JP5134570B2 - Turbine cooling structure and gas turbine - Google Patents

Description

  The present invention relates to a turbine cooling structure and a gas turbine.

In recent years, in gas turbines, the temperature of combustion gas led to turbine stationary blades and turbine rotor blades tends to increase in order to achieve higher output and higher efficiency.
Therefore, a cooling medium such as air or steam is supplied to the cooling flow passages inside the turbine stator blade and the turbine rotor blade for the purpose of ensuring heat resistance in the turbine stator blade and the turbine rotor blade.

  As described above, in order to supply a cooling medium, for example, cooling air, to the cooling flow path of the turbine blade, the cooling air is introduced into the rotating rotor from the air storage chamber provided on the stationary casing side. Has been.

  Simply introducing cooling air from the stationary air reservoir into the rotating rotor causes energy loss (discs) when cooling air that does not have a flow velocity component in the rotor rotation direction flows into the rotor. Pumping loss) occurred. Furthermore, an inlet pressure loss has occurred at the inlet where the cooling air flows into the rotor.

  In order to solve the above-mentioned problem, a pre-swirl nozzle that gives a flow velocity component in the rotation direction of the rotor to the cooling air flowing into the rotor rotating from the stationary reservoir, or A technique of providing a TOBI (Tangential OnBoard Injection) nozzle has been proposed (see, for example, Patent Documents 1 to 4).

  As described above, by giving a flow velocity component in the rotation direction of the rotor to the cooling air by the pre-swirl nozzle or the TOBI nozzle, by reducing the difference between the flow velocity component in the rotation direction of the cooling air and the rotation speed of the rotor, The occurrence of disk pumping loss and the like could be suppressed.

Japanese Patent No. 4088557 Japanese Patent No. 4146257 JP 2007-298020 A Special table 2003-500585 gazette

  In the techniques described in Patent Documents 1 to 4 described above, the cooling air given the pre-turn by the pre-turn nozzle or the TOBI nozzle is used for cooling the turbine blade, the first stage stationary blade (1C), and the first stage blade (1S). Used for rim purge between

  As described above, cooling air that has been given a pre-swirl is used for cooling turbine blades, and cooling that is not given a pre-swirl that has flowed in from a path other than the pre-swivel nozzle, that is, from a seal portion or the like. Air was also used.

  As described above, when the cooling air not given the pre-swirl is mixed with the cooling air given the pre-swirl, the flow velocity component in the rotation direction in the cooling air is decelerated. Then, the difference between the flow velocity component in the rotation direction of the cooling air and the rotation speed of the rotor becomes large, and there is a problem that the occurrence of disk pumping loss or the like cannot be suppressed.

  The present invention has been made to solve the above-described problems, and an object of the present invention is to provide a turbine cooling structure and a gas turbine capable of suppressing a decrease in performance.

In order to achieve the above object, the present invention provides the following means.
The turbine cooling structure according to the present invention includes a turbine disk that has a plurality of rotor blades arranged side by side in the circumferential direction on the outer periphery, and is rotatably supported around an axis, and the inside of the turbine disk is directed from the inside to the outside. And a plurality of cooling holes that are arranged side by side in the circumferential direction and supply a cooling medium to a cooling flow path inside the rotor blade, a first inlet space to which the cooling medium is supplied from the outside, and the plurality of the cooling holes A partition portion that divides the second inlet space communicated with the cooling hole, and a flow that extends inside the partition portion from the first inlet space toward the second inlet space so as to be inclined in the rotational direction of the turbine disk. A nozzle section having a passage; and an axially adjacent casing that is adjacent to the turbine disk in the axial direction and accommodates the plurality of rotor blades. A stationary blade that is mounted alternately with the moving blades, a rim seal space that communicates with a gap between the moving blades, and the interior of the partition portion, bypassing the nozzle portion, the first inlet space and the second inlet space. And a communication channel extending toward the rim seal space from a detour space connecting the two.

  According to the present invention, the cooling medium that has flowed into the bypass space from the first inlet space flows into the rim seal space via the communication flow path, and flows out from the gap between the moving blade and the stationary blade, in other words, the cooling medium is Used for rim seals. Therefore, inflow of the cooling medium from the bypass space to the second inlet space is suppressed, and deceleration of the flow velocity component in the rotation direction in the cooling medium flowing out from the nozzle portion can be suppressed.

  In the above-described invention, it is desirable that a plug portion that has a throttle for adjusting the flow rate of the cooling medium flowing through the communication channel and is detachably attached to the communication channel is provided.

  According to the present invention, the flow rate of the cooling medium flowing into the rim seal space can be changed by replacing the plug portion attached to the communication channel with another plug portion having a different throttle diameter. Therefore, for example, in the gas turbine having the turbine cooling structure of the present invention, even when the modification is accompanied by performance improvement and the flow rate of the cooling medium used for the rim seal needs to be changed, the plug portion is By exchanging, the flow rate of the cooling medium used for the rim seal can be easily changed.

  The gas turbine of the present invention includes a compression unit that sucks and compresses air, a combustion unit that generates a combustion gas by burning a mixture of compressed air and fuel supplied from the outside, and generated combustion And a turbine section for extracting a rotational driving force from the gas, and the turbine section is provided with the turbine cooling structure according to the present invention.

  According to the present invention, since the turbine cooling structure of the present invention is provided, it is possible to suppress performance degradation in the gas turbine.

  According to the turbine cooling structure and the gas turbine of the present invention, the communication medium flows out the cooling medium flowing into the bypass space from the first inlet space through the gap between the moving blade and the stationary blade through the rim seal space. There is an effect that a decrease in performance can be suppressed.

It is a mimetic diagram explaining the composition of the gas turbine concerning one embodiment of the present invention. It is a fragmentary sectional view explaining the structure in the turbine part of FIG. FIG. 3 is a partially enlarged cross-sectional view for explaining a configuration around a first stage moving blade in FIG. 2. FIG. 4 is a cross-sectional view taken along line AA for explaining the configuration around the first stage moving blade of FIG. 3. It is a partial expanded sectional view explaining the structure of the inner side partition part periphery of FIG.

A gas turbine according to an embodiment of the present invention will be described with reference to FIGS.
FIG. 1 is a schematic diagram illustrating the configuration of a gas turbine according to the present embodiment.
In the present embodiment, the gas turbine 1 of the present application will be described as applied to an application used to drive a generator, but may be used for other applications and is not particularly limited.
The gas turbine 1 is mainly provided with a compressor (compression unit) 2, a combustor (combustion unit) 3, a turbine unit 4, and a rotating shaft 5.

As shown in FIG. 1, the compressor 2 sucks and compresses atmospheric air, which is external air, and supplies the compressed air to the combustor 3.
In addition, as a compressor 2, a well-known structure can be used and the structure in particular is not limited.

As shown in FIG. 1, the combustor 3 mixes the air compressed by the compressor 2 and the fuel supplied from the outside, and burns the mixed gas mixture, thereby producing a high-temperature gas (combustion gas). Is generated.
In addition, as a combustor 3, a well-known structure can be used and the structure in particular is not limited.

  As shown in FIG. 1, the turbine unit 4 receives the supply of the high-temperature gas generated by the combustor 3 to generate a rotational driving force, and transmits the generated rotational driving force to the rotating shaft 5.

  FIG. 2 is a partial cross-sectional view for explaining the configuration in the turbine section of FIG. 1. FIG. 3 is a partially enlarged cross-sectional view for explaining the configuration around the first stage moving blade in FIG. 2. FIG. 4 is a cross-sectional view taken along line AA for explaining the configuration around the first stage moving blade in FIG. 3. FIG. 5 is a partially enlarged cross-sectional view for explaining the configuration around the inner partition wall of FIG.

  As shown in FIGS. 2 to 5, the turbine section 4 includes a first stage stationary blade (stationary blade) 11 </ b> S fixed to the casing 13 </ b> S and a first stage blade (moving blade) attached to a first stage turbine disk (turbine disk) 31 </ b> R. Blades) 21R, inlet cavities (first inlet spaces) 41C and outlet cavities (second inlet spaces) 42C to which cooling air (cooling medium) used for cooling the first stage blades 21R and the like is supplied, and inlet cavities 41C And an inner partition wall (partition) 61S that divides the outlet cavity 42C, and a rim seal cavity (rim seal space) 43C that supplies cooling air between the first-stage stationary blade 11S and the first-stage moving blade 21R are mainly provided. Yes.

  Here, for easy understanding, only the vicinity of the first stage stationary blade 11S and the first stage stationary blade 21R that are arranged near the combustor 3 and need the most cooling will be described. However, the turbine section 4 includes the first stage stationary blade 11S and In addition to the first stage blade 21R, a second stage blade, a second stage blade, a third stage blade, a third stage blade, and the like are provided as necessary along the rotating shaft 5 toward the downstream side of the combustion gas. .

As shown in FIGS. 2 and 3, the first stage stationary blade 11 </ b> S generates a rotational driving force from the combustion gas generated by the combustor 3 together with the first stage moving blade 21 </ b> R.
The first stage stationary blade 11S extends from the cylindrical casing 13S to the radially inner side (lower side in FIG. 2) around the rotation shaft 5, in other words, extends to the flow path side through which the combustion gas flows. It is the wing | blade arrange | positioned along with the circumferential direction at equal intervals.
Further, the first stage stationary blade 11S is arranged on the combustor 3 side (left side in FIG. 2) from the first stage moving blade 21R, in other words, on the upstream side of the combustion gas flow.

An inner shroud 12S is provided at an inner peripheral end of the first stage stationary blade 11S.
The inner shroud 12S is a plate-like member extending in the circumferential direction from the end of the first stage stationary blade 11S, and forms a part of the flow path through which the combustion gas flows. Further, the inner shroud 12S is fixed to the outer partition wall 51S.

  The outer partition wall 51S is a member formed in a cylindrical shape with the rotating shaft 5 as the center, and constitutes part of the wall surface of the inlet cavity 41C, the rim seal cavity 43C, and the like together with the inner partition wall 61S and the like. .

As shown in FIGS. 2 and 3, the first stage blade 21 </ b> R generates rotational driving force from the combustion gas generated by the combustor 3 together with the first stage stationary blade 11 </ b> S.
The first stage blade 21R extends from the outer peripheral surface of the first stage turbine disk 31R formed in a disk shape to the radially outer side (upper side in FIG. 2), in other words, the flow path side through which the combustion gas flows. The wings are arranged at equal intervals in the circumferential direction.
Further, the first stage moving blade 21R is arranged on the downstream side (right side in FIG. 2) of the combustion gas flow with respect to the first stage stationary blade 11S.

  As shown in FIGS. 3 and 4, the first stage blade 21 </ b> R is mainly provided with a blade portion 22 </ b> R, a platform 23 </ b> R, a blade attachment portion 24 </ b> R, and a cooling channel 25 </ b> R.

  As shown in FIGS. 3 and 4, the wing portion 22 </ b> R is a portion that receives a flow of combustion gas and generates a rotational driving force. Further, the blade portion 22R includes a negative pressure surface curved in a convex shape and a positive pressure surface curved in a concave shape in a sectional view along the circumferential direction, and a front edge and an end portion on the downstream side that are upstream ends in the combustion gas flow. It is a part composed of a certain trailing edge.

  As shown in FIGS. 3 and 4, the platform 23 </ b> R is a portion disposed at the radially inner end (lower side in FIG. 3) of the wing 22 </ b> R, in other words, between the wing 22 </ b> R and the wing attachment 24 </ b> R. It is a part arranged in. The platform 23R extends in the radial direction and forms part of a flow path through which combustion gas flows.

As shown in FIGS. 3 and 4, the blade attachment portion 24 </ b> R is a portion that is fitted to the first stage turbine disk 31 </ b> R, in other words, a portion that is disposed at the radially inner end of the first stage blade 21 </ b> R. . Further, the blade mounting portion 24R is provided with a portion that is formed in a so-called Christmas tree shape in a cross-sectional view along the radial direction and is fitted to the first stage turbine disk 31R.
On the other hand, the radially inner end of the blade attachment portion 24R constitutes a part of the wall surface of an axial communication path 36R described later.

The cooling flow path 25R is a flow path that guides cooling air to the inside of the wing part 22R and cools the wing part 22R. It is a flow path that is open to a surface constituting the wall surface of the axial communication path 36R in the blade attachment portion 24R and extends from the blade attachment portion 24R to the blade portion 22R.
Note that the configuration of the cooling flow path 25R inside the blade portion 22R can be a known configuration, and is not particularly limited.

As shown in FIGS. 2 and 3, the first stage turbine disk 31 </ b> R is a member formed in a disk shape, and is disposed so as to be able to transmit a rotational driving force to the rotary shaft 5. On the outer peripheral surface of the first stage turbine disk 31R, the first stage moving blades 21R are arranged at equal intervals in the circumferential direction.
On the other hand, the first stage turbine disk 31R constitutes a part of the wall surface of the rim seal cavity 43C together with the outer partition wall portion 51S and the inner partition wall portion 61S.

  As shown in FIGS. 3 and 4, the first-stage turbine disk 31R includes a fitting groove 32R, a first cooling hole (cooling hole) 33R, a second cooling hole (cooling hole) 34R, and a circumferential communication path 35R. In addition, an axial communication path 36R and a seal arm 37R are provided.

As shown in FIGS. 3 and 4, the fitting groove 32 </ b> R is a portion into which the blade attachment portion 24 </ b> R of the first stage moving blade 21 </ b> R is fitted. Therefore, the fitting groove 32R is formed in a shape that meshes with a so-called Christmas tree shape in the blade attachment portion 24R.
Further, the fitting grooves 32 </ b> R are formed on the outer peripheral surface of the first stage turbine disk 31 </ b> R so as to be arranged at equal intervals, and are formed so as to extend along the rotation shaft 5.
On the other hand, the radially inner end of the fitting groove 32R is connected to the groove constituting the axial communication path 36R.

As shown in FIGS. 2 to 4, the first cooling hole 33 </ b> R is a hole that guides cooling air from the outlet cavity 42 </ b> C to the cooling passage 25 </ b> R of the first stage blade 21 </ b> R together with the second cooling hole 34 </ b> R. It constitutes an air flow path.
The first cooling holes 33R are holes extending in the radial direction inside the first stage turbine disk 31R, and are arranged at equal intervals in the circumferential direction, like the first stage moving blades 21R. In other words, the first cooling hole 33R is disposed at the same position as the first stage blade 21R in the circumferential direction.
Further, one end of the first cooling hole 33R is opened to a surface constituting the wall surface of the outlet cavity 42C in the first stage turbine disk 31R, and the other end is an axial communication path 36R in the first stage turbine disk 31R. It is opened in the surface which comprises the wall surface.

As shown in FIGS. 2 to 4, the second cooling hole 34 </ b> R is a hole that guides cooling air from the outlet cavity 42 </ b> C to the cooling passage 25 </ b> R of the first stage blade 21 </ b> R together with the first cooling hole 33 </ b> R. It constitutes an air flow path.
The second cooling hole 34R is a hole extending in the radial direction inside the first stage turbine disk 31R, and is disposed between the first cooling holes 33R.
Further, one end of the second cooling hole 34R is opened to a surface constituting the wall surface of the outlet cavity 42C in the first stage turbine disk 31R, and the other end is a circumferential communication path 35R in the first stage turbine disk 31R. It is opened in the surface which comprises the wall surface.

As shown in FIGS. 3 and 4, the circumferential communication path 35R is a flow path that guides the cooling air that has flowed through the second cooling hole 34R to the axial communication path 36R.
The circumferential communication passage 35R is a flow passage formed in an annular shape extending in the circumferential direction at a radially outer position in the first stage turbine disk 31R, and is a flow connecting the first cooling hole 33R and the second cooling hole 34R. Road.

As shown in FIGS. 3 and 4, the axial communication path 36 </ b> R extends along the rotation shaft 5, and the cooling air flowing through the first cooling hole 33 </ b> R and the second cooling hole 34 </ b> R is supplied to the cooling flow of the first stage blade 21 </ b> R. This is a flow path leading to the path 25R.
The axial communication passage 36R is a passage disposed at a radially outer position in the first stage turbine disk 31R, and includes a groove formed in the first stage turbine disk 31R, and a blade attachment part 24R of the first stage rotor blade 21R. It is a passage surrounded by.

2 and 3, the seal arm 37R extends in a cylindrical shape from the first stage turbine disk 31R toward the inner partition wall 61S, and regulates the flow rate of the cooling air flowing from the outlet cavity 42C to the rim seal cavity 43C. It constitutes a seal structure.
Further, the seal arm 37R constitutes a part of the wall surface of the outlet cavity 42C.

  As shown in FIGS. 2 and 5, the inlet cavity 41C is a cylindrical space to which cooling air is supplied, and supplies the cooling air to the outlet cavity 42C via a TOBI nozzle 62S described later.

The inlet cavity 41C is a space formed between the outer partition wall portion 51S and the inner partition wall portion 61S, and is upstream of the combustion gas flow (left side in FIG. 2) with respect to the outlet cavity 42C and the rim seal cavity 43C. Is formed. Furthermore, the inlet cavity 41C is disposed radially outward (upper side in FIG. 2) with respect to the outlet cavity 42C, and is disposed radially inward (lower side of FIG. 2) with respect to the rim seal cavity 43C. ing.
On the other hand, the inlet cavity 41C is connected to a supply flow path 52S for supplying cooling air, and is also connected to a TOBI nozzle 62S and a partition wall flow path 63S described later.

  Here, the cooling air supplied to the inlet cavity 41 </ b> C is air extracted from the compressor 2 and cooled. More specifically, compressed air having a pressure required to supply cooling air from the inlet cavity 41C to the cooling flow path 25R of the first stage blade 21R, or cooling from the rim seal cavity 43C to the combustion gas flow path. Compressed air having a pressure required for causing air to flow out is extracted from the compressor 2.

  As shown in FIGS. 2 and 5, the exit cavity 42C is a cylindrical space to which cooling air is supplied, and supplies a part of the cooling air to the first cooling hole 33R and the second cooling hole 34R. is there.

  The exit cavity 42C is a space formed between the inner partition wall 61S, the connection rotor 38R, and the first stage turbine disk 31R, and is formed on the downstream side (right side in FIG. 2) of the inlet cavity 41C in the combustion gas flow. It is disposed at substantially the same position as the rim seal cavity 43C. Furthermore, the outlet cavity 42C is disposed radially inward (lower side in FIG. 2) with respect to the inlet cavity 41C and the rim seal cavity.

  On the other hand, the outlet cavity 42C is connected to a later-described TOBI nozzle 62S and a space (bypassing space) 44C between the inner partition wall portion 61S and the connection rotor 38R, as well as the first cooling hole 33R and the second cooling hole. 34R, the rim seal cavity 43C, and the disk flow path 39R are also connected.

  As shown in FIGS. 2, 3, and 5, the inner partition wall portion 61S is disposed with a space inside the outer partition wall portion 51S and is disposed with a space outside the connection rotor 38R. It is a member. The inner partition wall 61S forms an inlet cavity 41C together with the outer partition wall 51S, forms an outlet cavity 42C together with the connecting rotor 38R and the first stage turbine disk 31R, and further rims together with the outer partition wall 51S and the first stage turbine disk 31R. The seal cavity 43C is formed.

  Further, the inner partition wall 61S is provided with a cylindrical portion extending along the upstream rotary shaft 5 in the combustion gas flow and a flange-like portion extending in a ring plate shape in the downstream radial direction.

  The inner partition wall portion 61S includes a TOBI nozzle (nozzle portion) 62S, a partition wall flow channel 63S, a jumper tube (communication flow channel) 64S, a plug portion 65S, a first seal portion 66S and a second seal portion 67S, A protruding portion 68S is provided.

  As shown in FIGS. 2, 3 and 5, the TOBI nozzle 62S is a channel formed on the downstream side of the combustion gas flow in the inner partition wall 61S, and connects the inlet cavity 41C and the outlet cavity 42C. It is. Furthermore, the TOBI nozzle 62S is also a nozzle that extends from the inlet cavity 41C toward the outlet cavity 42C and jets cooling air in the rotation direction of the rotary shaft 5 and the first stage turbine disk 31R. The flow paths formed in the TOBI nozzle 62S are arranged at equal intervals in the circumferential direction.

  As shown in FIGS. 2 and 5, the partition wall channel 63S is a channel formed on the upstream side of the combustion gas flow in the inner partition wall portion 61S, and includes the inlet cavity 41C, the inner partition wall portion 61S, and the connection rotor 38R. It is a flow path connecting the space 44C between the two. Further, the partition channel 63S is formed to extend in the radial direction and is arranged at equal intervals in the circumferential direction.

  2, 3 and 5, the jumper tube 64S is a flow path formed on the downstream side of the combustion gas flow in the inner partition wall 61S, and is between the inner partition wall 61S and the connection rotor 38R. This is a flow path that connects the space 44C and the rim seal cavity 43C. Furthermore, the jumper tube 64S is also a flow path that is inclined radially outward from the space 44C toward the rim seal cavity 43C. The jumper tubes 64S are arranged at equal intervals in the circumferential direction between the flow paths formed in the TOBI nozzle 62S. In other words, the flow paths formed in the jumper tube 64S and the TOBI nozzle 62S are alternately arranged in the circumferential direction.

The plug portion 65S adjusts the flow rate of the cooling air flowing through the jumper tube 64S as shown in FIGS. 2, 3, and 5, and can be attached to and detached from the opening portion of the jumper tube 64S with respect to the rim seal cavity 43C. It can be attached.
The plug portion 65S is formed with a throttle through which cooling air flows, and the flow rate of the cooling air flowing through the jumper tube 64S is adjusted by this throttle.

  As shown in FIGS. 2 and 5, the first seal portion 66S is a space 44C between the inner partition wall 61S and the connection rotor 38R, between the opening of the partition channel 63S and the opening of the jumper tube 64S. It is the seal | sticker part arrange | positioned in. In other words, it is the cooling air that has flowed into the space 44C from the partition wall flow path 63S, and blocks the flow of the cooling air that flows toward the downstream side of the combustion gas flow.

  2 and 5, the second seal portion 67S is a space 44C between the inner partition wall portion 61S and the connection rotor 38R, between the opening portion of the jumper tube 64S and the connection portion with the outlet cavity 42C. It is the seal | sticker part arrange | positioned in. In other words, the cooling air that has flowed into the space 44C from the partition channel 63S and blocks the flow of the cooling air that passes through the first seal portion 66S and flows toward the downstream side of the combustion gas flow.

  The first seal portion 66S and the second seal portion 67S are seal members that extend from the inner partition wall portion 61S toward the connection rotor 38R, and include, for example, leaf seals, labyrinth seals, brush seals, and the like. A combination of the above-described seals may be used and is not particularly limited.

As shown in FIGS. 2, 3 and 5, the protrusion 68 </ b> S is a substantially cylindrical member extending from the downstream end of the combustion gas flow in the inner partition wall 61 </ b> S toward the first stage turbine disk 31 </ b> R. It is a member arranged radially outside the seal arm 37R of the turbine disk 31R.
Furthermore, the projecting portion 68S is provided with a third seal portion 69S that forms a seal structure with the seal arm 37R.

  As shown in FIGS. 2, 3 and 5, the third seal portion 69S is a seal portion disposed between the protruding portion 68S and the seal arm 37R. In other words, the flow of the cooling air flowing into the rim seal cavity 43C from the outlet cavity 42C is blocked.

  The third seal portion 69S is a seal member that extends from the projecting portion 68S toward the seal arm 37R, and is composed of, for example, a leaf seal, a labyrinth seal, a brush seal, or the like, and combines the above-described seals. There is no particular limitation.

  As shown in FIGS. 2 and 3, the rim seal cavity 43C is a space surrounded by the first stage turbine disk 31R, the outer partition wall 51S, and the inner partition wall 61S. Further, the rim seal cavity 43C is a space into which cooling air flows from the jumper tube 64S and the third seal portion 69S, and the cooling air flows through the gap between the first stage stationary blade 11S and the first stage moving blade 21R. It is also a space that flows into the road.

As shown in FIG. 1, the rotating shaft 5 is a columnar member that is rotatably supported around a rotating axis (axis) L, and the rotational driving force generated by the turbine unit 4 is used as a compressor 2 or a generator. To communicate.
In addition, as a rotating shaft 5, a well-known structure can be used and the structure in particular is not limited.

  Next, a general operation in the gas turbine 1 having the above-described configuration will be described, and then a cooling air supply method to the first stage blade 21R and a cooling air supply method to the rim seal cavity 43C will be described. .

  As shown in FIG. 1, the gas turbine 1 sucks air (air) when the compressor 2 is rotationally driven. The sucked air is compressed by the compressor 2 and sent out toward the combustor 3.

  The compressed air flowing into the combustor 3 is mixed with fuel supplied from the outside in the combustor 3. The mixture of air and fuel is combusted in the combustor 3, and hot gas is generated by the combustion heat.

  The hot gas generated in the combustor 3 is supplied from the combustor 3 to the downstream turbine unit 4. The turbine unit 4 is rotationally driven by the high-temperature gas, and the rotational driving force is transmitted to the rotary shaft 5. The rotating shaft 5 transmits the rotational driving force extracted in the turbine unit 4 to the compressor 2 and the generator.

  Next, a method for supplying cooling air to the first stage blade 21R and a method for supplying cooling air to the rim seal cavity 43C, which are features of the present embodiment, will be described.

  As shown in FIG. 1, a part of the air compressed by the compressor 2 is extracted from the compressor 2 and cooled for use as cooling air. Thereafter, as shown in FIGS. 2 and 5, the cooling air is supplied as cooling air having a pressure P0 from the supply flow path 52S to the inlet cavity 41C.

  A part of the cooling air supplied to the inlet cavity 41C flows into the TOBI nozzle 62S and into the outlet cavity 42C due to a pressure difference between the pressure in the inlet cavity 41C and the pressure in the outlet cavity 42C. At this time, the cooling air is given a flow velocity component in the rotation direction of the rotating shaft 5 and the first stage turbine disk 31R, and turns into a swirling flow.

  Here, the pressure of the cooling air in the outlet cavity 42C is P1, and the cooling air flowing into the outlet cavity 42C has a pressure difference (P0−P1) between the pressure P0 in the inlet cavity 41C and the pressure P2 in the outlet cavity 42C. ) Is given.

  The flow velocity component of the cooling air in the outlet cavity 42C in the circumferential direction is equal to or close to the rotational speed at the opening of the first cooling hole 33R and the second cooling hole 34R on the outlet cavity 42C side. Flows into the first cooling hole 33R and the second cooling hole 34R without causing energy loss.

  As shown in FIGS. 3 and 4, the cooling air that has flowed into the first cooling hole 33R flows radially outward and flows into the axial communication path 36R. On the other hand, the cooling air that has flowed into the second cooling hole 34R flows radially outward and flows into the axial communication path 36R via the circumferential communication path 35R and the first cooling hole 33R.

  The cooling air flowing into the axial communication path 36R flows in the direction in which the rotating shaft 5 extends, and flows into the cooling flow path 25R of the first stage blade 21R. The cooling air that has flowed into the cooling flow path 25R cools the blade portion 22R of the first stage moving blade 21R, and is then released, for example, into the combustion gas flow path.

  On the other hand, as shown in FIG. 5, the cooling air that has flowed into the inlet cavity 41C has a difference between the pressure P0 in the inlet cavity 41C and the pressure P3 in the space 44C between the inner partition wall 61S and the connection rotor 38R. In accordance with the pressure (P0-P3), it flows into the space 44C via the partition wall flow path 63S.

  A part of the cooling air flowing into the space 44 </ b> C flows toward the upstream side of the combustion gas flow, in other words, toward the compressor 2, and is used for cooling the compressor 2. The remaining cooling air flows toward the first seal portion 66S according to the pressure difference (P3-P1) between the pressure (P3) in the space 44C and the pressure (P1) in the outlet cavity 42C.

The cooling air that has passed through the first seal portion 66S is divided into cooling air that flows into the jumper tube 64S and cooling air that flows toward the second seal portion 67S.
The cooling air that has flowed into the jumper tube 64S passes through the restriction of the plug portion 65S and flows into the rim seal cavity 43C.

In addition, cooling air that has passed through the third seal portion 69S flows into the rim seal cavity 43C from the exit cavity 42C.
Since the pressure P4 in the rim seal cavity 43C is higher than the pressure P2 in the flow path through which the combustion gas flows, depending on the differential pressure (P4-P2), the cooling air in the rim seal cavity 43C and the first stage stationary blade 11S and It flows out from the gap between the first stage blades 21R into the combustion gas flow path.
Therefore, the combustion gas does not flow into the rim seal cavity 43C.

  According to the above configuration, the cooling air flowing into the space 44C from the inlet cavity 41C flows into the space 44C between the wall portion 61S and the connecting rotor 38R via the jumper tube 64S, and enters the first stage stationary blade 11S and the first stage. The cooling air flows out from the gap between the rotor blades 21R into the combustion gas flow path, in other words, the cooling air is used for the rim seal. Therefore, the inflow of cooling air from the space 44C to the outlet cavity 42C is suppressed, the deceleration of the flow velocity component in the rotational direction in the cooling air flowing out from the TOBI nozzle 62S can be suppressed, and the performance deterioration in the gas turbine 1 is suppressed. be able to.

  Furthermore, the flow rate of the cooling air flowing into the rim seal cavity 43C can be changed by replacing the plug portion 65S attached to the jumper tube 64S with another plug portion having a different throttle diameter. Therefore, for example, in the gas turbine 1 of the present embodiment, the plug portion 65S should be replaced even when refurbishment with performance improvement or the like is performed and the flow rate of the cooling air used for the rim seal needs to be changed. Thus, the flow rate of the cooling air used for the rim seal can be easily changed.

1 Gas turbine 2 Compressor (compression unit)
3 Combustor (combustion section)
4 Turbine part 11S 1st stage stationary blade (stator blade)
13S Casing 21R 1st stage blade (blade)
25R Cooling flow path 31R Single stage turbine disk (turbine disk)
33R 1st cooling hole (cooling hole)
34R Second cooling hole (cooling hole)
41C inlet cavity (first inlet space)
42C outlet cavity (second inlet space)
44C space (bypass)
43C Rim seal cavity (rim seal space)
61S Inner partition (partition)
62S TOBI nozzle (nozzle part)
64S jumper tube (communication flow path)
65S plug part L rotation axis (axis)

Claims (3)

A plurality of rotor blades are attached to the outer periphery side by side in the circumferential direction, and a turbine disk that is rotatably supported around an axis,
A plurality of cooling holes extending from the inside to the outside of the turbine disk and arranged side by side in the circumferential direction, and supplying a cooling medium to a cooling flow path inside the rotor blade;
A first inlet space to which the cooling medium is supplied from the outside, and a partition that divides the second inlet space communicated with the plurality of cooling holes;
A nozzle portion having a flow path that extends in the rotation direction of the turbine disk from the first inlet space toward the second inlet space inside the partition portion;
A stationary blade that is adjacent to the turbine disk in the axial direction and is attached to the inner peripheral surface of a casing that accommodates the plurality of moving blades inside the turbine disk, alternately along with the plurality of moving blades along the axial direction. And a rim seal space communicated with the gap between the rotor blades,
A communication channel extending from the bypass space that bypasses the nozzle portion and connects the first inlet space and the second inlet space to the rim seal space, inside the partition portion;
A cooling structure for a turbine, wherein:   2. The turbine according to claim 1, further comprising: a throttle portion that adjusts a flow rate of the cooling medium flowing through the communication flow path, and a plug portion that is detachably attached to the communication flow path. Cooling structure. A compression section that sucks and compresses air;
A combustion section that generates a combustion gas by combusting a mixture of compressed air and fuel supplied from the outside;
A turbine section for extracting rotational driving force from the generated combustion gas;
Is provided,
A gas turbine, wherein the turbine cooling structure according to claim 1 or 2 is provided in the turbine section.

Images