 140 extend from the outer shroud 120 through the wing body 110 to the inner shroud 130. Some of the multiple air passages 140 may communicate with each other in the radially outer Dro portion or the radially inner Dri portion. One of the multiple air passages 140 communicates with the first space S1 of the outer shroud 120. One of the multiple air passages 140 communicates with the second space S2 of the inner shroud 130. The air passages 140 communicate with multiple exhaust holes 110h of the wing body 110.

[0053] A portion of the cooling air Ac that flows into the inner circumference of the turbine blade ring 70 through the air intake port 72 of the turbine blade ring 70 flows into the air holes 123h of the impact plate 123 of the outer shroud 120 and cools the outer shroud 120 by flowing through the cavity CA of the outer shroud 120. Another portion of the cooling air Ac that flows into the inner circumference of the turbine blade ring 70 flows into the air passage 140 and cools the blade body 110 as it passes through the air passage 140. A portion of the cooling air Ac that has flowed through the air passage 140 is exhausted into the combustion gas passage 49 from a plurality of exhaust holes 110h provided in the blade body 110. Another portion of the cooling air Ac that has flowed through the air passage 140 flows into the second space S2 of the inner shroud 130 and cools the inner shroud 130. The cooling air Ac that has flowed through the second space S2 of the inner shroud 130 is exhausted into the combustion gas passage 49 through the exhaust port 131h of the inner shroud 130.

[0054] (Deflector) Next, I will explain the deflector 90. As shown in Figure 3, the multiple deflectors 90 are arranged on the outer circumference of at least a portion of the turbine blade ring 70. The multiple deflectors 90 are arranged in a circumferential direction Dc and are positioned corresponding to the multiple air intakes 72. The deflectors 90 and the turbine blade ring 70 are made of a material of the same composition (e.g., stainless steel) and have the same coefficient of thermal expansion. "Same composition" here means, for example, that the material used to manufacture the turbine blade ring 70 and the material used to manufacture the deflectors 90 are the same material.

[0055] Figure 6 is a plan view of the deflector 90 as seen from the radially outer Dro. Figure 7 is a front view of the deflector 90 as seen from the axially upstream Dau. As shown in Figure 6, the deflector 90 is positioned to block at least a portion of the air intake 72 when viewed from the radially outer Dro. As shown in Figures 3 to 7, the deflector 90 has a fixed base portion 91 and a plate portion 92.

[0056] (Fixed base) The fixed base portion 91 is positioned between two adjacent protrusions 73 (first protrusion 73a and second protrusion 73b) in the circumferential direction Dc and is attached to the turbine blade ring 70. In this embodiment, the fixed base portion 91 is detachably fixed to the deflector fixing portion 75 of the blade ring body 7. The fixed base portion 91 has a first portion 91a that abuts the fixing surface 75a of the deflector fixing portion 75 from the axial upstream side Dau, and a second portion 91b that abuts the support surface 75b of the deflector fixing portion 75 from the radially outward side Dro. Each of the first portion 91a and the second portion 91b is plate-shaped with a predetermined thickness and is integrally formed with its ends connected to each other. The fixed base portion 91 is L-shaped when viewed from the circumferential direction Dc.

[0057] The first portion 91a of the fixing base 91 has a through hole 93h through which a connector 93 is passed. The connector 93, passed through the through hole 93h of the first portion 91a, is coupled to the coupling hole 75h of the deflector fixing portion 75, thereby fixing the first portion 91a and the second portion 91b, which is integrated with the first portion 91a, to the deflector fixing portion 75. In other words, the fixing base 91 is detachable from the deflector fixing portion 75 by the connector 93.

[0058] Furthermore, the first portion 91a is provided with a coupling hole 91h through which a coupling device 95, such as a bolt, is passed. The first portion 91a, including the coupling hole 91h, is an example of a "fixing portion" capable of fixing the sealing member 85. In this embodiment, the first portion 91a described above has a coupling hole 91h and to which the sealing member 85 is attached. Therefore, the first portion 91a, including the coupling hole 91h, for fixing the deflector 90 to the turbine blade ring 70, is positioned between two protrusions 73 in the circumferential direction Dc.

[0059] Furthermore, the second portion 91b is provided with connecting holes 94h on the radially outward-facing Dro side, to which a connector 94 such as a bolt can be attached. Multiple connecting holes 94h (for example, two) are arranged on the end face of the second portion 91b facing radially outward-facing Dro.

[0060] (Itabe) The plate portion 92 is attached to the fixed base portion 91. The plate portion 92 is attached to the radially outer Dro end of the fixed base portion 91 and is supported by the fixed base portion 91. The plate portion 92 is, for example, plate-shaped along a direction intersecting (e.g., perpendicular to) the radial direction Dr. The plate portion 92 has, for example, a main surface 92a that extends in a direction intersecting (e.g., perpendicular to) the radial direction Dr. In this embodiment, when the plate portion 92 is attached to the fixed base portion 91, it has a pair of main surfaces 92a positioned to face the radially outer Dro and the radially inner Dri. Of the pair of main surfaces 92a of the plate portion 92, the main surface 92a facing the radially inner Dri is in contact with the radially outer Dro facing surface of the second portion 91b of the fixed base portion 91.

[0061] The plate portion 92 has a through hole 92h through which a connector 94 is passed. The through hole 92h penetrates a pair of main surfaces 92a in the thickness direction of the plate portion 92. The connector 94, passed through the through hole 92h of the plate portion 92, is coupled to the coupling hole 94h of the second portion 91b, thereby fixing the plate portion 92 to the second portion 91b of the fixed base portion 91. In other words, the plate portion 92 is detachable from the fixed base portion 91 by the connector 94. Therefore, the fixed base portion 91 can be removed from the turbine blade ring 70 with the plate portion 92 and the fixed base portion 91 as a single unit.

[0062] As shown in Figure 6, the plate portion 92 is positioned so as to overlap at least a portion of each of the two adjacent protrusions 73 (first protrusion 73a and second protrusion 73b) in the circumferential direction Dc when viewed from the radially outer side Dro. Furthermore, the plate portion 92 is positioned with a gap between it and the upstream surface 7s of the blade ring body 7 when viewed from the radially outer side Dro. Hereinafter, the gap that the plate portion 92 leaves between it and the upstream surface 7s of the blade ring body 7 when viewed from the radially outer side Dro will be referred to as "gap G1". Here, the plate portion 92 has a first end portion 92e1 which is the end of the downstream side Dad of the axis, a pair of second end portions 92e2 which are the ends of the first direction (left and right direction in Figures 6 and 7) where the two protrusions are adjacent, and a third end portion 92e3 which is the end of the upstream side Dau of the axis. That is, the first end portion 92e1 is positioned with a gap G1 between it and the upstream surface 7s of the turbine blade ring 70. The first end 92e1 and the third end 92e3 extend in the first direction. The second end 92e2 extends in the axial direction Da.

[0063] Furthermore, as shown in Figure 7, the plate portion 92 is positioned with a gap between it and the two protrusions 73 in the radial direction Dr. Hereinafter, the gap that the plate portion 92 has between it and the two protrusions 73 in the radially outward direction Dr will be referred to as "gap G2". That is, the second end portion 92e2 is positioned with a gap G2 between it and the two protrusions 73. Moreover, the width L1 of the plate portion 92 in the first direction where the two protrusions 73 are adjacent is larger than the width L2 of the narrow portion 72a of the air intake 72 in the first direction when viewed from the radial direction. Note that the width L1 of the plate portion 92 is the distance between the pair of second end portions 92e2 in the first direction.

[0064] Furthermore, as shown in Figure 6, the proportion of the air intake 72 obstructed by the deflector 90 when viewed from the radially outer side Dro is higher than the proportion of the air intake 72 obstructed by the deflector 90 when viewed from the axially upstream side Dau, as shown in Figure 7. In other words, the proportion of the air intake 72 obstructed by the fixed base portion 91 and plate portion 92 in the radial direction Dr when viewed from the radially outer side Dro is higher than the proportion of the air intake 72 obstructed by the fixed base portion 91 and plate portion 92 in the axial direction Da when viewed from the axially upstream side Dau.

[0065] Furthermore, as shown in Figure 5, the third end portion 92e3 of the plate portion 92 is located downstream of the axis Dad from the front surface 124a of the retainer 124 included in the stationary vane 46a1. Also, the third end portion 92e3 is located downstream of the axis Dad from the surface of the fixed base portion 91 facing upstream of the axis Dau. Furthermore, the third end portion 92e3 is located downstream of the axis Dad from the end face 73u of the projection portion 73 facing upstream of the axis Dau.

[0066] Furthermore, as shown in Figure 3, the deflector 90 is positioned on the inner circumference side of the cooling medium flow section 100. The plate section 92 is located in the radial direction Dr between the cooling medium flow section 100 and the air intake port 72.

[0067] (Sealing material) As shown in Figures 3 to 7, the sealing member 85 is positioned between the tailpipes 32 of two adjacent combustors 30 in the circumferential direction Dc, and airtightly seals the gap between the tailpipes 32 of the two combustors 30. For example, a spring seal is used for the sealing member 85. The sealing member 85 is positioned on the inner circumference side of the turbine blade ring 70 (see Figure 5). The sealing member 85 is positioned in a location that overlaps with the air intake 72 in the radial direction Dr (see Figures 5 and 6). The sealing member 85 is removable to the outer circumference side of the turbine blade ring 70 through the air intake 72 when the tailpipes 32 of the combustors 30 are detached from the combustor connecting member 80.

[0068] The sealing member 85 is provided with a through hole 95h through which a connector 95 is passed. The connector 95, passed through the through hole 95h of the sealing member 85, is coupled to the coupling hole 91h of the first portion 91a of the fixing base portion 91, thereby fixing the sealing member 85 to the first portion 91a. In other words, the sealing member 85 is detachable from the fixing base portion 91 by the connector 95.

[0069] (Shielding cover) Figure 8 is a cross-sectional view showing the shielding cover 150. The shielding cover 150 is attached to the inner shroud 130 and fixed to the inner shroud 130 by welding or the like. The shielding cover 150 covers at least a portion of the second space S2 of the inner shroud 130 from the radially inner Dri. In this embodiment, the shielding cover 150 is attached to the peripheral wall 132 of the inner shroud 130 and covers the entire second space S2 of the inner shroud 130. With the shielding cover 150 attached to the inner shroud 130, compressed air Ac is no longer supplied to the second space S2 of the inner shroud 130 from the radially inner Dri. Compressed air Ac that flows in from the air intake 72 of the turbine blade ring 70 and flows through the air passage 140 of the outer shroud 120 and blade body 110 is supplied to the second space S2 of the inner shroud 130.

[0070] (Effects / Actions) Next, the functions and effects exhibited by the configuration of this embodiment will be described. As shown in Figure 2, compressed air Ac compressed by the compressor 20 is supplied from the compressor 20 to the containment chamber R where the combustor 30 is located. A portion of the compressed air Ac supplied to the containment chamber R has its flow direction angle changed by the guide section 64 and flows toward the peripheral wall 62 of the gas turbine casing 15. The compressed air Ac that collides with the peripheral wall 62 (first peripheral wall 62a) then splits into two, for example. That is, a portion of the compressed air Ac (for example, most of it) changes its flow direction toward the intake section 31 of the combustor 30 and then flows toward the intake section 31 of the combustor 30 (see arrow A1 in Figure 2). On the other hand, another portion of the compressed air Ac, after colliding with the peripheral wall 62 of the gas turbine casing 15, changes its flow direction so that it flows from the peripheral wall 62 toward the rear wall 63 (see arrow A2 in Figure 2). As a result, inside the containment chamber R, a flow of compressed air Ac is generated in the space located radially outward Dro relative to the turbine blade ring 70, moving from radially outward Dro to radially inward Dri (see arrow A2 in Figures 2 and 9). A portion of this compressed air Ac flow then attempts to be drawn into the inner circumference of the turbine blade ring 70 through the air intake port 72 of the turbine blade ring 70. Furthermore, as shown in Figures 2 and 4, a portion of the compressed air Ac supplied to the containment chamber R has its flow direction angle changed by the guide 64, and then a flow of compressed air (see arrow A3 in Figures 2 and 4) is generated that flows directly from the upstream side of the axis Dau towards the air intake 72. At the same time, it collides with the sealing member 85 that airtightly seals the gap between the two tailpipes 32 of the combustor 30, and a flow of compressed air Ac (see arrow A4 in Figures 2 and 4) is generated along the sealing member 85 from the radially inward Dri to the radially outward Dro.

[0071] Figure 9 is a cross-sectional view showing the operation of the deflector 90. In this embodiment, a deflector 90 is attached to the turbine blade ring 70. The deflector 90 includes a plate portion 92 positioned to obstruct at least a part of the air intake 72 when viewed from the radial direction Dr. Therefore, most of the compressed air Ac flowing along the rear wall 63 of the gas turbine casing 15 does not flow directly into the air intake 72, but instead collides with the plate portion 92 of the deflector 90 and bounces back (see arrows A2 and A2' in Figure 9). In this process, foreign matter M (e.g., rust and dirt) contained in the compressed air Ac flows upstream of the axis Dau so as to move away from the air intake 72. Therefore, it becomes difficult for foreign matter M contained in the compressed air Ac to enter the air intake 72 from the radially outer direction Dro. As a result, clogging of the cooling structure of the stator blade 46a1 (e.g., the exhaust holes 110h of the blade body 110 and the exhaust holes 131h of the inner shroud 130) with foreign matter M is suppressed.

[0072] Foreign matter M, which flows upstream of the axis Dau away from the air intake 72 as the compressed air Ac bounces back, is taken into the combustor 30, for example, from the intake section 31 of the combustor 30, and is discharged to the outside of the gas turbine 10 as part of the combustion gas G. The combustion gas passage 49 has fewer or no small holes that can be clogged by foreign matter compared to the stator vanes 46a1, so malfunctions caused by foreign matter M are less likely to occur.

[0073] Furthermore, according to the configuration described in this embodiment, the above-mentioned functions can be achieved with a simple configuration such as the plate portion 92. Therefore, for example, it is possible to suppress an increase in the manufacturing cost of the deflector 90.

[0074] Furthermore, as shown in Figure 7, the width L1 of the plate portion 92 in the first direction where the two protrusions 73 are adjacent is larger than the width L2 of the narrow portion 72a of the cooling medium intake port 72 in the first direction when viewed from the radial direction Dr. Therefore, compared to, for example, the case where the width L1 of the plate portion 92 is the same as or smaller than the width L2 of the narrow portion 72a, compressed air Ac is less likely to flow from the radially outer Dro through the air intake port 72 to the inner circumference side of the turbine blade ring 70.

[0075] Furthermore, as shown in Figure 6, the first end portion 92e1 of the plate portion 92 is positioned with a gap G1 between it and the upstream surface 7s of the blade ring body 7 when viewed from the radially outer Dro. In addition, as shown in Figure 7, the pair of second end portions 92e2 of the plate portion 92 are positioned with a gap G2 between them and the two protrusions 73 when viewed from the axially upstream Dau. Therefore, even if tolerances occur in the dimensions such as length, width, and thickness during the manufacturing process of the plate portion 92, the gaps G1 and G2 prevent interference between the plate portion 92 and the turbine blade ring 70. As a result, for example, the man-hours required for assembling the blade ring assembly WS can be reduced.

[0076] Furthermore, as shown in Figure 5, the third end portion 92e3 of the plate portion 92 is located downstream of the front surface 124a of the retainer 124 included in the stationary vane 46a1, at Da. Therefore, compared to, for example, the case where the third end portion 92e3 is at the same position as the front surface 124a of the retainer 124 in the axial direction Da, or located upstream of the front surface 124a at Dau, foreign matter M contained in the compressed air Ac flowing from the upstream side Dau (see arrow A3 in Figures 2, 4, and 9) is less likely to collide with the main surface 92a of the plate portion 92 facing the radially inward Dri and bounce back. As a result, foreign matter M is less likely to be taken into the air intake 72. Also, foreign matter M contained in the compressed air Ac flowing from the radially inward Dri (see arrow A4 in Figures 2, 4, and 9) is less likely to collide with the main surface 92a of the plate portion 92 facing the radially inward Dri and bounce back. As a result, foreign matter M is less likely to be drawn into the air intake 72.

[0077] Furthermore, according to the configuration described in this embodiment, the fixed base portion 91 of the deflector 90 is positioned between two protrusions 73 in the circumferential direction Dc and attached to the turbine blade ring 70, and the plate portion 92 is attached to the outer end of the fixed base portion 91 in the radial direction Dr. As a result, the fixed base portion 91 in the cooling medium intake port can support the plate portion 92 from the radially inner Dri. Therefore, when compressed air Ac collides with the plate portion 92 from the radially outer Dro, displacement of the plate portion 92 in the radially inner Dri can be suppressed.

[0078] Furthermore, according to the configuration described in this embodiment, the third end portion 92e3 of the plate portion 92 is located downstream of the axis Da to the surface of the fixed base portion 91 facing upstream of the axis Dau. Therefore, compared to the case where the third end portion 92e3 is at the same position in the axial direction Da as the surface of the fixed base portion 91 facing upstream of the axis Dau, or located upstream of the axis Dau to the surface, foreign matter M contained in the compressed air Ac flowing from radially outward Dro from the fixed base portion 91 is less likely to collide with and bounce off the main surface 92a of the plate portion 92 facing radially inward Dri. As a result, foreign matter M is less likely to be taken into the air intake port 72.

[0079] Furthermore, according to the configuration described in this embodiment, the fixed base portion 91 can be removed from the turbine blade ring 70 while the plate portion 92 and the fixed base portion 91 are integrated. Therefore, even if, for example, a forced cooling component such as the cooling medium flow portion 100 is present as an obstruction radially outward Dro from the plate portion 92, the fixed base portion 91 can be removed toward the upstream side Dau of the axis while remaining integrated with the plate portion 92. As a result, compared to the case where, for example, the fixed base portion 91 and the plate portion 92 are not integrated, maintenance such as repairs can be easily performed on the fixed base portion 91 and the plate portion 92.

[0080] Furthermore, according to the configuration described in this embodiment, the fixed base portion 91 has a first portion 91a to which a sealing member 85 is fixed, which airtightly seals the gap between two tailpipes 32 of adjacent combustors 30 in the circumferential direction Dc. In other words, the fixed base portion 91 is a member that supports the plate portion 92 from the radially inward Dri, and at the same time serves as a member for fixing the sealing member 85. Therefore, there is no need to provide any member other than the fixed base portion 91 to fix the sealing member 85. As a result, the increase in the number of parts of the turbine blade ring 70 can be suppressed.

[0081] Furthermore, according to the configuration described in this embodiment, each of the multiple protrusions 73 has an end face 73u that includes a connecting hole 73h that can be connected to a combustor connecting member 80 to which the tailpipe 32 of the combustor 30 is fixed. Therefore, by utilizing the protrusions 73, the length of the plate portion 92 can be made relatively long in the axial direction Da. Consequently, the inflow of foreign matter M contained in the compressed air Ac flowing from the radially outer side Dro toward the deflector 90 into the air intake port 72 is suppressed.

[0082] Furthermore, according to the configuration described in this embodiment, since the turbine blade ring 70 and the deflector 90 are formed from materials of the same composition, the effect of thermal expansion of the turbine blade ring 70 acting between the turbine blade ring 70 and the deflector 90 can be minimized. Therefore, for example, malfunctions between the turbine blade ring 70 and the first portion 91a of the fixed base 91 can be suppressed. As a result, the lifespan of the blade ring assembly WS can be extended.

[0083] Furthermore, as shown in Figures 6 and 7, the proportion of the air intake 72 viewed from the radially outer Dro that the deflector 90 obstructs in the radial direction Dr is higher than the proportion of the air intake 72 viewed from the axially upstream Dau that the deflector 90 obstructs. As a result, the amount of cooling air Ac taken into the air intake 72 from the axially upstream Dau is greater than the amount of compressed air Ac taken into the air intake 72 from the radially outer Dro. The inventors analyzed the size of foreign matter M contained in the compressed air Ac in the containment chamber R. As a result, they found that the size of foreign matter M contained in the compressed air Ac flowing from the radially outer Dro toward the deflector 90 is larger than the size of foreign matter M contained in the compressed air Ac flowing from the axially upstream Dau toward the deflector 90. Therefore, with the above configuration, foreign matter M contained in the compressed air Ac flowing from the radially outer Dro is suppressed from being taken into the air intake 72, while cooling air Ac is taken into the air intake 72 from the axially upstream side Dau, thereby suppressing the clogging of the cooling structure of the stator vane 46a1 by foreign matter M.

[0084] Furthermore, as shown in Figures 6 and 7, the plate portion 92 is positioned such that, when viewed from the radially outer Dro, it overlaps with at least a portion of each of the two adjacent protrusions 73 in the circumferential Dc direction, and a gap G2 is left between these two protrusions 73 in the radial Dr direction. A portion of the compressed air Ac flowing from the radially outer Dro toward the plate portion 92 collides with the main surface 92a of the plate portion 92 facing the radially outer Dro direction, flows in the circumferential Dc direction, then turns at an angle greater than 90° (for example, about 180°) and flows into the air intake 72 through the gap G2 in the radial Dr direction. In this process, foreign matter M contained in the compressed air Ac cannot turn due to the inertial force of the foreign matter M and is hardly taken into the air intake 72. In other words, while suppressing the intake of foreign matter M flying in from the radially outer Dr into the air intake 72, a portion of the compressed air Ac from which the foreign matter M has been separated by the plate portion 92 can be taken into the air intake 72 through the gap G2 between the plate portion 92 and the protruding portion 73. Therefore, compared to, for example, the case where the plate portion 92 abuts the protruding portion 73 without leaving a gap G2 in the radial Dr, the amount of compressed air Ac from which the foreign matter M has been separated that is taken into the air intake 72 can be increased.

[0085] <Second Embodiment> In the following description, a wing ring assembly according to the second embodiment of this disclosure will be described with reference to Figures 10 to 12. In the following description, components having the same or similar functions as those in the first embodiment will be denoted by the same reference numerals, and their descriptions will be omitted.

[0086] Figure 10 is a perspective view showing a partially disassembled part of the blade ring assembly WS, and Figure 11 is a plan view of the deflector 97 as seen from the radially outer Dro. In this embodiment, cooling holes 76 formed in the radial direction Dr are formed in the deflector fixing portion 75. The cooling holes 76 penetrate the deflector fixing portion 75, and one end of the radially outer Dro of the cooling holes 76 opens to the support surface 75b adjacent to the upstream surface 7s of the blade ring body 7, while the other end of the radially inner Dri opens to the first space S1 of the outer shroud 120. In this embodiment, three cooling holes 76 are formed in one deflector fixing portion 75 spaced apart from each other in the circumferential direction Dc, but the number of cooling holes 76 is not limited to three and can be appropriately selected according to the cooling performance required for the blade ring assembly WS.

[0087] The deflector 97 of this embodiment has a fixed base portion 98 and a plate portion 99. The fixed base portion 98 is detachably fixed to the deflector fixing portion 75 by a connector 93 and has a first portion 98a that abuts the fixing surface 75a of the deflector fixing portion 75 from the upstream side Dau of the axis, and a pair of second portions 98b, 98c provided on the upper surface of the radially outer side Dro of the first portion 98a. The first portion 98a and the second portions 98b, 98c are a single integrally molded part. The second portions 98b, 98c are spaced apart in the circumferential direction Dc and do not obstruct the flow of the cooling medium between them. The fixed base portion 98 is L-shaped when viewed from the circumferential direction Dc, but unlike the fixed base portion 91 of the first embodiment described above, the second portions 98b, 98c do not abut the support surface 75b of the deflector fixing portion 75. The second portions 98b and 98c are provided on the upper surface of the radially outer Dro of the first portion 98a, so that a groove 98d is formed between one second portion 98b and the other second portion 98c, with the upper surface of the first portion 98a as its bottom. The groove 98d is formed in the axial direction Da and is flush with the support surface 75b of the deflector fixing portion 75. The groove 98d and the support surface 75b define a cooling medium intake passage 97a between them and the plate portion 99 fixed to the second portions 98b and 98c.

[0088] As shown in Figure 11, the plate portion 99 is plate-shaped, extending in a direction intersecting the radial direction Dr, and has two main surfaces 99a facing radially outward Dro and radially inward Dri. One of the main surfaces 99a facing radially inward Dri abuts against the radially outward Dro-facing surfaces of the second portions 98b and 98c of the fixed base portion 98. The plate portion 99 has two through holes formed in each part, spaced apart in the circumferential direction Dc, through which the connectors 94 are passed. On the other hand, the radially outward Dro-facing surfaces of the second portions 98b and 98c each have two coupling holes formed, corresponding to the through holes in the plate portion 99. The connectors 94 passed through the through holes in the plate portion 99 are coupled to the coupling holes in the second portions 98b and 98c, thereby detachably fixing the plate portion 99 to the second portions 98b and 98c of the fixed base portion 98.

[0089] The plate portion 99 has a first end portion 99c, which is the end on the downstream side Da of the axis; two second ends 99d and 99e, which are the ends of the plate portion 99 in the longitudinal direction; a third end portion 99f, which is the end on the upstream side Dau of the axis; and a rectangular projection portion 99g that extends from the third end portion 99f toward the upstream side Dau of the axis. The first end portion 99c is positioned with a gap G1 between it and the upstream surface 7s of the wing ring body 7 when viewed from the radially outward side Dro. The first end portion 99c and the third end portion 99f extend in the first direction, the same as the plate portion 92 of the first embodiment. Both the second ends 99d and 99e extend in the axial direction Da, and they are parallel to each other. The leading edge 99h of the protruding portion 99g is parallel to the first end portion 99c, and the distance between the two side edges 99i and 99j of the protruding portion 99g, which are spaced apart in the first direction, is substantially equal to the length Dc in the circumferential direction of the fixed base portion 98.

[0090] The second end 99d of the plate portion 99 protrudes toward the circumferential direction Dcn, and the second end 99e protrudes toward the circumferential direction Dcp. When viewed from the radially outer direction Dro, the plate portion 99 is positioned such that the second end 99d overlaps the first protrusion 73a and the second end 99e overlaps the second protrusion 73b of two adjacent protrusions in the circumferential direction Dc. Furthermore, the second ends 99d and 99e are positioned with a gap between them and the protrusions 73a and 73b in the radial direction Dr.

[0091] When viewed from the radially outer Dro, the leading edge 99h of the protruding portion 99g coincides with the surface of the first portion 98a of the fixed base portion 98 on the axially upstream side Dau, and the two side edges 99i and 99j of the protruding portion 99g coincide with two sides of the fixed base portion 98 that are spaced apart in the circumferential direction Dc. As a result, the protruding portion 99g is positioned to overlap the groove 98d of the fixed base portion 98, and the groove 98d is obscured by the protruding portion 99g when viewed from the radially outer Dro.

[0092] Figure 12 is a cross-sectional view showing the operation of the deflector 97. Compressed air Ac (A3), which serves as a cooling medium, flows into the intake passage 97a of the deflector 97, with its flow direction angle changed by the guide section (compressor turning vane), from the upstream side Dau along the axis. The compressed air Ac that flows into the intake passage 97a flows between the second sections 98b and 98c, along the groove 98d and the support surface 75b towards the downstream side Dad along the axis, and collides with the upstream surface 7s of the blade ring body 7. A portion of the compressed air Ac that collides with the upstream surface 7s changes direction radially inward Dri and flows into the first space S1 of the outer shroud 120 through the cooling holes 76. The cooling air Ac that flows into the inner circumference side of the turbine blade ring 70 through the cooling holes 76 cools the blade body 110, the outer shroud 120, and the inner shroud 130, similar to the first embodiment described above. In this embodiment, since a cooling medium intake passage 97a is formed in the fixed base portion 98 of the deflector 97, the amount of compressed air Ac flowing into the outer shroud 120 and the inner shroud 130 can be increased compared to the deflector 90 of the first embodiment. This improves the cooling effect of the stator vane 46a1.

[0093] Separate from the compressed air Ac(A3), the compressed air Ac(A2) flowing along the rear wall 63 of the gas turbine casing 15 does not flow into the air intake 72, but instead collides with the plate portion 99 of the deflector 97. At this time, foreign matter M contained in the compressed air Ac hits the plate portion 99 and bounces back, flowing toward the upstream side Dau of the axis so as to move away from the air intake 72. In this embodiment, the plate portion 99 of the deflector 97 has an overhang portion 99g that extends toward the upstream side Dau of the axis and blocks the groove 98d, so the foreign matter M hits the plate portion 99 and bounces back, and does not hit the groove 98d of the fixed base portion 98. In other words, even if a groove 98d that constitutes part of the compressed air intake passage 97a is formed in the fixed base portion 98 of the deflector 97, the overhang portion 99g can prevent foreign matter M from flowing from the intake passage 97a into the outer shroud 120 and the inner shroud 130. As a result, the accumulation of foreign matter M in the cooling structure of the stationary vane 46a1 is suppressed.

[0094] (Other embodiments) Although embodiments of this disclosure have been described in detail above with reference to the drawings, the specific configuration is not limited to that of the embodiments, and additions, omissions, substitutions, and other modifications to the configuration are possible without departing from the gist of this disclosure.

[0095] As shown in Figure 13, the plate portion 92' of the deflector 90 may be integrally formed with the fixing base portion 91'. In this case, the fixing base portion 91' has through holes 91c through which a connector 96 such as a bolt passes. Multiple (for example, two) through holes 91c are formed in the fixing base portion 91'. The fixing base portion 91' is fixed to the blade ring body 7 by connecting it to a pre-formed coupling hole 7a in the blade ring body 7 with the connector 96 passed through the through holes 91c. This eliminates the need to attach the fixing base portion 91' and the plate portion 92' separately, making it easier to attach the deflector 90 to the turbine blade ring 70, for example.

[0096] Furthermore, the plate portion 92 described in the above embodiment is not limited to the arrangement described above, and may, for example, be placed inside the air intake 72.

[0097] <Note> The blade ring assembly described in the embodiment, and the gas turbine equipped therewith, can be understood, for example, as follows:

[0098] (1) The first aspect of the blade ring assembly WS comprises a turbine blade ring 70 extending in the circumferential direction Dc centered on the axis Ar, a stationary blade 46a disposed on the inner circumference side of the turbine blade ring 70, and deflectors 90, 97 disposed on the outer circumference side of at least a portion of the turbine blade ring 70. The turbine blade ring 70 has a blade ring body 7 and a plurality of protrusions 73 that protrude from the blade ring body 7 toward the upstream side Dau of the axial direction Da on both sides of the axial direction Da on which the axis Ar extends, and are spaced apart from each other and located at different positions in the circumferential direction Dc. Two adjacent protrusions 73 included in the plurality of protrusions 73 in the circumferential direction Dc, together with the blade ring body 7, form a cooling medium intake port (air intake port 72) that extends from the outer circumference side to the inner circumference side of the turbine blade ring 70. The deflectors 90, 97 are arranged to block at least a portion of the cooling medium intake port when viewed from the radial direction Dr centered on the axis Ar.

[0099] As a result, most of the compressed air Ac flowing from the outside radially Dr to the blade ring assembly WS within the gas turbine casing 15 does not flow directly into the coolant intake, but instead collides with the deflectors 90 and 97 and bounces back. In this process, foreign matter M (e.g., rust and debris) contained in the compressed air Ac flows upstream along the axis Dau, away from the coolant intake. Therefore, it becomes difficult for foreign matter M contained in the compressed air Ac to enter the coolant intake from the outside radially Dr.

[0100] (2) The airfoil ring assembly WS according to the second embodiment is the airfoil ring assembly WS according to the first embodiment, wherein the deflectors 90, 97 may include plate portions 92, 92', 99 having main surfaces 92a, 99a that extend in a direction intersecting the radial direction Dr.

[0101] This makes it possible to achieve the above-mentioned function (1) with a simple configuration such as plate sections 92, 92', and 99.

[0102] (3) The third aspect of the blade ring assembly WS is the same as the second aspect of the blade ring assembly WS, wherein the cooling medium intake port has a minimum portion (narrow portion 72a) between the inner circumferential ends of the two protrusions 73, and the width L1 of the plate portions 92, 92' in the first direction where the two protrusions 73 are adjacent may be greater than the width L2 of the minimum portion of the cooling medium intake port in the first direction.

[0103] As a result, compared to the case where the width L1 of the plate portions 92,92' is the same as or smaller than the width L2 of the initial portion of the cooling medium intake, compressed air Ac is less likely to flow from the outside in the radial direction Dr through the cooling medium intake to the inner circumference side of the turbine blade ring 70.

[0104] (4) The blade ring assembly WS according to the fourth embodiment is the blade ring assembly WS according to the second or third embodiment, wherein the plate portions 92, 92' have a first end portion 92e1 which is the downstream Da end of the axial direction Da, and a pair of second end portions 92e2 which are the ends of the two protrusions 73 adjacent to each other in the first direction, and the first end portion 92e1 and the pair of second end portions 92e2 may be arranged with gaps G1, G2 between them and the turbine blade ring 70.

[0105] As a result, because of the gaps G1 and G2, even if tolerances occur in the dimensions such as length, width, and thickness during the manufacturing process of the plate portions 92 and 92', interference between the plate portions 92 and 92' and the turbine blade ring 70 can be suppressed. Consequently, the man-hours required for tasks such as assembling the blade ring assembly WS can be reduced.

[0106] (5) The fifth aspect of the wing ring assembly WS is one of the second to fourth aspects of the wing ring assembly WS, wherein the plate portions 92, 92' have a third end portion 92e3 which is the end of the upstream side Dau of the axis, and the third end portion 92e3 may be located on the downstream side Da of the axis, on either side of Da in the axial direction, rather than on the surface (front surface 124a) of the retainer 124 included in the stationary blade 46a that faces the upstream side Dau of the axis.

[0107] As a result, compared to the case where the third end portion 92e3 is at the same position in the axial direction Da as the surface of the retainer 124 facing the upstream side Dau of the axis, or is located further upstream in the axial direction Dau than that surface, foreign matter M contained in compressed air Ac flowing from the upstream side Dau of the axis is less likely to collide with and bounce off the main surface 92a of the plate portion 92, 92' facing inward in the radial direction Dr. Also, foreign matter M contained in compressed air Ac flowing from outside the radial direction Dr is less likely to collide with and bounce off the main surface 92a of the plate portion 92 facing inward in the radial direction Dr.

[0108] (6) The blade ring assembly WS according to the sixth embodiment is a blade ring assembly according to any one of the second to fifth embodiments, wherein the deflector 90 further includes fixed base portions 91, 91', 98 positioned between the two protrusions 73 in the circumferential direction Dc and attached to the turbine blade ring 70, and the plate portions 92, 92', 99 may be attached to the outer ends of the fixed base portions 91, 91', 98 in the radial direction Dr.

[0109] This allows the fixed base portions 91, 91' inside the cooling medium intake port to support the plate portions 92, 92' from the inside in the radial direction Dr.

[0110] (7) The wing ring assembly WS according to the seventh embodiment is the wing ring assembly WS according to the sixth embodiment, wherein the plate portion 92 has a third end portion 92e3 which is the end portion of the upstream side Dau of the axis, and the third end portion 92e3 may be located on the downstream side Da of the axis, on either side of Da in the axial direction, rather than on the surface of the fixed base portion 91 facing the upstream side Dau of the axis.

[0111] As a result, compared to the case where the third end portion 92e3 is at the same position in the axial direction Da as the surface of the fixed base portion 91 facing the axial upstream side Dau, or is located further upstream in the axial direction Dau than the said surface, foreign matter M contained in the compressed air Ac flowing from the inside of the fixed base portion 91 in the radial direction Dr is less likely to collide with and bounce off the main surface 92a of the plate portion 92 facing the inside in the radial direction Dr.

[0112] (8) The blade ring assembly WS according to the eighth embodiment is the blade ring assembly WS according to the sixth or seventh embodiment, wherein the fixing base portion 91 may be removable from the turbine blade ring 70 while the plate portion 92 and the fixing base portion 91 are integrated.

[0113] This makes it easier to perform maintenance, such as repairs, on the fixed base 91 and the plate 92, compared to, for example, a case where the fixed base 91 and the plate 92 are not integrated.

[0114] (9) The wing ring assembly WS according to the ninth embodiment is any one of the wing ring assembly WS from the sixth to the eighth embodiment, wherein the fixing base portion 91 may have a fixing portion (first portion 91a including a coupling hole 91h) capable of fixing a sealing member 85 that closes the gap between two tailpipes 32 of a combustor 30 adjacent in the circumferential direction Dc.

[0115] As a result, the fixing base portion 91 serves as a member that supports the plate portion 92 from the inside in the radial direction Dr, and at the same time, it also serves as a member for fixing the sealing member 85. Therefore, for example, there is no need to provide any member other than the fixing base portion 91 to fix the sealing member 85.

[0116] (10) The blade ring assembly WS according to the tenth embodiment is any one of the blade ring assembly WS from the sixth to ninth embodiments, wherein the fixed base portion 98 has a groove 98d formed in the axial direction Da on the radial side of the plate portion 99, and which defines a cooling medium intake passage 97a between it and the plate portion 99.

[0117] This increases the amount of cooling medium flowing into the cooling structure of the stator vanes, thereby improving the cooling effect of the stator vanes.

[0118] (11) The 11th aspect of the wing ring assembly WS is the 10th aspect of the wing ring assembly WS, wherein the plate portion 99 is formed to protrude on the upstream side Dau of the axis and may have a protruding portion 99g that overlaps with the groove 98d when viewed from the outside in the radial direction Dr.

[0119] As a result, even when a deflector 97 with a compressed air intake passage 97a is used, the protruding portion 99g can prevent foreign matter M from entering the cooling structure of the stationary vane 46a1.

[0120] (12) The wing ring assembly WS according to the tenth embodiment is any one of the wing ring assembly WS from the first to the eleventh embodiment, wherein each of the two protrusions 73 may have a connecting portion (end face 73u including a connecting hole 73h) that can be connected to a fixing component (combustion unit connecting member 80) to which the tail tube 32 of the combustor 30 is fixed.

[0121] This allows the lengths of the plate portions 92, 92', and 99 to be relatively long in the axial direction Da by utilizing the two protrusions 73.

[0122] (13) The blade ring assembly WS according to the 13th embodiment is any one of the blade ring assembly WS from the 1st to the 12th embodiments, wherein the turbine blade ring 70 and the deflector 90 may be formed of a material of the same composition.

[0123] This makes it possible to minimize the effect of thermal expansion of the turbine blade ring 70 acting between the turbine blade ring 70 and the deflector 90. Therefore, it is possible to suppress the occurrence of malfunctions between the turbine blade ring 70 and the deflector 90.

[0124] (14) The 14th aspect of the blade ring assembly WS is any one of the first to 13 aspects of the blade ring assembly WS, wherein the proportion of the cooling medium intake port that the deflector 90 obstructs when viewed from the outside in the radial direction Dr is higher than the proportion of the cooling medium intake port that the deflector 90 obstructs when viewed from the upstream side Dau of the axis.

[0125] As a result, the amount of cooling air Ac taken in from the upstream side of the axis Dau to the cooling medium intake becomes greater than the amount of compressed air Ac taken in from the outside of the radial direction Dr to the cooling medium intake.

[0126] (15) The blade ring assembly WS according to the 15th embodiment is any one of the blade ring assembly WS from the 1st to 14th embodiments, wherein the fixing portion (first portion 91a including the insertion hole 93h) for fixing the deflector 90 to the turbine blade ring 70 may be arranged between the two protrusions 73 in the circumferential direction Dc.

[0127] This allows for space savings compared to, for example, the case where the fixing portion of the deflector 90 is located on the outer or inner circumference side of the turbine blade ring 70.

[0128] (16) The blade ring assembly WS according to the 16th embodiment is a blade ring assembly WS according to any one of the first to 15 embodiments, further comprising a forced cooling component (cooling medium flow section 100) through which steam flows, provided on the outer circumference side of the turbine blade ring 70 and positioned to overlap with the cooling medium intake when viewed in the radial direction Dr, and the deflector 90 may be provided on the inner circumference side of the forced cooling component.

[0129] (17) The gas turbine 10 according to the 17th embodiment comprises a rotor (gas turbine rotor 11) that can rotate about the axis Ar, a casing (gas turbine casing 15) that covers the rotor from the outer circumference, one of the first to 16 embodiments of blade ring assembly WS that covers the rotor from the outer circumference while being supported by the casing, and a combustor 30 that generates combustion gas G by burning fuel F and sends the combustion gas G into the casing, wherein the stator blade 46a included in the blade ring assembly WS is the first stage turbine stator blade (stator blade 46a1) in the axial direction Da. [Explanation of symbols]

[0130] 7...Wing ring body 7a,73h,75h,91h,94h…joining hole 7s…Upstream side 10... Gas turbine 11…Gas turbine rotor 15…Gas turbine casing 20... Compressor 21... Compressor rotor 22,42…Rotor shaft 23,43...Rotor blade stage 23a,43a...moving blade 25...Compressor compartment 26,46…Stator vane stage 26a…Shizuyuki 30... Combustor 31... Intake section 32...tail tube 40... Turbine 41... Turbine rotor 45... Turbine casing 45a…Outer compartment 45b…Inner compartment 45c…Divided ring 45p...Cooling air passage 46a... Gas turbine stator blades 46a1...Silent Wing 49…Combustion gas flow path 50…Cooling device 51...Foreign object collector 52...Cooler 53... Boost compressor 54…Cooling air line 61...Front wall 61a...Cylinder part 61b...Cylindrical lid 61h…Aperture 62...peripheral wall 62a...First peripheral wall 62b…Second peripheral wall 63…Back wall 64…Information Department 70... Turbine blade ring 70o...outer surface 70i…Inner peripheral surface 70d…Downstream end face 70u…Upstream end face 71...Blade ring fixed part 72... Air intake 72a…Narrow area 73...Protrusion 73a...first protrusion 73b…Second protrusion 73u...end face 75... Deflector fixing part 75a…Fixed surface 75b…support surface 76…Cooling hole 80... Combustion unit connection component 81...Frame part 82…Flange 82h, 91c, 92h, 93h, 95h… Through hole 83, 93, 94, 95, 96… Connectors 85...Sealing material 90, 97… Deflector 91,91´,98…Fixed base part 91a…Part 1 91b…Second part 92,92´,99…Plate part 92a,99a…main surface 92e1...First end 92e2…Second end 92e3...Third end 98d…Groove 99g…Protrusion 100...Cooling medium distribution section 110...wing body 110h... Exhaust vent 120... Outer shroud 121,131... Shroud body 121a, 131a... Gas path surface 121b…Outer inner surface 122,132…peripheral wall 122a, 132a...Front wall part 122b, 132b...Rear wall part 122c, 132c...Dorsal wall part 123,133…Collision plate 123h... Air vent 124... Retainer 124a...Front 130...Inner shroud 131b...Inner surface 140... Air passage 150...Shielding cover A...Air A1, A2, A2', A3, A4... arrows Ac... Compressed air Ar…Axis line CA... Cavity Da…Axis direction Dad... downstream of the axis Dau…Axis upstream side Dc, Dcn, Dcp…Circumferential direction Dr…Radial direction Dri...Radial inner side Dro...Radial outward F…Fuel G... Combustion gas G1, G2... Gap GEN... Generator L1…Width of the plate in the first direction L2…Width of the narrow section in the first direction R... Containment room S1…first space S2…Second space WS... Wing Ring Assembly

Claims

1. A turbine blade ring extending in the circumferential direction centered on the axis, A stationary vane positioned on the inner circumference side of the turbine blade ring, The turbine blade ring comprises a deflector disposed on the outer circumference of at least a portion of the turbine blade ring, The turbine blade ring is The wing ring body and The blade ring body has a plurality of protrusions that extend from both sides in the axial direction in which the axis extends toward the upstream side of the axis, and are spaced apart from each other and positioned at different locations in the circumferential direction. Two of the aforementioned multiple protrusions, which are adjacent in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference to the inner circumference of the turbine blade ring. The deflector is positioned such that, when viewed from the radial direction centered on the axis, it obstructs at least a portion of the cooling medium intake port. The deflector is a blade ring assembly including a plate portion having a main surface that extends in a direction intersecting the radial direction.

2. The cooling medium intake port has the smallest portion where the cooling medium intake port is narrowest, between the inner circumferential ends of the two protrusions. The blade ring assembly according to claim 1, wherein the width of the plate portion in the first direction in which the two protrusions are adjacent to each other is greater than the width of the minimum portion of the cooling medium intake port in the first direction.

3. The plate portion has a first end which is the downstream end of the axial direction, and a pair of second ends which are the two adjacent ends in the first direction. The blade ring assembly according to claim 1, wherein the first end and the pair of second ends are arranged with a gap between them and the turbine blade ring.

4. The plate portion has a third end which is the end on the upstream side of the axis, The wing ring assembly according to claim 1, wherein the third end is located on the downstream side of the axis in the axial direction, more so than the surface of the retainer included in the stator vane that faces the upstream side of the axis.

5. The deflector further includes a fixed base portion positioned between the two protrusions in the circumferential direction and attached to the turbine blade ring, The blade ring assembly according to claim 1, wherein the plate portion is attached to the radially outer end of the fixed base portion.

6. The plate portion has a third end which is the end on the upstream side of the axis, The wing ring assembly according to claim 5, wherein the third end is located on the downstream side of the axis in the axial direction, more so than the surface of the fixed base portion facing the upstream side of the axis.

7. The blade ring assembly according to claim 5, wherein the fixing base portion is removable from the turbine blade ring while the plate portion and the fixing base portion are integrated.

8. The blade ring assembly according to claim 5, wherein the fixing base portion has a fixing portion capable of fixing a sealing member that closes the gap between two tailpipes of combustors adjacent to each other in the circumferential direction.

9. The blade ring assembly according to claim 5, wherein the fixed base portion has a groove formed in the axial direction radially inward from the plate portion, and defining a cooling medium intake passage between it and the plate portion.

10. The blade ring assembly according to claim 9, wherein the plate portion is formed to protrude toward the upstream side of the axis and has a protruding portion that overlaps the groove when viewed from the outside in the radial direction.

11. A turbine blade ring extending in the circumferential direction with respect to the axis, A stationary vane positioned on the inner circumference side of the turbine blade ring, The turbine blade ring comprises a deflector disposed on the outer circumference of at least a portion of the turbine blade ring, The turbine blade ring is The wing ring body and The blade ring body has a plurality of protrusions that extend from both sides in the axial direction in which the axis extends toward the upstream side of the axis, and are spaced apart from each other and positioned at different locations in the circumferential direction. Two of the aforementioned multiple protrusions, which are adjacent in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference to the inner circumference of the turbine blade ring. The deflector is positioned such that, when viewed from the radial direction centered on the axis, it obstructs at least a portion of the cooling medium intake port. Each of the two aforementioned protrusions is a blade ring assembly having a connecting portion that can be connected to a fixing component to which the tail section of the combustor is fixed.

12. A turbine blade ring extending in the circumferential direction with respect to the axis, A stationary vane positioned on the inner circumference side of the turbine blade ring, The turbine blade ring comprises a deflector disposed on the outer circumference of at least a portion of the turbine blade ring, The turbine blade ring is The wing ring body and The blade ring body has a plurality of protrusions that extend from both sides in the axial direction in which the axis extends toward the upstream side of the axis, and are spaced apart from each other and positioned at different locations in the circumferential direction. Two of the aforementioned multiple protrusions, which are adjacent in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference to the inner circumference of the turbine blade ring. The deflector is positioned such that, when viewed from the radial direction centered on the axis, it obstructs at least a portion of the cooling medium intake port. A blade ring assembly in which the deflector obstructs the cooling medium intake port when viewed from the radially outer side, is higher than the deflector obstructs the cooling medium intake port when viewed from the upstream side of the axis.

13. A turbine blade ring extending in the circumferential direction with respect to the axis, A stationary vane positioned on the inner circumference side of the turbine blade ring, The turbine blade ring comprises a deflector disposed on the outer circumference of at least a portion of the turbine blade ring, The turbine blade ring is The wing ring body and The blade ring body has a plurality of protrusions that extend from both sides in the axial direction in which the axis extends toward the upstream side of the axis, and are spaced apart from each other and positioned at different locations in the circumferential direction. Two of the aforementioned multiple protrusions, which are adjacent in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference to the inner circumference of the turbine blade ring. The deflector is positioned such that, when viewed from the radial direction centered on the axis, it obstructs at least a portion of the cooling medium intake port. The fixing portion for securing the deflector to the turbine blade ring is a blade ring assembly positioned between the two protrusions in the circumferential direction.

14. A turbine blade ring extending in the circumferential direction with respect to the axis, A stationary vane positioned on the inner circumference side of the turbine blade ring, The turbine blade ring comprises a deflector disposed on the outer circumference of at least a portion of the turbine blade ring, The turbine blade ring is The wing ring body and The blade ring body has a plurality of protrusions that extend from both sides in the axial direction in which the axis extends toward the upstream side of the axis, and are spaced apart from each other and positioned at different locations in the circumferential direction. Two of the aforementioned multiple protrusions, which are adjacent in the circumferential direction, together with the blade ring body, form a cooling medium intake port that extends from the outer circumference to the inner circumference of the turbine blade ring. The deflector is positioned such that, when viewed from the radial direction centered on the axis, it obstructs at least a portion of the cooling medium intake port. The turbine blade ring is provided on the outer circumference side and positioned to overlap with the cooling medium intake port when viewed in the radial direction, and further comprises a forced cooling component through which the cooling medium flows, The deflector is a blade ring assembly positioned on the inner circumference side of the forced cooling component.

15. A rotor that can rotate about the aforementioned axis, A casing that covers the rotor from the outer circumference, A blade ring assembly according to any one of claims 1 to 14, which covers the rotor from the outer circumference side while being supported by the casing, The system comprises a combustor that generates combustion gas by burning fuel and sends the combustion gas into the casing, A gas turbine in which the stator blades included in the blade ring assembly are first-stage turbine stator blades in the axial direction.

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