JP2011043118A - Cooling structure for turbine, and turbine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To reduce a stress on the trailing edge of a turbine stator blade 23 while sufficiently ensuring the cooling performance of the impingement cooling of a cooling structure 53 for a turbine. <P>SOLUTION: A cooling chamber 55 is partitioned into an inner divided chamber 65 allowing a leakage flow of combustion gas FA on the radial inside and the radial outside and an outer divided chamber 67 suppressing the leakage flow of the combustion gas FA by a partition member 63. The inside of a front insert 73 communicates with the outer divided chamber 67 and is isolated from the inner divided chamber 65, and the inside of a rear insert 77 communicates with the inner divided chamber 65 and outer divided chamber 67. <P>COPYRIGHT: (C)2011,JPO&INPIT

Description

  The present invention relates to a turbine cooling structure that performs impingement cooling on inner wall surfaces of a leading edge side cooling passage and a trailing edge side cooling passage in a turbine stator blade in a turbine stator of a turbine.

  A conventional turbine cooling structure and the like will be described with reference to FIG. In the drawings, “F” indicates the front direction (upstream direction), “R” indicates the rear direction (downstream direction), “In” indicates the radially inner side, and “Out” indicates the radially outer side. .

  As shown in FIG. 5, turbine stationary blades 103 in a turbine stator 101 of a turbine used in a gas turbine engine (not shown) have a hollow structure, and usually increase the mechanical strength and rigidity of the turbine stationary blades 103. Therefore, a reinforcing rib 105 is provided inside the turbine stationary blade 103. Further, during operation of the gas turbine engine, the turbine stationary blade 103 is exposed to a high temperature. Usually, in order to suppress the high temperature of the turbine stationary blade 103, the front edge side in which the interior of the turbine stationary blade 103 is partitioned by the reinforcing rib 105. Impingement cooling is performed on the inner wall surfaces of the cooling channel 107 and the trailing edge side cooling channel 109. A structure (turbine cooling structure) 111 that performs impingement cooling has the following configuration.

  That is, an annular cooling chamber 117 is defined between the inner peripheral surface of the turbine case 113 of the turbine and the outer peripheral surface of the outer band 115 of the turbine stator 101, and this cooling chamber 117 is compressed by the gas turbine engine. It is connected to a cooling air extraction source (not shown) such as a machine, and accommodates cooling air CA extracted from the cooling air extraction source. Further, a pipe-like front insert 119 is inserted into the cooling passage 107 on the leading edge side of the turbine vane 103. The inside of the front insert 119 communicates with the cooling chamber 117, and the front insert 119 is inserted. A plurality of front ejection holes 121 for ejecting the cooling air CA toward the inner wall surface of the leading edge side cooling flow path 107 of the turbine stationary blade 103 are formed. Further, a pipe-like rear insert 123 is inserted into the cooling passage 109 on the trailing edge side of the turbine vane 103. The interior of the rear insert 123 communicates with the cooling chamber 117, and the rear insert 123 A plurality of rear ejection holes 125 for ejecting the cooling air CA toward the inner wall surface of the trailing edge side cooling flow path 109 of the turbine stationary blade 103 are formed.

  Accordingly, when the cooling air CA extracted from the cooling air extraction source flows into the front insert 119 and the rear insert 123 via the cooling chamber 117, the front of the turbine stationary blade 103 is forwarded from the plurality of front ejection holes 121 of the front insert 119. The cooling air CA is ejected toward the inner wall surface of the edge side cooling flow path 107 and the cooling air is directed from the plurality of rear ejection holes 125 of the rear insert 123 toward the inner wall surface of the rear edge side cooling flow path 109 of the turbine stationary blade 103. CA is ejected. Thereby, impingement cooling can be performed on the inner wall surfaces of the leading edge side cooling channel 107 and the trailing edge side cooling channel 109 of the turbine stationary blade 103. At this time, it is known that a leakage flow of the combustion gas FA occurs from the gap between the hook 127 of the turbine case 113 and the outer band 115 to the cooling chamber 111.

  In addition, there exist some which are shown to patent document 1, patent document 2, and patent document 3 as a prior art relevant to this invention.

JP-A-5-240003 JP 2001-254604 A JP 2001-295604 A

  By the way, in order to sufficiently ensure the impingement cooling performance of the turbine cooling structure 111 described above, the pressure of the cooling air CA flowing into the front insert 119 and the rear insert 123 is increased, and the front insert 119 It is necessary to increase the ejection pressure of the front ejection holes 121 and the rear ejection holes 125 of the rear insert 123.

  On the other hand, when the pressure of the cooling air CA flowing into the front insert 119 and the rear insert 123 is increased, the pressure difference between the inside and outside of the turbine stationary blade 103 increases, and the swelling of the turbine stationary blade 103 increases. Therefore, excessive stress is generated at the trailing edge of the turbine stationary blade 103, which is a portion having a very small curvature, and damage such as cracks is likely to occur at the trailing edge of the turbine stationary blade 103.

  That is, there is a problem that it is difficult to prevent the rear edge of the turbine stationary blade 103 from being damaged, such as cracks, while sufficiently securing the impingement cooling performance of the turbine cooling structure 111.

  Therefore, an object of the present invention is to provide a turbine cooling structure having a novel configuration that can solve the above-described problems.

  A first feature of the present invention is used in an inner wall surface of a leading edge side cooling channel and a trailing edge side cooling channel, which is used in a turbine of a gas turbine engine and in which the interior of a turbine stator blade in the turbine stator of the turbine is partitioned by reinforcing ribs. In the turbine cooling structure that performs impingement cooling, the cooling air extraction is defined between an inner peripheral surface of a turbine case of the turbine and an outer peripheral surface of an outer band of the turbine stator, and is connected to a cooling air extraction source. An annular cooling chamber for containing cooling air extracted from a source and an outer band of the turbine stator are provided so as to surround the outer side of the turbine stator. An internal combustion gas (mainstream gas) that allows leakage flow and communicates with the cooling passage on the trailing edge side of the turbine vane. An annular partition member that partitions the split chamber and an outer split chamber that can suppress (prevent) leakage of combustion gas, and is inserted into the cooling passage on the leading edge side of the turbine stationary blade, and the inside is formed in the outer split chamber A pipe-like front insert that is in communication with and cut off from the inner divided chamber and has a plurality of front injection holes for injecting cooling air toward the inner wall surface of the cooling passage on the leading edge side of the turbine stationary blade And the inside of the turbine stationary blade is inserted into the trailing edge side cooling channel, the inside communicates with the inner divided chamber and the outer divided chamber, and the inner wall surface of the trailing edge side cooling channel of the turbine stationary blade. And a pipe-like rear insert in which a rear ejection hole for ejecting cooling air is formed.

  In the specification and claims of the present application, “provided” means not only directly provided but also indirectly provided via an interposed member such as a bracket. The term “annular” is not limited to the state of being continuous in the circumferential direction, but includes a state in which even if there is a discontinuous portion (divided portion), it can be regarded as a ring as a whole. The “cooling air extraction source” includes the compressor of the gas turbine engine.

  According to the first feature, when the cooling air extracted from the cooling air extraction source flows into the front insert through the outer divided chamber of the cooling chamber during operation of the gas turbine engine, Cooling air is ejected from the plurality of front ejection holes of the insert toward the inner wall surface of the leading edge side cooling passage of the turbine stationary blade. Further, when the cooling air extracted from the cooling air extraction source flows into the rear insert through the outer divided chamber and the inner divided chamber, the turbine stationary blades from the plurality of rear ejection holes of the rear insert. The cooling air is jetted toward the inner wall surface of the trailing edge side cooling channel. Thereby, impingement cooling can be performed on the inner wall surfaces of the leading edge side cooling channel and the trailing edge side cooling channel of the turbine stationary blade.

  Here, the inside of the cooling chamber is partitioned into the inner divided chamber and the outer divided chamber on the radially inner side and the radially outer side by the partition member, and the inside of the front insert communicates with the outer divided chamber and the inner side. Since the interior of the rear insert is in communication with the inner and outer divided chambers, the cooling air flowing into the front insert usually contains low-pressure combustion gas. On the other hand, the cooling air flowing into the rear insert contains low-pressure combustion gas. Accordingly, the pressure of the cooling air flowing into the front insert is increased, while the increase in the pressure of the cooling air flowing into the rear insert is suppressed, and the trailing edge side cooling flow of the turbine stationary blade is suppressed. The difference between the pressure in the road and the pressure outside the turbine stationary blade can be reduced to reduce the swelling of the trailing edge of the turbine stationary blade.

  The gist of the second feature of the present invention is that the turbine driven by the expansion of the combustion gas (mainstream gas) from the combustor of the gas turbine engine includes the turbine cooling structure according to the first feature.

  According to the 2nd characteristic, there exists an effect | action similar to the effect | action by a 1st characteristic.

  According to the present invention, while increasing the pressure of the cooling air flowing into the front insert, the difference between the pressure in the trailing edge side cooling passage of the turbine stationary blade and the pressure outside the turbine stationary blade. , And the swelling of the trailing edge of the turbine stationary blade can be reduced, so that the injection pressure of the front injection hole of the front insert is increased to sufficiently secure the impingement cooling performance of the turbine cooling structure. In addition, it is possible to reduce the stress at the trailing edge of the turbine stationary blade and suppress the occurrence of damage such as cracks at the trailing edge of the turbine stationary blade.

It is sectional drawing which shows a part of turbine which concerns on embodiment of this invention. It is an enlarged view of the arrow I part in FIG. FIG. 3 is an enlarged view along III-III in FIG. 2. FIG. 4A is a contour diagram showing the stress distribution of the trailing edge of the turbine vane as seen from the inside of the trailing edge cooling channel of the turbine vane in the case of the embodiment, and FIG. 4B is a comparative example. FIG. 6 is a diagram showing the stress distribution of the trailing edge of the turbine vane seen from the inside of the trailing edge cooling channel of the turbine vane in the case of. It is a figure explaining the conventional cooling structure for turbines.

  An embodiment of the present invention will be described with reference to FIGS. 1 to 3. In the drawings, “F” indicates the front direction (upstream direction), “R” indicates the rear direction (downstream direction), “In” indicates the radially inner side, and “Out” indicates the radially outer side. .

  As shown in FIGS. 1 and 2, a high-pressure turbine 1 according to an embodiment of the present invention is used for a jet engine (an example of a gas turbine engine), and is a combustion gas (mainstream gas) from a combustor 3. ) A rotational force is obtained by the expansion of the FA to drive a high-pressure compressor (not shown). The high-pressure turbine 1 includes a cylindrical high-pressure turbine case 5, and the high-pressure turbine case 5 extends in the engine axial direction (front-rear direction).

  In the high-pressure turbine case 5, a high-pressure turbine rotor 7 that is rotated by the expansion of the combustion gas FA from the combustor 3 is provided to be rotatable around an engine axis (not shown). The specific configuration of the high-pressure turbine rotor 7 is as follows.

  That is, a turbine disk 9 is provided in the high-pressure turbine case 5 so as to be rotatable around the engine axis, and this turbine disk 9 is integrally connected to a high-pressure compressor rotor (not shown) of the high-pressure compressor. It is. Further, a plurality of mounting grooves 11 are formed at equal intervals in the circumferential direction on the peripheral edge of the turbine disk 9, and turbine blades 13 are fitted into the mounting grooves 11 of the turbine disk 9. ing. Further, a cylindrical front holding member 15 that holds a plurality of turbine rotor blades 13 from the front direction is provided on the front side of the turbine disk 9, and a plurality of turbine rotor blades 13 are provided on the rear side of the turbine disk 9. A cylindrical rear holding member 17 is provided to hold the rear side from behind.

  An annular shroud 19 is provided on the inner peripheral surface of the high-pressure turbine case 5 so as to surround the plurality of turbine rotor blades 13, and the inner peripheral side of the shroud 19 can allow contact with the tips of the turbine rotor blades 13. It is like a honeycomb. Further, the shroud 19 is segmented in the circumferential direction, in other words, includes a plurality of arc-shaped shroud segments.

  An annular high-pressure turbine stator 21 that rectifies the combustion gas FA into an axial flow is provided on the front side of the high-pressure turbine rotor 7 in the high-pressure turbine case 5. The specific configuration of the high-pressure turbine stator 21 is as follows.

  That is, the high-pressure turbine stator 21 is connected to a plurality (only one is shown) of turbine stationary blades 23 arranged at equal intervals in the circumferential direction, and outer end portions (radially outer end portions) of the plurality of turbine stationary blades 23. And an arc-shaped outer band 25 provided to be connected to each other and an arc-shaped inner band 27 provided so as to be connected to the inner ends (radially inner ends) of the plurality of turbine vanes 23. Yes. The high-pressure turbine stator 21 is segmented in the circumferential direction, in other words, a plurality of arc-shaped turbine stator segments.

  An annular front claw 29 is formed on the front side of the outer band 25, and the front claw 29 of the outer band 25 can be locked to an annular front hook 31 formed on the inner peripheral surface of the high-pressure turbine case 5. . An annular rear claw 33 is formed on the rear side of the outer band 25, and the rear claw 33 of the outer band 25 is an annular shape formed behind the front hook 31 on the inner peripheral surface of the high-pressure turbine case 5. The rear hook 35 can be locked. Further, an annular projecting piece 37 is formed on the inner peripheral surface of the inner band 27, and the projecting piece 37 of the inner band 27 is fitted in the peripheral groove 41 of the stator support member 39 fixed in the high-pressure turbine case 5. Is possible.

  As shown in FIGS. 2 and 3, each turbine stationary blade 23 is made of a ceramic matrix composite material and has a hollow structure. Further, a reinforcing rib 43 is provided inside each turbine vane 23, and the inside of each turbine vane 23 is cooled by the reinforcing rib 43 with a leading edge side cooling channel (front cooling channel) 45 and a trailing edge side cooling. A flow path (rear cooling flow path) 47 is partitioned. A plurality of blowout holes 49 for blowing out the cooling air CA are formed in the front edge and the abdominal surface of each turbine stationary blade 23, and a plurality of outlets for discharging the cooling air CA are discharged at the rear edge of each turbine stationary blade 23. The discharge hole 51 is formed.

  The high-pressure turbine 1 includes a turbine cooling structure 53, and the turbine cooling structure 53 impinges the inner wall surfaces of the leading edge side cooling passage 45 and the trailing edge side cooling passage 47 of each turbine vane 23. In addition, film cooling is performed on the surface (including the front edge, the abdominal surface, and the back surface) of each turbine stationary blade 23. And the concrete structure of the cooling structure 53 for turbines which is the principal part of embodiment of this invention is as follows.

  That is, an annular cooling chamber 55 is defined (partitioned) between the inner peripheral surface of the high-pressure turbine case 5 and the outer peripheral surface of the outer band 25. The cooling chamber 55 includes the connection port 57 and the connection pipe. 59 is connected to a cooling air extraction source 61 such as a low-pressure compressor (not shown) or a high-pressure compressor, and accommodates the cooling air CA noted from the cooling air extraction source 61. An annular partition member 63 is provided between the front hook 31 and the rear hook 35 in the high-pressure turbine case 5 so as to surround and press-contact the outer band 25, and the partition member 63 is provided in the cooling chamber 55. Is divided into an inner divided chamber 65 and an outer divided chamber 67 on the radially inner side and the radially outer side. Further, an opening 69 is formed at a position (a position facing) the leading edge side cooling channel 45 of each turbine stationary blade 23 in the partition member 63, and the rear portion of each turbine stationary blade 23 in the partition member 63 is formed. A through hole (communication hole) 71 is formed at a position aligned with the edge side cooling flow path 47.

  Here, as shown in FIG. 2, the inner divided chamber 65 has a lower pressure than the cooling air CA that passes through the gap between the front hook 31 and the outer band 31 and the gap between the partition member 63 and the outer band 31. The combustion gas FA can be allowed to leak, and is communicated with the cooling passage 47 on the trailing edge side of each turbine vane 23. Further, the outer divided chamber 67 is capable of suppressing (blocking) the leakage flow of the combustion gas FA by the partition member 63 being pressed against the front hook 31 and the rear hook 35 in the high-pressure turbine case 5. Since the pressure in the outer divided chamber 67 is higher than the pressure in the inner divided chamber 65, the inflow of the cooling air CA from the inner divided chamber 65 to the outer divided chamber 67 can be suppressed through the through hole 71. It has become.

  A pipe-like front insert 73 is provided at the periphery of each opening 69 of the partition member 63, and the inside of each front insert 73 communicates with the outer divided chamber 67 and is blocked by the inner divided chamber 65. . Each front insert 73 is inserted into the leading edge side cooling channel 45 of each turbine vane 23, and each front insert 73 has an inner wall surface of the leading edge side cooling channel 45 of each turbine vane 23. On the other hand, a plurality of front ejection holes 75 for ejecting the cooling air CA are formed. Here, by inserting each front insert 73 into the leading edge side cooling flow path 45 of each turbine stationary blade 23, the front edge side cooling flow path 45 and the inner divided chamber 65 of each turbine stationary blade 23 are blocked. ing.

  A pipe-like rear insert 77 is provided at a position aligned with each through hole 71 in each outer band 25, and the inside of each outer band 25 is connected to the inner divided chamber 65 and the outer divided chamber 67 with the partition member 63. It communicates via each through hole 71. Each rear insert 77 is inserted into the trailing edge side cooling flow path 47 of each turbine stationary blade 23, and the inner wall surface of the rear edge side cooling flow path 47 of each turbine stationary blade 23 is inserted into each rear insert 77. On the other hand, a plurality of rear ejection holes 79 for ejecting the cooling air CA are formed.

  Then, the effect | action and effect of embodiment of this invention are demonstrated.

  When the jet engine is operated, the plurality of turbine rotor blades 13 and the turbine disk 9 (in other words, the high-pressure turbine rotor 7) can be rotated by the expansion of the combustion gas FA from the combustor 3. On the other hand, the high-pressure compressor rotor can be driven by rotating the high-pressure compressor rotor by the rotation of the high-pressure turbine rotor 7.

  During operation of the jet engine, when the cooling air CA extracted from the cooling air extraction source 61 flows into the front inserts 73 via the outer divided chamber 67 of the cooling chamber 55, a plurality of front jets of the front inserts 73 are supplied. The cooling air CA is ejected from the hole 75 toward the inner wall surface of the cooling passage 45 on the leading edge side of each turbine stationary blade 23. Further, when the cooling air CA extracted from the cooling air extraction source 61 flows into the rear inserts 77 via the outer divided chamber 67 and the inner divided chamber 65, the cooling air CA is extracted from the plurality of rear ejection holes 79 of the rear inserts 77. Cooling air CA is ejected toward the inner wall surface of the trailing edge side cooling passage 47 of the turbine stationary blade 23. Thereby, impingement cooling can be performed on the inner wall surfaces of the leading edge side cooling flow path 45 and the trailing edge side cooling flow path 47 of each turbine stationary blade 23. Further, the cooling air CA ejected from the plurality of front ejection holes 75 of each front insert 73 is ejected from the plurality of ejection holes 49 of each turbine stationary blade 23, thereby covering the surface of each turbine stationary blade 23. Film cooling can be performed on the surface of each turbine vane 23 while generating a film (not shown). Note that the cooling air CA ejected from the plurality of rear ejection holes 79 of each rear insert 77 is exhausted from the plurality of exhaust holes 51 of each turbine stationary blade 23.

  Here, the inside of the cooling chamber 55 is divided into an inner divided chamber 65 and an outer divided chamber 67 on the radially inner side and radially outer side by the partition member 63, and the inside of the front insert 73 communicates with the outer divided chamber 67 and is divided into the inner divided chambers. Since the rear insert 77 is cut off from the chamber 65 and communicates with the inner divided chamber 65 and the outer divided chamber 67, the cooling air CA that flows into the front insert 73 is usually low pressure combustion gas. While FA is not included, the cooling air CA that flows into the rear inserts 77 includes a low-pressure combustion gas FA. Thus, the pressure of the cooling air CA flowing into the front inserts 73 is increased, while the increase in the pressure of the cooling air CA flowing into the rear inserts 77 is suppressed, so that the turbine stationary blades 23 By reducing the difference between the pressure in the trailing edge side cooling flow path 47 and the pressure outside the turbine stationary blades 23, the swelling of the trailing edge of each turbine stationary blade 23 can be reduced.

  Therefore, according to the embodiment of the present invention, while increasing the pressure of the cooling air CA flowing into each front insert 73, the pressure in the trailing edge side cooling passage 47 of each turbine stationary blade 23 and each turbine static Since the difference between the pressure outside the blades 23 can be reduced and the bulge of the trailing edge of each turbine vane 23 can be reduced, the jet pressure of the front jet holes 75 of each front insert 73 is increased, and the turbine cooling structure 53, after ensuring sufficient cooling performance for impingement cooling and film cooling, the stress at the trailing edge of each turbine vane 23 is reduced, and damage such as cracks occurs at the trailing edge of each turbine vane 23. Can be suppressed.

  In addition, this invention is not restricted to description of the above-mentioned embodiment, In addition, it can implement in a various aspect. Further, the scope of rights encompassed by the present invention is not limited to these embodiments.

  An embodiment of the present invention will be described with reference to FIGS.

  When the low-pressure combustion gas is contained in the cooling air flowing into the rear insert (in the embodiment), and when the low-pressure combustion gas is not contained in the cooling air flowing into the rear insert A structural analysis of the hoop direction stress distribution at the trailing edge of the turbine stationary blade in the case of the comparative example is as shown in FIGS. 4 (a) and 4 (b).

  That is, in the case of the example, it was confirmed that the stress at the trailing edge of the turbine blade, particularly the stress near the ejection hole at the trailing edge of the turbine blade can be reduced as compared with the comparative example. More specifically, the stress in the vicinity of the ejection hole at the trailing edge of the turbine blade in the case of the comparative example is 41.4 MPa, whereas the ejection hole at the trailing edge of the turbine blade in the example. The stress in the vicinity was 36.4 MPa, and it was confirmed that the stress in the vicinity of the ejection hole at the trailing edge of the turbine blade could be reduced by 12.1%.

CA Cooling air FA Combustion gas (Mainstream gas)
DESCRIPTION OF SYMBOLS 1 High pressure turbine 3 Combustor 5 High pressure turbine case 7 High pressure turbine rotor 9 Turbine disk 13 Turbine blade 21 High pressure turbine stator 23 Turbine stationary blade 25 Outer band 27 Inner band 29 Front claw 31 Front hook 33 Rear claw 35 Rear hook 43 Reinforcement rib 45 Front edge side cooling flow path 47 Rear edge side cooling flow path 49 Blowout hole 51 Discharge hole 53 Cooling structure for turbine 55 Cooling chamber 61 Cooling air extraction source 63 Partition member 65 Inner divided chamber 67 Outer divided chamber 69 Opening 71 Through hole 73 Front insert 75 Front ejection hole 77 Rear insert 79 Rear ejection hole

Claims (5)

Cooling for turbines used in a turbine of a gas turbine engine and performing impingement cooling on the inner wall surfaces of the leading edge side cooling flow path and the trailing edge side cooling flow path in which the inside of the turbine stationary blade in the turbine stator of the turbine is partitioned by reinforcing ribs In structure
Annular cooling that is defined between the inner peripheral surface of the turbine case of the turbine and the outer peripheral surface of the outer band of the turbine stator, is connected to a cooling air extraction source, and stores cooling air extracted from the cooling air extraction source A chamber;
The turbine stator is provided so as to surround the outer band, and is capable of allowing a leakage flow of combustion gas having a pressure lower than that of cooling air inside the cooling chamber on the radially inner side and the radially outer side. An inner partition chamber communicating with the trailing edge side cooling flow path and an annular partition member partitioning into an outer partition chamber capable of suppressing a leakage flow of combustion gas;
The turbine stationary blade is inserted into the leading edge side cooling flow path, and the inside communicates with the outer divided chamber and is cut off from the inner divided chamber. A pipe-shaped front insert formed with a plurality of front ejection holes for ejecting cooling air toward the wall surface;
The turbine stationary blade is inserted into the trailing edge side cooling channel, and the inside communicates with the inner divided chamber and the outer divided chamber, and faces the inner wall surface of the trailing edge side cooling channel of the turbine stationary blade. And a pipe-like rear insert in which a rear ejection hole for ejecting cooling air is formed.   The front insert is inserted into the leading edge side cooling flow path of the turbine stationary blade, whereby the leading edge side cooling flow path of the turbine stationary blade and the inner divided chamber are shut off. The turbine cooling structure according to claim 1.   The opening part is formed in the partition member in a position that matches the leading edge side cooling flow path of the turbine stationary blade, and the front insert is provided on the periphery of the opening part of the partition member. The cooling structure for turbines of Claim 1 or Claim 2.   A through hole is formed in the partition member at a position aligned with the cooling passage on the trailing edge side of the turbine stationary blade, and the rear insert has the through hole of the partition member in the outer divided chamber and the inner divided chamber. The turbine cooling structure according to claim 1, wherein the turbine cooling structure is communicated with each other. In a turbine that obtains rotational power by expansion of combustion gas from a combustor of a gas turbine engine,
A turbine comprising the turbine cooling structure according to any one of claims 1 to 4.

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