JP6914643B2 - Static wing segment, gas turbine and gas turbine equipment equipped with it - Google Patents

Description

The present invention relates to a stationary blade segment including a stationary blade, a gas turbine including the stationary blade, and a gas turbine equipment.

The gas turbine includes a compressor that compresses the atmosphere to generate compressed air, a combustor that burns fuel in the compressed air to generate combustion gas, and a turbine that is driven by the combustion gas. The turbine has a turbine rotor that rotates about an axis, a plurality of vane trains that are aligned in the axial direction in which the axis extends, and a turbine casing that rotatably covers the turbine rotor. The plurality of vane rows are arranged in the axial direction in which the axis extends. Each vane row has a plurality of vanes aligned in the circumferential direction with respect to the axis. The stationary blade has a blade body extending in the radial direction with respect to the axis, an inner shroud provided on the radial inner side of the blade body, and an outer shroud provided on the radial outer side of the blade body. The inner shroud of the vane defines the inner periphery of the combustion gas flow path through which the combustion gas flows. The outer shroud also defines the outer perimeter of the combustion gas flow path. The blade body is arranged in this combustion gas flow path.

Since the stationary blade is exposed to high-temperature combustion gas, it is necessary to cool it with, for example, cooling air.

For example, in the stationary blade of Patent Document 1 below, two blade passages are formed which penetrate from the outer shroud, through the blade body, and to the inner shroud. Both of the two wing passages have radial outer ends open. In addition, the two wing passages communicate with each other on the inner side in the radial direction. Of the two wing passages, cooling air flows into the first wing passage from the outside in the radial direction. This cooling air flows inward in the radial direction in the first wing passage and flows into the second wing passage. The cooling air that has flowed into the second wing passage flows outward in the radial direction and is discharged to the outside of the stationary wing through the opening of the second wing passage. The cooling air is heated by heat exchange with the stationary blade in the process of passing through the stationary blade. The cooling air discharged from the vane is used as a part of the combustion air in the combustor.

As described above, in the technique described in Patent Document 1, the cooling air heated by the cooling of the stationary blade is used as the combustion air.

Japanese Unexamined Patent Publication No. 2013-09348

In the technique described in Patent Document 1, the cooling air heated by the cooling of the stationary blade is used as the combustion air. As the cooling air, air compressed by a compressor of a gas turbine is used. Therefore, effective use of cooling air leads to an increase in efficiency of the gas turbine. Therefore, it is desired to use the cooling air more effectively.

Therefore, an object of the present invention is to provide a stationary blade segment capable of effectively utilizing the cooling air that cools the stationary blade, and a gas turbine and gas turbine equipment provided with the segment.

The stationary wing segment as one aspect of the invention for achieving the above object is
A stationary blade having a blade body extending in the radial direction with respect to the axis and an outer wing ring provided on the radial outer side of the stationary blade are provided, and the stationary blade is provided with cooling air from the radial inside with respect to the stationary blade. A first passage that flows in and causes the cooling air to flow out in the radial direction, and a temperature different from the temperature of the cooling air that flows in from the radial inside with respect to the stationary blade and flows out from the first passage. The outer wing ring has a second passage for discharging cooling air outward in the radial direction, and the outer wing ring is a first exhaust port for exhausting cooling air flowing outward in the radial direction from the first passage to the first outside. And a second exhaust port for exhausting the cooling air flowing out from the second passage to the outside in the radial direction to the second outside. The first passage is a passage that penetrates the stationary blade in the radial direction. Further, the stationary wing has a tubular shape with the first wing passage penetrating the inside of the wing body in the radial direction, and the stationary wing defines the first wing passage with respect to the inner surface of the first wing passage. It has a first insert, which is spaced apart from the first wing passage. The tubular first insert has a plurality of through holes that eject inner cooling air to the inner surface of the first wing passage. The inside of the tubular first insert forms part of the first passage.

In the stationary wing segment, the temperature of the cooling air flowing into the first exhaust port of the outer wing ring through the first passage of the stationary wing and the temperature of the cooling air flowing into the second exhaust port of the outer wing ring through the second passage of the stationary wing. It will be different from the temperature of the cooling air. The air that has flowed into the first exhaust port is discharged to the first outside. Further, the air flowing into the second exhaust port is discharged to a second outside different from the first outside. Therefore, in the stationary blade segment, the temperature of the cooling air flowing out from the first passage of the stationary blade is taken into consideration, and the temperature of the air flowing out from the second passage is sent to the first outside suitable for using the cooling air. In consideration of the above, the cooling air is sent to a second outside suitable for use. Therefore, in the stationary blade segment, the cooling air flowing out from the stationary blade can be effectively reused.

Here, in the stationary blade segment, the passage length of the second passage may be longer than the passage length of the first passage.

In the stationary blade segment, the temperature of the cooling air flowing into the second exhaust port via the second passage can be increased.

Further, in any of the above-mentioned stationary blade segments, the first passage is a passage that penetrates the stationary blade in the radial direction, and the second passage contains a directional component perpendicular to the radial direction. A portion extending in the holding direction may be included.

In the stationary blade segment, the temperature of the cooling air flowing into the second exhaust port via the second passage can be increased.

In any of the above-mentioned stationary blade segments, the stationary blade may have a communication passage that communicates with the second passage and the first passage.

In any of the above-mentioned stationary blade segments, the stationary blade has an inner shroud provided on the radial inner side of the blade body and an inner impinging plate attached to the inner shroud. The inner shroud extends from the radial inner end of the wing body in a direction having a directional component perpendicular to the radial direction, and from the peripheral edge along the peripheral edge of the inner shroud body. It has an inner peripheral wall that projects inward in the radial direction, and the inner impinge plate is arranged at intervals in the radial direction with respect to the inner shroud main body, and the inner shroud main body and the inner peripheral wall. Together, a plurality of inner cavities are formed radially outside the inner impinge plate, and the inner impinging plate guides cooling air from the inside radially inside the inner impinging plate to the inner cavity. A through hole may be formed and the inner cavity may form part of the second passage.

In the stationary blade segment, cooling air from the inside in the radial direction of the inner impinge plate flows into the inner cavity through a plurality of through holes of the inner impinging plate. The cooling air flowing into the inner cavity collides with the inner shroud body to impinge-cool the inner shroud body. In the stationary blade segment, since the inner cavity forms a part of the second passage, the temperature of the cooling air flowing into the second exhaust port through the second passage can be increased.

In the vane segment having the inner shroud, the inner shroud may have a plurality of ejection holes for ejecting cooling air in the inner cavity to the outside of the vane.

In the vane segment, the cooling air in the inner cavity is ejected out of the vane from the plurality of ejection holes. A part of the cooling air ejected from the ejection hole cools the surface of the inner shroud with a film.

As another aspect of the invention for achieving the above object, the stationary blade segment is
A stationary blade having a blade body extending in the radial direction with respect to the axis and an outer wing ring provided on the radial outer side of the stationary blade are provided, and the stationary blade is provided with cooling air from the radial inside with respect to the stationary blade. A first passage that flows in and causes the cooling air to flow out in the radial direction, and a temperature different from the temperature of the cooling air that flows in from the radial inside with respect to the stationary blade and flows out from the first passage. The outer wing ring has a second passage for discharging cooling air outward in the radial direction, and the outer wing ring is a first exhaust port for exhausting cooling air flowing outward in the radial direction from the first passage to the first outside. And a second exhaust port for exhausting the cooling air flowing out from the second passage to the outside in the radial direction to the second outside. Further, the stationary blade has an outer shroud provided on the radial outer side of the blade body and an outer impinge plate attached to the outer shroud, and the outer shroud is the outer shroud of the blade body. An outer shroud body extending in a direction having a directional component perpendicular to the radial direction from the radial outer end, and an outer peripheral wall protruding radially outward from the peripheral edge along the peripheral edge of the outer shroud body. The outer impinge plate has An outer cavity is formed on the inner side in the radial direction, and the outer impinge plate is formed with a plurality of through holes for guiding cooling air from the outer side in the radial direction to the outer cavity, and the outer cavity is formed. It forms a part of the second passage.

In the stationary blade segment, cooling air from the outer side in the radial direction of the outer impinge plate flows into the outer cavity through a plurality of through holes of the outer impinging plate. The cooling air that has flowed into the outer cavity collides with the outer shroud body to impinge-cool the outer shroud body. In the stationary blade segment, since the outer cavity forms a part of the second passage, the temperature of the cooling air flowing into the second exhaust port through the second passage can be increased.

In the vane segment having the outer shroud, the outer shroud may have a plurality of ejection holes for ejecting cooling air in the outer cavity to the outside of the vane.

In the vane segment, the cooling air in the outer cavity is ejected out of the vane from the plurality of ejection holes. A part of the cooling air ejected from the ejection hole cools the surface of the outer shroud with a film.

In any of the above vane segments having the inner shroud and the outer shroud, the vane penetrates the blade body in the radial direction and cools air from the inner cavity from the outer impinge plate. Also has a second wing passage leading outward in the radial direction, and the second wing passage may form a part of the second passage.

In the stationary wing segment having the second wing passage, the stationary wing has a tubular shape, and the stationary wing is spaced from the inner surface of the second wing passage that defines the second wing passage. The second insert having a second insert arranged in the second wing passage, and the tubular second insert may have a plurality of through holes for ejecting inner cooling air to the inner surface of the second wing passage. ..

In the stationary blade segment, the cooling air in the second insert is ejected to the outside of the second insert through the plurality of through holes of the second insert. The cooling air ejected to the outside of the second insert collides with the inner surface of the second wing passage to impinge-cool the inner surface of the second wing passage. In the stationary wing segment, since the second wing passage forms a part of the second passage, the temperature of the cooling air flowing into the second exhaust port through the second passage can be increased.

As yet another aspect of the invention for achieving the above object, the stationary blade segment is
A stationary blade having a blade body extending in the radial direction with respect to the axis and an outer wing ring provided on the radial outer side of the stationary blade are provided, and the stationary blade is provided with cooling air from the radial inside with respect to the stationary blade. A first passage that flows in and causes the cooling air to flow out in the radial direction, and a temperature different from the temperature of the cooling air that flows in from the radial inside with respect to the stationary blade and flows out from the first passage. The outer wing ring has a second passage for discharging cooling air outward in the radial direction, and the outer wing ring is a first exhaust port for exhausting cooling air flowing outward in the radial direction from the first passage to the first outside. And a second exhaust port for exhausting the cooling air flowing out from the second passage to the outside in the radial direction to the second outside. Further, the stationary wing has a tubular shape with the first wing passage penetrating the inside of the wing body in the radial direction, and the stationary wing defines the first wing passage with respect to the inner surface of the first wing passage. The tubular first insert has a first insert that is spaced into the first wing passage, and the tubular first insert has a plurality of through holes that eject inner cooling air to the inner surface of the first wing passage. The inside of the tubular first insert forms a part of the first passage, and the outside of the first insert forms a part of the second passage.

In the stationary blade segment, a part of the cooling air in the second insert is ejected to the outside of the first insert through the plurality of through holes of the first insert. The cooling air ejected to the outside of the first insert collides with the inner surface of the first wing passage to impinge-cool the inner surface of the first wing passage. In the stationary blade segment, since the outside of the first insert forms a part of the second passage, the temperature of the cooling air flowing into the second exhaust port through the second passage can be increased.

The gas turbine as one aspect of the invention for achieving the above object is
A turbine having any of the above-mentioned stationary blade segments, an air compressor that compresses air to generate compressed air, and fuel is burned in the compressed air to generate combustion gas, and the combustion gas is produced. A compressor that guides the turbine into the turbine is provided, and the turbine further covers a turbine rotor that rotates about the axis and an outer peripheral side of the turbine rotor, and the stationary blade segment is attached to the inner peripheral side. The turbine casing and the turbine rotor are arranged at different positions in the axial direction with respect to the rotor shaft extending in the axial direction with the axis as the center and fixed to the rotor shaft. It has a turbine and a moving wing.

The gas turbine equipment as one aspect of the invention for achieving the above object is
A stationary blade that discharges from the gas turbine and the air compressor and guides the compressed air before being used for burning the fuel in the combustor as the cooling air from the radial inside of the stationary blade into the stationary blade. A cooling line, a pressure-pressing compressor arranged in the stationary blade cooling line, which boosts the air flowing into the stationary blade cooling line and sends it to the stationary blade, and a high-temperature component in contact with the combustion gas in the gas turbine. A first line that connects the first high temperature component excluding the stationary blade and the first exhaust port of the outer wing ring and guides the cooling air in the first exhaust port to the first high temperature component. Among the high temperature parts in contact with the combustion gas in the gas turbine, the second high temperature parts excluding the stationary blade and the first high temperature part are connected to the second exhaust port of the outer wing ring, and the first (Ii) A second line for guiding the cooling air in the exhaust port to the second high temperature component is provided.

In the gas turbine equipment, the cooling air flowing out from the stationary blade can be reused in the first high temperature component and the second high temperature component which are the parts of the gas turbine.

Here, in the gas turbine equipment, the step-up compressor may be configured to have a compressor impeller fixed to the turbine rotor.

In the gas turbine equipment, since the step-up compressor is arranged inside the gas turbine, the equipment can be miniaturized. Further, in the gas turbine equipment, it is not necessary to separately provide a drive source for driving the step-up compressor.

As another aspect of the invention for achieving the above object, the gas turbine equipment is
It includes a gas turbine, a vane cooling line, a step-up compressor, a first line, and a second line. The gas turbine includes a turbine having a stationary blade segment, an air compressor that compresses air to generate compressed air, and fuel is burned in the compressed air to generate combustion gas, and the combustion gas is produced. It is equipped with a compressor that guides the turbine into the turbine. The turbine, in addition, a turbine rotor rotating about an axis, covering the outer periphery of the turbine rotor, a turbine casing, wherein the stator vane segments on the inner circumferential side is attached has. The turbine rotor about said axis, having a rotor shaft extending in the axial direction, and a moving blade which is fixed to the rotor shaft. The stationary blade segment is arranged at a different position in the axial direction with respect to the moving blade, and has a blade body having a blade body extending in the radial direction with respect to the axial line, and an outer side provided on the radial outer side of the stationary blade. It is equipped with a wing ring. In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. It has a second passage for allowing cooling air having a temperature different from the temperature of the cooling air flowing out from the first passage to flow outward in the radial direction. The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port for exhausting to the outside. The stationary blade cooling line discharges from the air compressor and guides the compressed air before being used for combustion of the fuel in the combustor as the cooling air from the radial inside of the stationary blade into the stationary blade. .. The step-up compressor is arranged in the vane cooling line, and boosts the air flowing into the vane cooling line and sends it to the vane. The first line connects the first high temperature component excluding the stationary blade and the first exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine, and the first line. The cooling air in the exhaust port is guided to the first high temperature component. The second line includes the second high temperature component excluding the stationary blade and the first high temperature component and the second exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine. Connect and guide the cooling air in the second exhaust port to the second high temperature component. The combustor has a tail cover that guides the combustion gas into the turbine, and the first high temperature component includes the tail cover.

In the gas turbine equipment, the tail cover of the combustor can be cooled by a part of the cooling air flowing out from the stationary blade.

As yet another aspect of the invention for achieving the above object, the gas turbine equipment is
It includes a gas turbine, a vane cooling line, a step-up compressor, a first line, and a second line. The gas turbine includes a turbine having a stationary blade segment, an air compressor that compresses air to generate compressed air, and fuel is burned in the compressed air to generate combustion gas, and the combustion gas is produced. It is equipped with a compressor that guides the turbine into the turbine. The turbine, in addition, a turbine rotor rotating about an axis, covering the outer periphery of the turbine rotor, a turbine casing, wherein the stator vane segments on the inner circumferential side is attached has. The turbine rotor about said axis, a rotor shaft extending in the axial direction, are arranged at positions different in the axial direction with respect to the stationary blade, a moving blade which is fixed to the rotor shaft, the Have. The stationary blade segment is arranged at a different position in the axial direction with respect to the moving blade, and has a blade body having a blade body extending in the radial direction with respect to the axial line, and an outer side provided on the radial outer side of the stationary blade. It is equipped with a wing ring. In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. It has a second passage for allowing cooling air having a temperature different from the temperature of the cooling air flowing out from the first passage to flow outward in the radial direction. The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port for exhausting to the outside. The stationary blade cooling line discharges from the air compressor and guides the compressed air before being used for combustion of the fuel in the combustor as the cooling air from the radial inside of the stationary blade into the stationary blade. .. The step-up compressor is arranged in the vane cooling line, and boosts the air flowing into the vane cooling line and sends it to the vane. The first line connects the first high temperature component excluding the stationary blade and the first exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine, and the first line. The cooling air in the exhaust port is guided to the first high temperature component. The second line includes the second high temperature component excluding the stationary blade and the first high temperature component and the second exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine. Connect and guide the cooling air in the second exhaust port to the second high temperature component. Further, it is provided with a cooler that is arranged in the second line and cools the air that has flowed into the second line. The second high temperature component includes the turbine rotor.

In the gas turbine equipment, the turbine rotor can be cooled by a part of the cooling air flowing out from the stationary blade.

In one aspect of the present invention, the cooling air that cools the stationary blade can be effectively used.

It is a notch side view of the main part of the gas turbine in one Embodiment which concerns on this invention. It is sectional drawing of the main part of the gas turbine equipment in one Embodiment which concerns on this invention. It is sectional drawing of the stationary blade segment in one Embodiment which concerns on this invention. FIG. 3 is a sectional view taken along line IV in FIG. FIG. 3 is a cross-sectional view taken along line V in FIG. FIG. 3 is a sectional view taken along line VI in FIG. FIG. 3 is a sectional view taken along line VII in FIG. FIG. 3 is a sectional view taken along line VIII in FIG. FIG. 3 is a cross-sectional view taken along the line IX in FIG. It is an X-ray sectional view in FIG. It is sectional drawing of the stationary blade segment in one modification of one Embodiment which concerns on this invention. It is sectional drawing of the main part of the gas turbine equipment in one modification of one Embodiment which concerns on this invention.

Hereinafter, an embodiment of a gas turbine facility including a stationary blade segment according to the present invention, and a modified example of the stationary blade segment will be described in detail with reference to the drawings.

"Embodiment"
An embodiment of the gas turbine equipment according to the present invention will be described with reference to FIGS. 1 to 10.

As shown in FIG. 1, the gas turbine equipment according to the embodiment of the present invention includes a gas turbine.

The gas turbine 10 includes an air compressor 20 that compresses air A, a combustor 30 that burns fuel F in the air A compressed by the air compressor 20 to generate combustion gas, and a turbine driven by the combustion gas. 40 and.

The air compressor 20 includes a compressor rotor 21 that rotates about an axis Ar, a compressor casing 25 that covers the compressor rotor 21, and a plurality of stationary blade rows 26. The turbine 40 has a turbine rotor 41 that rotates about the axis Ar, a turbine casing 45 that covers the turbine rotor 41, and a plurality of stationary blade rows 46.

The compressor rotor 21 and the turbine rotor 41 are located on the same axis Ar and are connected to each other to form the gas turbine rotor 11. For example, a rotor of a generator is connected to the gas turbine rotor 11. The gas turbine 10 further includes an intermediate casing 14 arranged between the compressor casing 25 and the turbine casing 45, and an inner casing 16 arranged in the intermediate casing 14. The compressor casing 25, the intermediate casing 14, and the turbine casing 45 are connected to each other to form the gas turbine casing 15. In the following, the direction in which the axis Ar extends is referred to as the axial direction Da, the circumferential direction centered on the axis Ar is simply referred to as the circumferential direction Dc, and the direction perpendicular to the axis Ar is referred to as the radial direction Dr. Further, in the axial direction Da, the air compressor 20 side is referred to as the axial upstream side Dau, and the opposite side is referred to as the axial downstream side Dad with reference to the turbine 40. Further, the side approaching the axis Ar in the radial direction is the radial inner Dri, and the opposite side is the radial outer Dro.

The compressor rotor 21 has a rotor shaft 22 extending in the axial direction Da about the axis Ar, and a plurality of rotor blade rows 23 attached to the rotor shaft 22. The plurality of blade rows 23 are arranged in the axial direction Da. Each rotor blade row 23 is composed of a plurality of rotor blades 23a arranged in the circumferential direction Dc. A stationary blade row 26 is arranged on the Dad on the downstream side of each axis of the plurality of blade rows 23. Each vane row 26 is provided inside the compressor casing 25. Each of the stationary blade rows 26 is composed of a plurality of stationary blades 26a arranged in the circumferential direction Dc. A suction port 27 is formed in the Dau on the upstream side of the axis of the compressor casing 25.

The turbine rotor 41 has a rotor shaft 42 extending in the axial direction Da about the axis Ar, and a plurality of blade rows 43 attached to the rotor shaft 42. The plurality of blade rows 43 are arranged in the axial direction Da. Each rotor blade row 43 is composed of a plurality of rotor blades 43a arranged in the circumferential direction Dc. A stationary blade row 46 is arranged on each axis upstream Dau of the plurality of rotor blade rows 43. Each vane row 46 is provided inside the turbine casing 45. Each of the stationary blade rows 46 is composed of a plurality of stationary blades 50 arranged in the circumferential direction Dc.

The air compressor 20 takes in air A from the suction port 27 and compresses the air A to generate compressed air. This compressed air flows out from the discharge port 28 of the air compressor 20 and flows into the combustor 30. In addition to the compressed air, the fuel F from the fuel supply source is also supplied to the combustor 30. In the combustor 30, the fuel F is burned in the compressed air to generate the combustion gas G. This combustion gas G flows into the turbine casing 45 and rotates the turbine rotor 41.

As shown in FIG. 2, both the intermediate casing 14 and the inner casing 16 have a tubular shape centered on the axis Ar. The inner casing 16 is arranged on the inner peripheral side of the intermediate casing 14 at intervals from the intermediate casing 14. An intermediate casing 13 of the gas turbine 10 is formed between the inner peripheral side of the intermediate casing 14 and the outer peripheral side of the inner casing 16. A gas turbine rotor 11 is arranged on the inner peripheral side of the tubular inner casing 16. The combustor 30 is arranged in the intermediate vehicle compartment 13. The combustor 30 is fixed to the intermediate casing 14 by a support 39.

The combustor 30 includes a diffuser 31, a combustion cylinder 32, a tail cylinder 33, a pilot burner 34, and a plurality of main burners 35. The diffuser 31 has a tubular shape. Both ends of the diffuser 31 are open. The first end of the diffuser 31 is connected to the discharge port 28 of the air compressor 20. Therefore, the compressed air from the air compressor 20 flows into the diffuser 31. The cross-sectional area of the flow path in the diffuser 31 gradually increases from the first end to the second end of the diffuser 31. Therefore, in the process of passing through the diffuser 31, the flow velocity of the compressed air gradually decreases, while the pressure of the compressed air gradually increases. Therefore, the diffuser 31 plays a role of converting the dynamic pressure of the compressed air into a static pressure. The diffuser 31 is formed with a plurality of through holes 31a penetrating from the inner peripheral side to the outer peripheral side. Therefore, a part of the compressed air that has flowed into the diffuser 31 flows into the intermediate vehicle interior 13 through the plurality of through holes 31a.

The combustion cylinder 32 has a cylindrical shape. Both ends of the combustion cylinder 32 are open. The first end of the combustion cylinder 32 is connected to the second end of the diffuser 31. Compressed air from the diffuser 31 flows into the combustion cylinder 32. The tail cover 33 has a tubular shape. Both ends of the tail cover 33 are open. The first end of the tail cylinder 33 is connected to the second end of the combustion cylinder 32. The second end of the tail cover 33 is connected to a part of the turbine 40.

The pilot burner 34 is arranged on the central axis of the cylindrical combustion cylinder 32. The plurality of main burners 35 are arranged in the circumferential direction with respect to the central axis of the combustion cylinder 32 with the pilot burner 34 as the center. Fuel F from a fuel supply source is supplied to the pilot burner 34 and the plurality of main burners 35 via the fuel line 38. The pilot burner 34 injects fuel F together with compressed air into the combustion cylinder 32. This fuel F is diffusely burned in the combustion cylinder 32 and the tail cylinder 33. In the plurality of main burners 35, the fuel F and the compressed air are mixed to generate a premixed gas. Each main burner 35 injects this premixed gas into the combustion cylinder 32. This premixed gas is premixed and burned in the combustion cylinder 32 and the tail cylinder 33. The high-temperature and high-pressure combustion gas G generated by the combustion of the fuel F is sent to the turbine 40 via the tail cover 33.

The stationary blade 50 of the turbine 40 has a blade body 51 extending in the radial direction Dr, an inner shroud 60i provided on the radial inner Dri of the blade body 51, and an outer side provided on the radial outer side Dr of the blade body 51. It has a shroud 60o and. An outer wing ring 90 is arranged on the radial inner Dri of the turbine casing 45, which is the radial outer Dro of the outer shroud 60o. The outer shrouds 60o of the plurality of vanes 50 constituting one vane row 46 are attached to the outer wing ring 90. The outer wing ring 90 is fixed to the turbine casing 45. In the present embodiment, the outer wing ring 90 and the plurality of stationary blades 50 attached to the outer wing ring 90 form a stationary blade segment S.

The rotor blades 43a of the turbine 40 include a blade body 43b extending in the radial direction, a platform 43f provided on the radial inner Dri of the blade body 43b, and a blade root 43r extending from the platform 43f to the radial inner Dri. Have. The blade root 43r of the rotor blade 43a is embedded and fixed in the rotor shaft 42. A dividing ring 48 is arranged on the radial outer Dro of the rotor blade 43a. The split ring 48 is attached to the outer wing ring 90 described above. The position of the radial Dr of the platform 43f substantially coincides with the position of the radial Dr of the inner shroud 60i of the stationary blade 50 arranged on the Dau on the upstream side of the axis. The position of the radial Dr of the dividing ring 48 substantially coincides with the position of the radial Dr of the outer shroud 60o of the stationary blade 50 arranged on the Dau on the upstream side of the axis.

The combustion gas flow path 49 through which the combustion gas G flows in the turbine 40 forms an annular shape with the axis Ar as the center. The inner peripheral edge of the annular combustion gas flow path 49 is defined by the inner shroud 60i of the stationary blade 50 and the platform 43f of the moving blade 43a. The outer peripheral edge of the combustion gas flow path 49 is defined by the outer shroud 60o of the vane 50 and the dividing ring 48. Both the blade body 51 of the stationary blade 50 and the blade body 43b of the moving blade 43a are arranged in the combustion gas flow path 49. The combustion gas inlet 49i in the combustion gas flow path 49 is the inner shroud 60i and the outer side of the plurality of stationary blades 50a constituting the first-stage stationary blade row 46a located on the most upstream side Dau of the axis lines among the plurality of stationary blade rows 46. It is formed by a shroud 60o. Therefore, the second end of the tail cover 33 of the combustor 30 is connected to the inner shroud 60i and the outer shroud 60o of the plurality of stationary blades 50a constituting the first stage stationary blade row 46a.

The combustion cylinder 32 and tail cylinder 33 of the combustor 30, the stationary blade 50, the moving blade 43a, and the split ring 48 of the turbine 40 constitute high-temperature components that are exposed to the high-temperature combustion gas G. Therefore, these high temperature parts need to be cooled by some means. In the present embodiment, the stationary blade 50 and the moving blade 43a are cooled from the inside with cooling air. Therefore, the moving blade 43a has a cooling air passage 43p through which cooling air flows. Further, the first passage 81 and the second passage 82 are formed as cooling air passages in the plurality of stationary blades 50a constituting the first stage stationary blade row 46a.

An inner wing ring 100 is arranged on the radial inner Dri of the first-stage stationary blade row 46a. The inner blade ring 100 is formed with stationary blade cooling air passages 101 and 108 for supplying cooling air from the outside from the radial inner Dri into the stationary blade 50a.

The inner casing 16 has a mounting portion 17, an outer cylinder 18, and an inner cylinder 19. The mounting portion 17 is mounted on the compressor casing 25 or the intermediate casing 14. The mounting portion 17 extends radially inwardly from a portion mounted on the compressor casing 25 or the intermediate casing 14. Both the outer cylinder 18 and the inner cylinder 19 have a cylindrical shape centered on the axis Ar. The end of the Dau on the upstream side of the axis of the outer cylinder 18 is fixed to the radial inner Dri end of the mounting portion 17. Further, the end of the Dad on the downstream side of the axis of the outer cylinder 18 is attached to the inner wing ring 100 of the first-stage stationary blade row 46a. The space between the outer peripheral side of the outer cylinder 18 and the inner peripheral side of the intermediate casing 14 forms the above-mentioned intermediate casing 13. The inner cylinder 19 is a radial inner Dri of the outer cylinder 18 and is arranged on the radial outer Dro of the gas turbine rotor 11. The end of the Dau on the upstream side of the axis of the inner cylinder 19 is fixed to the outer cylinder 18. Further, the end of the Dad on the downstream side of the axis of the inner cylinder 19 is attached to the inner wing ring 100 of the first-stage stationary blade row 46a.

A stationary blade cooling air passage 108 through which cooling air for cooling the stationary blades 50a constituting the first-stage stationary blade row 46a flows between the radial outer Dro of the inner cylinder 19 and the radial inner Dri of the outer cylinder 18. To form. The stationary blade cooling air passage 108 communicates with the stationary blade cooling air passage 101 of the inner wing ring 100. A rotor cooling air passage 109 through which cooling air for cooling a part of the turbine rotor 41 flows is formed between the radial outer Dro of the gas turbine rotor 11 and the radial inner Dri of the inner cylinder 19. In the turbine rotor 41, a rotor blade cooling air passage 41p for supplying cooling air from the radial inner Dri is formed in a plurality of rotor blades 43a constituting the first stage rotor blade row 43. The first end of the moving blade cooling air passage 41p communicates with the rotor cooling air passage 109 on the inner peripheral side of the inner casing 16, and the second end of the moving blade cooling air passage 41p is the cooling air passage of the moving blade 43a. It communicates with 43p.

As shown in FIG. 7, the blade body 51 of the stationary blade 50a has an airfoil having a cross-sectional shape perpendicular to the radial direction Dr. In the wing body 51, the end of the Dau on the upstream side of the axis forms the leading edge 52, and the end of the Dad on the downstream side of the axis forms the trailing edge 53. On the surface of the wing body 51, of the surfaces facing the circumferential direction Dc, the convex surface forms the dorsal side surface (= negative pressure surface) 54, and the concave surface forms the ventral side surface (= positive pressure surface) 55. For convenience of the following explanation, the ventral side (= positive pressure surface side) of the wing body 51 is the circumferential ventral DcpDcp, and the dorsal side (= negative pressure surface side) of the wing body 51 is the circumferential dorsal DcnDcn in the circumferential direction Dc. And.

As shown in FIGS. 3 to 10, the stationary blade 50a constituting the first-stage stationary blade row 46a has a first wing passage 75, a second wing passage 76, and a third wing passage group 77. The first wing passage 75, the second wing passage 76, and the third wing passage group 77 all penetrate the wing body 51 in the radial direction Dr. The third wing passage group 77 is a collection of a large number of passages extending in the radial direction Dr. The first wing passage 75, the second wing passage 76, and the third wing passage group 77 are arranged in this order from the front edge side to the trailing edge side of the wing body 51. That is, the first wing passage 75 is located closest to the leading edge portion 52 in the wing body 51, and the third wing passage group 77 is located closest to the trailing edge portion 53 in the wing body 51. .. The second wing passage 76 is located between the first wing passage 75 and the third wing passage. The first wing passage 75 and the second wing passage 76 are separated by a first partition wall 78 formed in the wing body 51. The second wing passage 76 and the third wing passage group 77 are separated by a second partition wall 79 formed in the wing body 51.

The inner shroud 60i of the stationary blade 50a has an inner shroud main body 61i and an inner peripheral wall 63i, as shown in FIGS. 3 and 9. The inner shroud body 61i extends from the end of the radial inner Dri of the wing body 51 in a direction having a directional component perpendicular to the radial Dr. The surface of the inner shroud body 61i facing the outer Dro in the radial direction forms a gas path surface 62ip in contact with the combustion gas G. Further, the surface of the inner shroud main body 61i facing the inner Dri in the radial direction forms an anti-gas path surface 62io. The inner shroud body 61i has a parallelogram shape when viewed from the radial direction Dr. Of the pair of sides parallel to each other in the inner shroud main body 61i, one side faces Dau on the upstream side of the axis, and the other side faces Dad on the downstream side of the axis. Further, of the remaining pair of the inner shroud main body 61i parallel to each other, one side faces the circumferential ventral Dcp and the other side faces the circumferential dorsal Dcn. The inner peripheral wall 63i projects along the peripheral edge of the inner shroud main body 61i from this peripheral edge to the inner Dri in the radial direction. The inner peripheral wall 63i has a front peripheral wall 63f, a rear peripheral wall 63b, a ventral peripheral wall 63p, and a dorsal peripheral wall 63n. The front peripheral wall 63f projects from the side of the inner shroud main body 61i facing the upstream side Dau to the inner Dri in the radial direction. The rear peripheral wall 63b projects from the side of the inner shroud main body 61i facing the downstream side of the axis to the inner Dri in the radial direction. The ventral peripheral wall 63p projects from the side of the medial shroud body 61i facing the ventral Dcp in the circumferential direction to the medial Dri in the radial direction. The dorsal peripheral wall 63n projects from the side of the inner shroud main body 61i facing the dorsal Dcn in the circumferential direction toward the inner Dri in the radial direction. The inner shroud 60i is formed with a recess 64i recessed in the radial outer Dro. The side peripheral surface of the recess 64i is formed by an inner peripheral wall 63i. The bottom surface of the recess 64i is formed by the inner shroud body 61i.

In the recess 64i of the inner shroud 60i, the airfoil inner protruding portion 67i protrudes from the inner shroud main body 61i. The airfoil inner projecting portion 67i projects radially inward from a position along the blade surface of the blade body 51. The airfoil inner protrusion 67i has substantially the same shape as the airfoil of the airfoil body 51 when viewed from the radial direction Dr. The airfoil inner protruding portion 67i includes a portion of the radial inner Dri of the first wing passage 75, a portion of the radial inner Dri of the second wing passage 76, and a portion of the radial inner Dri of the third wing passage group 77. It is formed. The airfoil inner protrusion 67i is formed with a communication hole 67ic that penetrates from the second blade passage 76 to the outside of the airfoil inner protrusion 67i.

The vane 50a further has an inner impinging plate 68i and an inner sealing plate 69i. Both the inner impinging plate 68i and the inner sealing plate 69i are arranged in the recess 64i of the inner shroud 60i at intervals in the radial inner Dri with respect to the inner shroud main body 61i. The inner impinge plate 68i extends from the airfoil inner protrusion 67i in a direction perpendicular to the radial direction Dr. The inner impinge plate 68i, together with the inner shroud main body 61i, the inner peripheral wall 63i, and the airfoil inner protruding portion 67i, forms an inner cavity 65i on the outer Dro in the radial direction from the inner impinging plate 68i. The inner impinge plate 68i is formed with a large number of through holes 68t penetrating in the radial direction Dr. The inner shroud 60i is formed with a plurality of ejection holes 60e penetrating the combustion gas flow path 49 from the inside of the inner cavity 65i. The inner sealing plate 69i closes the opening of the radial inner Dri among the openings at both ends of the radial Dr in the second wing passage 76.

The outer shroud 60o of the stationary blade 50a has an outer shroud main body 61o and an outer peripheral wall 63o, as shown in FIGS. 3 to 6. The outer shroud body 61o extends from the end of the radial outer Dr of the blade body 51 in a direction having a directional component perpendicular to the radial Dr. The surface of the outer shroud body 61o facing the inner Dri in the radial direction forms a gas path surface 62op in contact with the combustion gas G. Further, the surface of the outer shroud body 61o facing the outer Dro in the radial direction forms an anti-gas path surface 62oo. Like the inner shroud main body 61i, the outer shroud main body 61o has a parallelogram shape when viewed from the radial direction Dr. Of the pair of sides parallel to each other in the outer shroud body 61o, one side faces the Axis upstream side Dau, and the other side faces the Axis downstream side Dad. Further, of the remaining pair of outer shroud bodies 61o parallel to each other, one side faces the circumferential ventral Dcp and the other side faces the circumferential dorsal Dcn. The outer peripheral wall 63o projects along the peripheral edge of the outer shroud main body 61o from this peripheral edge to the outer Dro in the radial direction. The outer peripheral wall 63o has a front peripheral wall 63f, a rear peripheral wall 63b, a ventral peripheral wall 63p, and a dorsal peripheral wall 63n. The front peripheral wall 63f projects from the side of the outer shroud main body 61o facing the upstream side Dau to the outer Dro in the radial direction. The rear peripheral wall 63b projects from the side of the outer shroud main body 61o facing the downstream side of the axis to the outer Dro in the radial direction. The ventral peripheral wall 63p projects from the side of the outer shroud body 61o facing the ventral Dcp in the circumferential direction to the outer Dro in the radial direction. The dorsal peripheral wall 63n projects from the side of the outer shroud main body 61o facing the dorsal Dcn in the circumferential direction toward the outer Dro in the radial direction. The outer shroud 60o is formed with a recess 64o recessed in the radial inner Dri. The side peripheral surface of the recess 64o is formed by an outer peripheral wall 63o. The bottom surface of the recess 64o is formed by the outer shroud body 61o.

In the recess 64o of the outer shroud 60o, the airfoil outer protrusion 67o protrudes from the outer shroud main body 61o. The airfoil outer protrusion 67o protrudes radially outward from a position along the blade surface of the blade body 51. The airfoil outer protrusion 67o has substantially the same shape as the airfoil of the airfoil body 51 when viewed from the radial direction Dr. The airfoil outer protrusion 67o includes a radial outer Dro portion of the first wing passage 75, a radial outer Dro portion of the second wing passage 76, and a radial outer Dro portion of the third wing passage group 77. It is formed. The airfoil outer protrusion 67o is formed with a communication hole 67oc that penetrates from the second wing passage 76 to the outside of the airfoil outer protrusion 67o. The first partition wall 78 and the second partition wall 79 project radially outward from the airfoil outer protrusion 67o. The positions of the ends of the radial outer Dro of the first partition wall 78 and the second partition wall 79 are substantially the same as the positions of the ends of the outer peripheral wall 63o on the radial outer Dro in the radial direction Dr.

The vane 50a further has an outer impinging plate 68o and an outer sealing plate 69o. Both the outer impinging plate 68o and the outer sealing plate 69o are arranged in the recess 64o of the outer shroud 60o at intervals in the outer Dro in the radial direction with respect to the outer shroud main body 61o. The outer impinging plate 68o extends from the airfoil outer protrusion 67o in a direction perpendicular to the radial direction Dr. The outer impinge plate 68o, together with the outer shroud main body 61o, the outer peripheral wall 63o, and the airfoil outer protruding portion 67o, forms an outer cavity 65o on the inner Dri in the radial direction from the outer impinging plate 68o. A large number of through holes 68t penetrating in the radial direction Dr are formed in the outer impinging plate 68o. The outer shroud 60o is formed with a plurality of ejection holes 60e penetrating the combustion gas flow path 49 from the inside of the outer cavity 65o.

The vane 50a further has a first insert 70a and a second insert 70b. The first insert 70a has a tubular insert body 71a and a sealing portion 72a extending outward from the tubular insert body 71a. Both ends of the tubular insert body 71a are open. The insert body 71a is formed with a large number of through holes 71t penetrating from the inside to the outside. The first insert 70a is arranged in the first wing passage 75 so that both ends of the insert body 71a face the radial direction Dr. There is a gap between the outer peripheral surface of the insert body 71a and the inner surface 75i of the first wing passage that defines the first wing passage 75 with the stationary blade 50a. A front gap passage 83 is formed between the outer peripheral surface of the insert body 71a and the inner surface 75i of the first wing passage. A seal portion 72a is provided at the end of the radial inner Dri of the insert body 71a. The seal portion 72a is in contact with the inner surface 75i of the first wing passage. Therefore, the end of the radial inner Dri of the front gap passage 83 is closed by the seal portion 72a. The end of the radial outer Dro of the insert body 71a protrudes radially outer Dro from the airfoil outer protruding portion 67o in which a part of the first blade passage 75 is formed.

The outer sealing plate 69o is arranged in the recess 64i of the outer shroud 60o at intervals in the outer Dro in the radial direction with respect to the outer impinging plate 68o. The outer sealing plate 69o is the radial direction of the airfoil outer protrusion 67o forming the end of the radial outer Dro of the first insert 70a, a part of the second wing passage 76, and a part of the third wing passage group 77. As shown in FIG. 4, it extends from the end of the outer Dro in a direction perpendicular to the radial direction Dr and is in contact with the outer peripheral wall 63o of the outer shroud 60o. The inside of the recess 64o of the outer shroud 60o is partitioned into a space radially inner Dri from the outer sealing plate 69o and a space radially outer Dro from the outer sealing plate 69o. Within the recess 64o of the outer shroud 60o, the space radially inner Dri from the outer sealing plate 69o and radially outer Dro from the outer impinge plate 68o forms the impinge air retention space 66. The front gap passage 83 communicates with the impinge air retention space 66. The cooling air that has flowed into the impinge air retention space 66 from the front gap passage 83 flows into the outer cavity 65o through a large number of through holes 68t of the outer impinging plate 68o.

The second insert 70b has a tubular insert body 71b, a seal portion 72b extending outward from the tubular insert body 71b, and a lid portion 73b. Both ends of the tubular insert body 71b are open. The insert body 71b is formed with a large number of through holes 71t penetrating from the inside to the outside. The second insert 70b is arranged in the second wing passage 76 so that both ends of the insert body 71b face the radial direction Dr. There is a gap between the outer peripheral surface of the insert body 71b and the inner surface 76i of the second wing passage that defines the second wing passage 76 with the stationary blade 50a. An intermediate gap passage 84 is formed between the outer peripheral surface of the insert body 71b and the inner surface 76i of the second wing passage. A seal portion 72b is provided at the end of the radial outer Dro of the insert body 71b. The seal portion 72b is in contact with the inner surface 76i of the second wing passage. Therefore, the end of the radial inner Dri of the intermediate gap passage 84 is closed by the seal portion 72b. The opening of the radial outer Dro of the insert body 71b is closed by the lid portion 73b. The lid portion 73b is located on the inner Dri in the radial direction with respect to the outer sealing plate 69o. The inside of the second insert 70b forms an intermediate passage 85.

In the space of the outer Dro in the radial direction from the outer sealing plate 69o in the recess 64o of the outer shroud 60o, as shown in FIG. 4, the above-mentioned first partition wall 78 and the second partition wall 79 extend in the circumferential direction Dc. , In contact with the ventral peripheral wall 63p and the dorsal peripheral wall 63n of the outer shroud 60o. Therefore, in the recess 64o of the outer shroud 60o, the space of the outer Dr in the radial direction from the outer sealing plate 69o is formed by the first partition wall 78 and the second partition wall 79 in the front space 67s, the intermediate space 68s, and the rear space 69s. It is partitioned into. The front space 67s, the intermediate space 68s, and the posterior space 69s are arranged in this order toward the cad on the downstream side of the axis.

As described above, the stationary blade 50a is formed with a first passage 81 and a second passage 82 as cooling air passages. The first passage 81 is formed in the recess 64i of the inner shroud 60i by the space radially inner Dri from the inner impinge plate 68i, the space inside the first insert 70a, and the front space 67s of the outer shroud 60o. .. The first passage 81 is a passage that extends substantially linearly in the radial direction Dr and penetrates the stationary blade 50a from the radial inner Dri to the radial outer Dro. The second passage 82 includes a space other than the space forming the first passage 81 in the stationary blade 50a, a space in the recess 64i of the inner shroud 60i, and a space in the inner Dri in the radial direction of the inner impinge plate 68i. It is formed with the space in the insert 70a. Therefore, in the recess 64i of the inner shroud 60i, the space of the inner Dri in the radial direction of the inner impinge plate 68i and the space in the first insert 70a are shared by the first passage 81 and the second passage 82. Further, as described above, a large number of through holes 71t penetrating from the inside to the outside are formed in the insert body 71a of the first insert 70a. Therefore, the second passage 82 communicates with the first passage 81 through a large number of through holes 71t of the first insert 70a.

As shown in FIG. 3, the outer wing ring 90 includes a wing ring plate 91 extending in the circumferential direction Dc, a wing ring peripheral wall 92 extending from the peripheral edge of the wing ring plate 91 to the inner Dri in the radial direction, and a first partition wall 93. It has a second partition wall 94 and a casing mounting portion 95. The outer wing ring 90 is formed with a recess 96 recessed in the radial outer Dro. The side peripheral surface of the recess 96 is formed by a wing ring peripheral wall 92. Further, the bottom surface of the recess 96 is formed of a wing ring plate 91. The first partition wall 93 and the second partition wall 94 are arranged in the recess 96 and are arranged in the axial direction Da. The recess 96 is divided into three spaces by a first partition wall 93 and a second partition wall 94. Of the three spaces, the space on the upstream side of the axis Dau forms the first exhaust port 97, and the space adjacent to the Dad on the downstream side of the axis of the first exhaust port 97 forms the second exhaust port 98. The blade ring peripheral wall 92 is formed with a first exhaust port 97o that penetrates from the inside of the first exhaust port 97 to the outside of the outer blade ring 90. Further, the wing ring plate 91 is formed with a second exhaust port 98o that penetrates from the inside of the second exhaust port 98 to the outside of the outer wing ring 90. The first exhaust port 97 communicates with the front space 67s of the stationary blade 50a. Therefore, this exhaust port communicates with the first passage 81 of the stationary blade 50a. Further, the second exhaust port 98 communicates with the intermediate space 68s of the stationary blade 50a. Therefore, the second exhaust port 98 communicates with the second passage 82 of the stationary blade 50a.

Of the wing ring peripheral wall 92, the front peripheral wall 92f facing the upstream side Dau of the axis line is substantially in contact with the front peripheral wall 63f of the outer shroud 60o. The first partition wall 93 of the outer wing ring 90 is in contact with the first partition wall 78 of the stationary blade 50a, and supports the first partition wall 78 of the stationary blade 50a from the radial outer Dro. The second partition wall 94 of the outer wing ring 90 is in contact with the second partition wall 79 of the stationary blade 50a, and supports the second partition wall 79 of the stationary blade 50a from the radial outer Dro. Of the wing ring peripheral wall 92, the rear peripheral wall 92b facing the downstream side of the axis Dad is in contact with the rear peripheral wall 63b of the outer shroud 60o, and supports the rear peripheral wall 63b of the outer shroud 60o from the radial outer Dro. A sealing material 99 is arranged between the rear peripheral wall 92b of the wing ring peripheral wall 92 and the rear peripheral wall 63b of the outer shroud 60o. The rear peripheral wall 63b of the outer shroud 60o has a more rigid structure than the front peripheral wall 63f of the outer shroud 60o, the first partition wall 78 of the outer shroud 60o, and the second partition wall 79 of the outer shroud 60o.

The casing mounting portion 95 projects from the wing ring plate 91 on the downstream side of the axis to the radial outer Dro. As shown in FIG. 2, the casing mounting portion 95 is mounted on the turbine casing 45.

In addition to the gas turbine 10, the gas turbine equipment includes a stationary blade cooling line 105, a first line 111, a second line 112, a step-up compressor 115, a cooler 116, and the like, as shown in FIG. To be equipped with.

The vane cooling line 105 is composed of a vane cooling pipe 106, a vane cooling air passage 108 of the inner casing 16, and a vane cooling air passage 101 of the inner wing ring 100. The first end of the stationary blade cooling pipe 106 is connected to the intermediate casing 14, and the second end of the stationary blade cooling pipe 106 is connected to the outer cylinder 18 of the inner casing 16. The stationary blade cooling pipe 106 guides the compressed air in the intermediate casing 13 into the stationary blade cooling air passage 108 of the inner casing 16 as cooling air. A boost compressor 115 is provided in the vane cooling pipe 106 to boost the compressed air flowing into the vane cooling pipe 106 and send it into the vane cooling air passage 108 of the inner casing 16.

One end of the first line 111 is connected to the first exhaust port 97o of the outer wing ring 90, and the second end of the first line 111 is connected to the tail cover 33 of the combustor 30. A plurality of passages are formed in the plate forming the tail cover 33. The first line 111 sends the air in the first exhaust port 97 of the outer wing ring 90 to the passage of the tail cover 33. The first line 111 may be formed of, for example, a steel pipe, or may be formed of a flexible pipe.

The second line 112 includes a rotor cooling pipe 113 and a rotor cooling air passage 109 of the inner casing 16. The first end of the rotor cooling pipe 113 is connected to the second exhaust port 98o of the outer wing ring 90, and the second end of the rotor cooling pipe 113 is connected to the inner cylinder 19 of the inner casing 16. The rotor cooling pipe 113 guides the air in the second exhaust port 98 of the outer wing ring 90 into the rotor cooling air passage 109 of the inner casing 16. A cooler 116 for cooling the air flowing into the rotor cooling pipe 113 is provided in the rotor cooling pipe 113. The cooler 116 is, for example, a heat exchanger that cools the air by exchanging heat between the air and the cooling medium. As the cooling medium, for example, fuel F flowing through the fuel line 38 can be considered. When the fuel F is used as a cooling medium, the fuel F is heated by heat exchange with air. Therefore, the cooler 116 in this case also functions as a fuel preheater.

Next, the flow of the compressed air discharged from the air compressor 20 will be described.

The compressed air discharged from the air compressor 20 flows into the diffuser 31 of the combustor 30. The temperature of this compressed air is, for example, 200 ° C. As described above, a part of the compressed air flowing into the diffuser 31 is used for combustion of the fuel F in the combustor 30. The rest of the compressed air that has flowed into the diffuser 31 flows into the intermediate passenger compartment 13 through a large number of through holes 31a formed in the diffuser 31. The compressed air that has flowed into the intermediate vehicle compartment 13 flows into the stationary blade cooling pipe 106. The compressed air that has flowed into the stationary blade cooling pipe 106 is boosted by the step-up compressor 115 provided in the stationary blade cooling pipe 106, and then is used as cooling air in the stationary blade cooling air passage 108 of the inner casing 16. Inflow into. The cooling air that has flowed into the stationary blade cooling air passage 108 of the inner casing 16 passes through the stationary blade cooling air passage 101 of the inner blade ring 100, and is the first of the plurality of stationary blades 50a constituting the first stage stationary blade row 46a. It flows into the passage 81 and the second passage 82.

The cooling air that has flowed into the first passage 81 exchanges heat with the stationary blade 50a in the process of flowing through the first passage 81 of the stationary blade 50a, and is heated while cooling the stationary blade 50a. The cooling air heated in the process of flowing through the first passage 81 flows into the first exhaust port 97 of the outer wing ring 90. The cooling air that has flowed into the first exhaust port 97 flows into the passage of the tail cover 33 via the first line 111. The cooling air that has flowed into the passage of the tail cover 33 is, for example, exhausted into the tail cover 33. The cooling air cools the tail cover 33 in the process of flowing through the passage of the tail cover 33. The plate forming the combustion cylinder 32 may be formed with a plurality of passages in the same manner as the plate forming the tail cylinder 33. In this case, the passage may be communicated with the passage of the tail cylinder 33, and cooling air may be passed through the passage of the combustion cylinder 32 to cool the combustion cylinder 32.

Although the pressure inside the tail cover 33 is slightly lower than the pressure of the compressed air discharged from the air compressor 20, it is not much different from the pressure of the compressed air. In order to send the cooling air that has flowed into the first exhaust port 97 of the outer wing ring 90 into the tail cover 33, in the present embodiment, the compressed air in the intermediate cabin 13 is boosted by the boost compressor 115, and then the pressure is increased. This compressed air is sent into the stationary blade 50a as cooling air.

The cooling air that has flowed into the second passage 82 of the stationary blade 50a exchanges heat with the stationary blade 50a in the process of flowing through the second passage 82 of the stationary blade 50a, and is heated while cooling the stationary blade 50a. .. The cooling air heated in the process of flowing through the second passage 82 flows into the second exhaust port 98 of the outer wing ring 90. The cooling air that has flowed into the second exhaust port 98 flows into the rotor cooling air passage 109 of the inner casing 16 via the rotor cooling pipe 113. The cooling air is cooled by the cooler 116 in the process of flowing through the rotor cooling pipe 113. The cooling air cooled by the cooler 116 flows from the rotor cooling pipe 113 into the rotor cooling air passage 109 of the inner casing 16.

The amount of heat exchanged with the stationary blade 50a in the process of the cooling air flowing through the first passage 81 of the stationary blade 50a is compared with the amount of heat exchanged with the stationary blade 50a in the process of the cooling air flowing through the second passage 82 of the stationary blade 50a. And much less. Therefore, the temperature of the cooling air flowing into the first exhaust port 97 through the first passage 81 of the stationary blade 50a is several tens of degrees higher than the temperature of the compressed air immediately after being discharged from the air compressor 20 of 200 ° C. It is a high degree. Therefore, even if the cooling air flowing into the first exhaust port 97 is sent to the tail cover 33 without being separately cooled, the tail cover 33 can be effectively cooled. On the other hand, the amount of heat exchanged with the stationary blade 50a in the process of the cooling air flowing through the second passage 82 of the stationary blade 50a is large, and the temperature of the air flowing into the second exhaust port 98 through the second passage 82 is 500 ° C. to It is a high temperature of 600 ° C. Therefore, the cooling air flowing into the second exhaust port 98 is cooled by the cooler 116 and then sent to the rotor cooling air passage 109 of the inner casing 16. The temperature of the cooling air cooled by the cooler 116 is, for example, about 200 ° C.

The cooling air that has flowed into the rotor cooling air passage 109 of the inner casing 16 passes through the rotor blade cooling air passage 41p of the turbine rotor 41 and is inside the cooling air passage 43p of the plurality of rotor blades 43a constituting the first stage rotor blade row 43. Inflow to. This cooling air cools the moving blade 43a in the process of flowing in the cooling air passage 43p of the moving blade 43a. This cooling air flows out from the cooling air passage 43p to the combustion gas passage 49. A part of the cooling air flowing out to the combustion gas flow path 49 cools the surface of the moving blade 43a with a film.

Next, the flow of cooling air in the plurality of stationary blades 50a constituting the first-stage stationary blade row 46a will be described in detail with reference to FIGS. 3 to 10.

As shown in FIGS. 3 and 10, the cooling air from the stationary blade cooling air passage 101 of the inner wing ring 100 flows into the space of the inner Dri in the radial direction from the inner impinge plate 68i in the recess 64i of the inner shroud 60i. .. As described above, this space forms a part of the first passage 81 and a part of the second passage 82 of the stationary blade 50a. A part of the cooling air that has flowed into the space of the inner Dri in the radial direction from the inner impinge plate 68i in the recess 64i of the inner shroud 60i passes through the first wing passage 75 as shown in FIGS. 3, 9 and 10. It flows into the first insert 70a extending in the radial direction Dr. As shown in FIGS. 3 and 7, a part of the cooling air flowing into the first insert 70a passes through a large number of through holes 71t formed in the first insert 70a and is on the front side outside the first insert 70a. It flows into the gap passage 83. The cooling air that has flowed into the front gap passage 83 collides with the inner surface 75i of the first blade passage to impinge-cool the portion on the front edge side of the blade body 51. This cooling air convection-cools the inner surface 75i of the first wing passage in the process of flowing in the front gap passage 83 to the outer Dro in the radial direction. As shown in FIGS. 3 and 5, the cooling air flows into the impinge air retention space 66 from the front gap passage 83. The cooling air that has flowed into the impinge air retention space 66 flows into the outer cavity 65o through a large number of through holes 68t of the outer impinging plate 68o, as shown in FIGS. 3 and 6. The cooling air flowing into the outer cavity 65o collides with the anti-gas path surface 62oo of the outer shroud main body 61o, and impinge-cools the outer shroud main body 61o. The cooling air obtained by impingi-cooling the outer shroud main body 61o flows out to the combustion gas flow path 49 from a large number of ejection holes 60e formed in the outer shroud 60o. A part of the cooling air flowing out to the combustion gas flow path 49 cools the gas path surface 62op of the outer shroud 60o and the like with a film.

The rest of the cooling air that has flowed into the first insert 70a flows into the front space 67s of the outer shroud 60o, as shown in FIGS. 3 and 4. The cooling air that has flowed into the front space 67s flows into the first exhaust port 97 of the outer wing ring 90. As described above, since the first passage 81 is a passage that substantially linearly extends in the radial direction Dr and penetrates the stationary blade 50a from the radial inner Dri to the radial outer Dro, the inside of the first passage 81. The amount of heat exchanged with the stationary blade 50a in the process of flowing cooling air is small, and the temperature of the cooling air immediately after flowing into the first passage 81 and the cooling flowing out of the first passage 81 and flowing into the first exhaust port 97. The temperature difference from the air temperature is small.

The other part of the cooling air that has flowed into the space of the inner Dri in the radial direction of the inner impinge plate 68i in the recess 64i of the inner shroud 60i is the inner impinging plate 68i as shown in FIGS. 3, 8 and 9. It flows into the inner cavity 65i forming a part of the second passage 82 through a large number of through holes 68t. The cooling air flowing into the inner cavity 65i collides with the anti-gas path surface 62io of the inner shroud main body 61i, and impinge-cools the inner shroud main body 61i. A part of the cooling air obtained by impingi-cooling the inner shroud main body 61i flows out to the combustion gas flow path 49 from a large number of ejection holes 60e formed in the inner shroud 60i. A part of the cooling air flowing out to the combustion gas flow path 49 cools the gas path surface 62ip of the inner shroud 60i and the like with a film.

As shown in FIGS. 3 and 8, the other part of the cooling air impinge-cooled the inner shroud 60i passes through the communication hole 67ic of the airfoil inner protrusion 67i in the recess 64i of the inner shroud 60i, and the intermediate passage 85. Inflow into. As shown in FIGS. 3 and 7, the cooling air flowing into the intermediate passage 85 passes through a large number of through holes 71t formed in the second insert 70b and enters the intermediate gap passage 84 outside the second insert 70b. Inflow. The cooling air flowing into the intermediate gap passage 84 collides with the inner surface 76i of the second blade passage to impinge-cool the intermediate portion of the blade main body 51. This cooling air flows into the intermediate space 68s of the outer shroud 60o, as shown in FIGS. 3 and 4. The intermediate space 68s forms a part of the second passage 82 of the stationary blade 50a as described above. The cooling air that has flowed into the intermediate space 68s flows into the second exhaust port 98 of the outer wing ring 90. The second passage 82 has a portion extending in a direction having a directional component perpendicular to the radial direction Dr, and the passage length of the second passage 82 is longer than the passage length of the first passage 81. Further, the cooling air that has flowed into the second passage 82 flows into the second exhaust port 98 after impinging cooling the stationary blade 50a at least once. Therefore, a large amount of heat exchanges heat with the stationary blade 50a in the process of the cooling air flowing in the second passage 82, and the temperature of the cooling air immediately after flowing into the second passage 82 and the outflow from the second passage 82. The temperature difference from the temperature of the cooling air flowing into the second exhaust port 98 is large.

The rest of the cooling air that has flowed into the space of the inner Dri in the radial direction of the inner impinge plate 68i in the recess 64i of the inner shroud 60i flows into the third wing passage group 77 as shown in FIGS. 3 and 9. The cooling air that has flowed into the third wing passage group 77 convectally cools the trailing edge side of the wing body 51 in the process of flowing through the third wing passage group 77 toward the radial outer Dro. As shown in FIGS. 3 and 4, this cooling air flows from the third wing passage group 77 into the rear space 69s of the outer shroud 60o. The cooling air that has flowed into the rear space 69s flows out to the combustion gas flow path 49 from the ejection hole 60e formed in the rear peripheral wall 63b of the outer shroud 60o. A part of the cooling air flowing out to the combustion gas flow path 49 cools the gas path surfaces 62ip, 62op and the like of the outer shroud 60o with a film. Further, a part of the cooling air that has flowed into the rear space 69s is used as sealing air between the rear peripheral wall 92b of the blade ring peripheral wall 92 and the rear peripheral wall 63b of the outer shroud 60o.

As described above, in the present embodiment, the temperature difference between the cooling air flowing out from the first passage 81 of the stationary blade 50a and the air flowing out from the second passage 82 is taken into consideration, and each cooling air is used as the gas turbine 10. It is reused as cooling air for the constituent parts. Specifically, since the cooling air flowing out of the first passage 81 is hardly heated in the process of flowing through the first passage 81, the tail is directly passed through the first exhaust port 97 and the first line 111 of the outer wing ring 90. It is sent to the cylinder 33 to cool the tail cylinder 33. Further, since the cooling air flowing out from the second passage 82 is heated in a large amount in the process of flowing through the second passage 82, the turbine rotor 41 passes through the second exhaust port 98 and the second line 112 of the outer wing ring 90. In the process of sending to, it is cooled by the cooler 116. The cooling air cooled by the cooler 116 passes through the second line 112 and the rotor blade cooling air passage 41p of the turbine rotor 41, and is contained in the cooling air passage 43p of the plurality of rotor blades 43a constituting the first stage rotor blade row 43. The rotor blade 43a is cooled.

Therefore, in the present embodiment, the cooling air that has cooled the stationary blade 50a can be effectively used.

"Modification example"
In the above embodiment, the first partition wall 78 and the second partition wall 79 form the space of the outer Dro in the radial direction from the outer sealing plate 69o in the recess 64o of the outer shroud 60o, and the front space 67s, the intermediate space 68s, and the rear space. It is divided into three spaces with the side space 69s. However, as shown in FIG. 11, the second partition wall 79 is not projected outward in the radial direction from the airfoil outer protrusion 67o, and the space of the outer Dro in the radial direction from the outer sealing plate 69o in the recess 64o of the outer shroud 60o. The inside may be divided into two spaces, a front space 67s and a rear space 69sa, by only the first partition wall 78. In this case, the front space 67s communicates with the first exhaust port 97 of the outer wing ring 90, as in the above embodiment. On the other hand, the rear space 69sa communicates with the second exhaust port 98 of the outer wing ring 90.

The step-up compressor 115 of the above embodiment is an external step-up compressor arranged outside the gas turbine 10. However, the step-up compressor may be provided in the gas turbine 10. Specifically, as shown in FIG. 12, the compressor impeller 119 is fixed to the gas turbine rotor 11, and the compressor impeller 119 is arranged in the stationary blade cooling air passage 108a of the inner casing 16a. The stationary blade cooling air passage 108a of the inner casing 16a has a communication opening 108ac that communicates with the intermediate passenger compartment 13 at a position of Dad on the downstream side of the axis of the compressor impeller 119. The step-up compressor 115a in this case includes a gas turbine rotor 11 and a compressor impeller 119 fixed to the gas turbine rotor 11. Further, the vane cooling line 105a in this case is composed of the vane cooling air passage 108a of the inner casing 16. The compressed air that has flowed into the intermediate casing 13 from the diffuser 31 through a large number of through holes 31a flows into the stationary blade cooling air passage 108a as cooling air through the communication opening 108ac of the inner casing 16a. The cooling air that has flowed into the stationary blade cooling air passage 108a is boosted by the step-up compressor 115a and then flows into the stationary blade cooling air passage 101 of the inner wing ring 100 from the stationary blade cooling air passage 108a.

In the above embodiment, the second line 112 communicates with the rotor blade cooling air passage 41p of the turbine rotor 41. However, the second line may communicate with the combustion cylinder 32. In this case, the cooler 116 is not provided in the second line. The high-temperature cooling air in the second exhaust port 98 of the outer wing ring 90 is sent into the combustion cylinder 32 via the second line without being cooled by the cooler 116, and is used as air for combustion of the fuel F. Will be done.

In the above embodiment, the first line 111 communicates with the passage of the tail cover 33. However, the first line may communicate with the rotor blade cooling air passage 41p of the turbine rotor 41. In this case, the cooling air in the first exhaust port 97 of the outer blade ring 90 is sent to the rotor blade 43a via the first line and the rotor blade cooling air passage 41p of the turbine rotor 41 for cooling the rotor blade 43a. Used as air.

Further, in the above embodiment, the first passage 81 and the second passage 82 of the stationary blade 50a communicate with each other. However, the first passage 81 and the second passage 82 do not have to communicate with each other. Further, in the above embodiment, a part of the space in the stationary blade 50a is shared by the first passage 81 and the second passage 82. However, the first passage 81 and the second passage 82 may be independent passages.

10: Gas turbine 11: Gas turbine rotor 13: Intermediate cabin 14: Intermediate casing 15: Gas turbine casing 16, 16a: Inner casing 17: Mounting part 18: Outer cylinder 19: Inner cylinder 20: Air compressor 21: Compressor Rotor 25: Compressor casing 27: Suction port 28: Discharge port 30: Combustor 31: Diffuser 31a: Through hole 32: Combustion cylinder 33: Tail cylinder 34: Pilot burner 35: Main burner 38: Fuel line 39: Support 40: Turbine 41: Turbine rotor 41p: Moving blade cooling air passage 42: Rotor shaft 43: Moving blade row 43a: Moving blade 43b: Blade body 43f: Platform 43p: Cooling air passage 43r: Blade root 45: Turbine casing 46: Static blade row 46a: First stage blade row 48: split ring 49: combustion gas flow path 49i: combustion gas inlet 50, 50a: blade S: blade segment 51: blade body 52: front edge 53: trailing edge 54: Dorsal side surface 55: Ventral side surface 60i: Inner shroud 60o: Outer shroud 60e: Ejection hole 61i: Inner shroud body 61o: Outer shroud body 62ip, 62op: Gas path surface 62io, 62oo: Anti-gas path surface 63i: Inner peripheral wall 63o: Outer peripheral wall 63f : Front peripheral wall 63b: Rear peripheral wall 63p: Ventral peripheral wall 63n: Dorsal peripheral wall 64i, 64o: Recessed 65i: Inner cavity 65o: Outer cavity 66: Impinge air retention space 67s: Front space 68s: Intermediate space 69s, 69sa: Rear side Space 67i: Blade-shaped inner protruding portion 67o: Blade-shaped outer protruding portion 67ic, 67oc: Communication hole 68i: Inner impinging plate 68o: Outer impinging plate 68t: Through hole 69i: Inner sealing plate 69o: Outer sealing plate 70a: No. One insert 70b: Second insert 71a, 71b: Insert body 71t: Through hole 72a, 72b: Seal portion 73b: Lid portion 75: First blade passage 75i: First blade passage inner surface 76: Second blade passage 76i: Second Blade passage inner surface 77: Third blade passage group 78: First partition wall 79, 79a: Second partition wall 81: First passage 82: Second passage 83: Front gap passage 84: Intermediate gap passage 85: Intermediate passage 90: Outer blade ring 91: Blade ring plate 92: Blade ring peripheral wall 92f: Front peripheral wall 92b: Rear peripheral wall 93: First partition wall 94: Second partition wall 95: Casing mounting portion 96: Recess 97: First exhaust port 98: First Two exhaust ports 97o: First exhaust port 98o: Second exhaust port 99: Sealing material 100: Inner blade ring 101: Surface cooling air passage 105: Surface cooling line 106: Surface cooling pipes 108, 108a: Surface blade Cooling sky Air passage 109: Rotor cooling air passage 111: First line 112: Second line 113: Rotor cooling pipe 115, 115a: Pressurizing compressor 116: Cooler 119: Compressor Impeller A: Air F: Fuel G: Combustion gas S: Static wing segment Ar: Axis Da: Axis direction Dau: Axis upstream side Dad: Axis downstream side Dc: Circumferential Dcn: Circumferential dorsal Dcp: Circumferential ventral Dr: Radial Dr: Radial inner Dr: Diameter Direction outside

Claims (18)

A stationary wing with a wing body that extends radially with respect to the axis,
The outer wing ring provided on the radial outer side of the stationary wing and
With
In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. The second passage has a second passage that allows cooling air having a temperature different from the temperature of the cooling air that has flowed out from the first passage to flow outward in the radial direction.
The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port that exhausts to the outside,
The first passage is a passage that penetrates the stationary blade in the radial direction.
Further, the stationary wing has a tubular shape with the first wing passage penetrating the inside of the wing body in the radial direction, and the stationary wing defines the first wing passage with respect to the inner surface of the first wing passage. With a first insert located in the first wing passage at intervals,
The tubular first insert has a plurality of through holes that eject inner cooling air to the inner surface of the first wing passage.
The inside of the tubular first insert forms part of the first passage.
Static wing segment. A stationary wing with a wing body that extends radially with respect to the axis,
The outer wing ring provided on the radial outer side of the stationary wing and
With
In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. The second passage has a second passage that allows cooling air having a temperature different from the temperature of the cooling air that has flowed out from the first passage to flow outward in the radial direction.
The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port that exhausts to the outside,
Further, the stationary blade has an inner shroud provided on the radial inner side of the blade body and an inner impinging plate attached to the inner shroud.
The inner shroud extends from the radial inner end of the wing body in a direction having a directional component perpendicular to the radial direction, and from the peripheral edge along the peripheral edge of the inner shroud body. It has an inner peripheral wall that projects inward in the radial direction.
The inner shroud plate is arranged at intervals in the radial direction with respect to the inner shroud main body, and in cooperation with the inner shroud main body and the inner peripheral wall, the inner shroud plate is radially outer side than the inner shroud plate. Form an inner cavity,
The inner impinge plate is formed with a plurality of through holes for guiding cooling air from the inside in the radial direction of the inner impinging plate to the inner cavity.
The inner cavity forms part of the second passage.
Static wing segment. In the stationary wing segment according to claim 2.
The inner shroud has a plurality of ejection holes for ejecting cooling air in the inner cavity to the outside of the stationary blade.
Static wing segment. In the stationary wing segment according to claim 2 or 3.
The stationary blade has an outer shroud provided on the radial outer side of the blade body and an outer impinging plate attached to the outer shroud.
The outer shroud extends from the radial outer end of the wing body in a direction having a directional component perpendicular to the radial direction, and from the peripheral edge along the peripheral edge of the outer shroud body. It has an outer peripheral wall that projects outward in the radial direction.
The outer impinge plate is arranged at intervals on the outer side in the radial direction with respect to the outer shroud main body, and in cooperation with the outer shroud main body and the outer peripheral wall, is radially inside the outer impinge plate. Form the outer cavity,
The outer impinge plate is formed with a plurality of through holes for guiding cooling air from the outer side in the radial direction of the outer impinging plate to the outer cavity.
The outer cavity forms part of the second passage.
Static wing segment. In the stationary wing segment according to claim 4.
The outer shroud has a plurality of ejection holes for ejecting cooling air in the outer cavity to the outside of the stationary blade.
Static wing segment. In the stationary wing segment according to claim 4 or 5.
The stationary blade has a second blade passage that penetrates the inside of the blade body in the radial direction and guides cooling air from the inner cavity to the outer side in the radial direction with respect to the outer impinge plate.
The second wing passage forms a part of the second passage.
Static wing segment. In the stationary wing segment according to claim 6.
The stationary wing has a tubular shape, and the second insert is arranged in the second wing passage at intervals from the inner surface of the second wing passage that defines the second wing passage with the stationary wing. Have,
The tubular second insert has a plurality of through holes that eject inner cooling air to the inner surface of the second wing passage.
Static wing segment. A stationary wing with a wing body that extends radially with respect to the axis,
The outer wing ring provided on the radial outer side of the stationary wing and
With
In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. The second passage has a second passage that allows cooling air having a temperature different from the temperature of the cooling air that has flowed out from the first passage to flow outward in the radial direction.
The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port that exhausts to the outside,
Further, the stationary wing has a tubular shape with the first wing passage penetrating the inside of the wing body in the radial direction, and the stationary wing defines the first wing passage with respect to the inner surface of the first wing passage. Having first inserts located in the first wing aisle at intervals,
The tubular first insert has a plurality of through holes that eject inner cooling air to the inner surface of the first wing passage.
The inside of the tubular first insert forms part of the first passage, and the outside of the first insert forms part of the second passage.
Static wing segment. In the stationary wing segment according to any one of claims 1 to 8.
The passage length of the second passage is longer than the passage length of the first passage.
Static wing segment. In the stationary wing segment according to any one of claims 1 to 9.
The first passage is a passage that penetrates the stationary blade in the radial direction.
The second passage includes a portion extending in a direction having a directional component perpendicular to the radial direction.
Static wing segment. In the stationary wing segment according to any one of claims 1 to 10.
The stationary wing has a communication passage that communicates with the second passage and the first passage.
Static wing segment. A turbine having a stationary blade segment according to any one of claims 1 to 11.
An air compressor that compresses air to generate compressed air,
A combustor that burns fuel in the compressed air to generate combustion gas and guides the combustion gas into the turbine.
With
The turbine further includes a turbine rotor that rotates about the axis, and a turbine casing that covers the outer peripheral side of the turbine rotor and has the stationary blade segment attached to the inner peripheral side.
The turbine rotor includes a rotor shaft extending in the axial direction with the axis as the center, and a moving blade arranged at a different position in the axial direction with respect to the stationary blade and fixed to the rotor shaft. Have, have
gas turbine. The gas turbine according to claim 12 and
A stationary blade cooling line that discharges from the air compressor and guides the compressed air before being used for combustion of the fuel in the combustor as the cooling air from the radial inside of the stationary blade into the stationary blade.
A step-up compressor arranged in the stationary blade cooling line, which boosts the air flowing into the stationary blade cooling line and sends it to the stationary blade.
Among the high-temperature parts in contact with the combustion gas in the gas turbine, the first high-temperature parts other than the stationary blade and the first exhaust port of the outer wing ring are connected, and the cooling in the first exhaust port is performed. The first line that guides air to the first high temperature component,
Among the high temperature parts in contact with the combustion gas in the gas turbine, the second high temperature parts excluding the stationary blade and the first high temperature part are connected to the second exhaust port of the outer wing ring, and the second A second line that guides the cooling air in the exhaust port to the second high temperature component,
Gas turbine equipment equipped with. In the gas turbine equipment according to claim 13.
The step-up compressor is configured to have a compressor impeller fixed to the turbine rotor.
Gas turbine equipment. In the gas turbine equipment according to claim 13 or 14.
The combustor has a tail cover that guides the combustion gas into the turbine.
The first high temperature component includes the tail cover.
Gas turbine equipment. With a gas turbine
With a stationary wing cooling line,
With a step-up compressor
The first line and
The second line and
With
The gas turbine
With a turbine with a stationary wing segment,
An air compressor that compresses air to generate compressed air,
A combustor that burns fuel in the compressed air to generate combustion gas and guides the combustion gas into the turbine.
With
The turbine, in addition, a turbine rotor rotating about an axis, covering the outer periphery of the turbine rotor, a turbine casing, wherein the stator vane segments on the inner circumferential side is attached, has,
The turbine rotor about said axis, having a rotor shaft extending in the axial direction, and blades are fixed to the rotor shaft, and
The stationary wing segment
A stationary blade that is arranged at a different position in the axial direction with respect to the moving blade and has a blade body extending in the radial direction with respect to the axial direction.
The outer wing ring provided on the radial outer side of the stationary wing and
With
In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. The second passage has a second passage that allows cooling air having a temperature different from the temperature of the cooling air that has flowed out from the first passage to flow outward in the radial direction.
The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port that exhausts to the outside,
The stationary blade cooling line discharges from the air compressor and guides the compressed air before being used for combustion of the fuel in the combustor into the stationary blade from the radial inside of the stationary blade as the cooling air. ,
The step-up compressor is arranged in the vane cooling line, boosts the air flowing into the vane cooling line, and sends the air to the vane.
The first line connects the first high temperature component excluding the stationary blade and the first exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine, and the first line. Guide the cooling air in the exhaust port to the first high temperature component and
The second line includes the second high temperature component excluding the stationary blade and the first high temperature component and the second exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine. Connect and guide the cooling air in the second exhaust port to the second high temperature component.
The combustor has a tail cover that guides the combustion gas into the turbine.
The first high temperature component includes the tail cover.
Gas turbine equipment. In the gas turbine equipment according to any one of claims 13 to 16.
A cooler arranged in the second line and cooling the air flowing into the second line is provided.
The second high temperature component includes the turbine rotor.
Gas turbine equipment. With a gas turbine
With a stationary wing cooling line,
With a step-up compressor
The first line and
The second line and
With
The gas turbine
With a turbine with a stationary wing segment,
An air compressor that compresses air to generate compressed air,
A combustor that burns fuel in the compressed air to generate combustion gas and guides the combustion gas into the turbine.
With
The turbine, in addition, a turbine rotor rotating about an axis, covering the outer periphery of the turbine rotor, a turbine casing, wherein the stator vane segments on the inner circumferential side is attached, has,
The turbine rotor about said axis, having a rotor shaft extending in the axial direction, and blades are fixed to the rotor shaft, and
The stationary wing segment
A stationary blade that is arranged at a different position in the axial direction with respect to the moving blade and has a blade body extending in the radial direction with respect to the axial direction.
The outer wing ring provided on the radial outer side of the stationary wing and
With
In the stationary blade, cooling air flows in from the radial inside with respect to the stationary blade, and a first passage for causing the cooling air to flow out to the radial outside and cooling air from the radial inside with respect to the stationary blade flow in. The second passage has a second passage that allows cooling air having a temperature different from the temperature of the cooling air that has flowed out from the first passage to flow outward in the radial direction.
The outer wing ring has a first exhaust port that exhausts the cooling air that has flowed out from the first passage to the outside in the radial direction, and a second exhaust port that exhausts the cooling air that has flowed out from the second passage to the outside in the radial direction. It has a second exhaust port that exhausts to the outside,
The stationary blade cooling line discharges from the air compressor and guides the compressed air before being used for combustion of the fuel in the combustor into the stationary blade from the radial inside of the stationary blade as the cooling air. ,
The step-up compressor is arranged in the vane cooling line, boosts the air flowing into the vane cooling line, and sends the air to the vane.
The first line connects the first high temperature component excluding the stationary blade and the first exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine, and the first line. Guide the cooling air in the exhaust port to the first high temperature component and
The second line includes the second high temperature component excluding the stationary blade and the first high temperature component and the second exhaust port of the outer wing ring among the high temperature components in contact with the combustion gas in the gas turbine. Connect and guide the cooling air in the second exhaust port to the second high temperature component.
A cooler arranged in the second line and cooling the air flowing into the second line is provided.
The second high temperature component includes the turbine rotor.
Gas turbine equipment.

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