JP4764219B2 - Gas turbine seal structure - Google Patents

Description

  The present invention relates to a gas turbine seal structure, and more particularly, is disposed between an inner shroud disposed radially inward of a stationary blade blade body and a platform disposed radially inward of the blade blade body. The present invention relates to a gas turbine seal structure.

As a gas turbine seal structure disposed between an inner shroud disposed radially inward of a stationary blade blade body and a platform disposed radially inward of the blade blade body, for example, Patent Document 1, 2 is known.
JP 2004-150435 A JP 11-324608 A

In the gas turbine seal structure disclosed in Patent Document 1, the combustion gas that has passed between the inner peripheral surface of the rear edge of the inner shroud and the outer peripheral surface of the front edge of the platform passes radially inward through the wheel space. It is guided linearly (straightly) to the (rotor side). Therefore, the combustion gas that has passed between the inner peripheral surface of the rear edge of the inner shroud and the outer peripheral surface of the front edge of the platform efficiently collides with the cooling air that flows along the inner peripheral surface of the platform arm. In order to sufficiently reduce the temperature of the combustion gas, a large amount of cooling air is required, and the efficiency of the gas turbine is reduced.
The problem that a large amount of cooling air is required and the efficiency of the gas turbine is reduced is that an industrial gas turbine adopting a half-crack structure in which the outer peripheral side is divided into two vertically, that is, in FIGS. This is particularly a problem in gas turbines having the indicated gap t.
In the gas turbine seal structure disclosed in Patent Document 2, a guide plate member is provided in a portion of the inner peripheral casing of the stationary blade facing the turbine disk of the moving blade. Therefore, there is a problem that the sealing structure becomes complicated and the manufacturing cost increases.

The present invention has been made in view of the above circumstances, and between a rear edge inner peripheral surface of an inner shroud and a front edge outer peripheral surface of a platform with a small amount of cooling air without complicating the structure. It is an object of the present invention to provide a gas turbine seal structure capable of reducing the amount of passing combustion gas.
Another object of the present invention is to provide a gas turbine seal structure that can be applied to an industrial gas turbine having a half-cracked structure.

The present invention employs the following means in order to solve the above problems.
A gas turbine seal structure according to the present invention includes a gas turbine seal disposed between an inner shroud disposed radially inward of a stationary blade blade body and a platform disposed radially inward of the blade blade body. a structure, an arm portion provided on at least one side of the front edge or rear edge of the platform, at least one side of edge or front edge after said inner shroud, and surrounds the distal end portion of the arm portion An inner peripheral surface located radially inward of the arm portion, extending from the base end portion to the distal end surface toward the radially outer side and extending toward the recess. is a, the recesses, together with the located radially inward from the distal end surface, previous SL platform of the blade disk mounted A side surface extending in the axial direction so as to be rotating shaft and the flat row, from one end of the side surface, a bottom surface extending to be parallel with the tip surface with radially toward the outside, radially outward of the bottom surface And an inclined surface extending outward in the radial direction and extending toward the arm portion .
According to such a gas turbine seal structure, the cooling air that has been supplied (flowed) to the side of the arm portion (radially outward) with the rotation of the rotor blade disk is the inner peripheral surface of the arm portion. Is smoothly guided toward the distal end surface (that is, toward the recess). On the other hand, the combustion gas that has entered through between the inner shroud and the arm portion is guided toward the distal end surface along the side surface forming the recess. The cooling air guided toward the tip surface along the inner peripheral surface of the arm portion and the combustion gas guided toward the tip surface along the side surface are radially inward of the tip surface (for example, In a portion surrounded by a broken line in FIG. 3, they collide (collide). At this time, since the cooling air and the combustion gas are sufficiently mixed, it is possible to sufficiently reduce the temperature of the combustion gas that has entered, and to the inside in the radial direction (that is, on the blade disk side). The combustion gas having a low temperature can be allowed to flow. Further, since the combustion gas is blocked by the cooling air inside the tip surface in the radial direction (for example, a portion surrounded by a broken line in FIG. 3), the combustion gas is prevented from entering. Can be reduced.

Moreover, as shown in FIG. 3, the combustion gas that has entered between the inner shroud and the arm portion is radially inward of the front end surface (for example, a portion surrounded by a broken line in FIG. 3). The traveling direction can be changed. In other words, the combustion gas that has passed through the recess (and comes out of the recess) proceeds in a direction substantially parallel to the rotating shaft of the rotor blade disk. Thereby, the entering combustion gas can be efficiently collided with the cooling air, and the temperature of the combustion gas can be lowered more efficiently.
Further, as shown in FIG. 2, the cooling air C flowing along the inner peripheral surface 41a is caused by the inclined surface 42b between the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d. In the meantime, the direction of travel will be changed. As a result, the cooling air C flowing along the inner peripheral surface 41a can smoothly flow out from between the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d. The inclined surface 42b, the rear edge inner peripheral surface 42a, and the tip of the arm portion 41 can be efficiently cooled.

A gas turbine seal structure according to the present invention includes a gas turbine seal disposed between an inner shroud disposed radially inward of a stationary blade blade body and a platform disposed radially inward of the blade blade body. a structure, an arm portion provided on at least one side of the front edge or rear edge of the platform, at least one side of edge or front edge after said inner shroud, and surrounds the distal end portion of the arm portion An inner peripheral surface located radially inward of the arm portion, extending from the base end portion to the distal end surface toward the radially outer side and extending toward the recess. is a, the recesses, together with the located radially inward from the distal end surface, toward the other end from one end and toward the radially outer DOO A side surface extending toward the tip surface, a bottom surface extending radially outward from one end of the side surface and extending parallel to the tip surface, and one end positioned radially outward of the bottom surface And an inclined surface extending outward in the radial direction and extending toward the arm portion .
According to such a gas turbine seal structure, the cooling air that has been supplied (flowed) to the side of the arm portion (radially outward) with the rotation of the rotor blade disk is the inner peripheral surface of the arm portion. Is smoothly guided toward the distal end surface (that is, toward the recess). On the other hand, the combustion gas that has entered through between the inner shroud and the arm portion is guided toward the distal end surface along the side surface forming the recess. The cooling air guided toward the tip surface along the inner peripheral surface of the arm portion and the combustion gas guided toward the tip surface along the side surface are radially inward of the tip surface (for example, In a portion surrounded by a broken line in FIG. 3, they collide (collide). At this time, since the cooling air and the combustion gas are sufficiently mixed, it is possible to sufficiently reduce the temperature of the combustion gas that has entered, and to the inside in the radial direction (that is, on the blade disk side). The combustion gas having a low temperature can be allowed to flow. Further, since the combustion gas is blocked by the cooling air inside the tip surface in the radial direction (for example, a portion surrounded by a broken line in FIG. 3), the combustion gas is prevented from entering. Can be reduced.

Further, as shown in FIG. 3, the combustion gas that has entered between the inner shroud and the arm portion has a side surface and a traveling direction on the tip surface (for example, a portion surrounded by a one-dot chain line in FIG. 4). Will be changed. That is, the main flow of the combustion gas that has passed through the recess (out of the recess) proceeds in a direction away from the rotating shaft of the rotor blade disk, and a part of the main stream flows to the rotating shaft of the rotor blade disk. It will proceed in a substantially parallel direction. As a result, the efficiency of the combustion gas that has entered in the vicinity of the tip surface (for example, a portion surrounded by a one-dot chain line in FIG. 4) and in the radial direction inside the tip surface (for example, a portion surrounded by a broken line in FIG. 4) is improved. It can be made to collide well with the cooling air, and the temperature of the combustion gas can be lowered more efficiently.
Furthermore, since the main flow of the combustion gas that has passed through the recess (exited from the recess) proceeds in a direction away from the rotating shaft of the blade disk, the combustion gas that travels to the blade disk side. The amount of can be reduced.
Furthermore, as shown in FIG. 2, the cooling air C flowing along the inner peripheral surface 41a is caused by the inclined surface 42b to the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d. During that time, the direction of travel will be changed. As a result, the cooling air C flowing along the inner peripheral surface 41a can smoothly flow out from between the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d. The inclined surface 42b, the rear edge inner peripheral surface 42a, and the tip of the arm portion 41 can be efficiently cooled.

In the gas turbine seal structure described above, it is further preferable that the inner shroud is provided with a cooling air supply hole for jetting cooling air toward the combustion gas flowing along the wall surface forming the recess. It is.
According to such a gas turbine seal structure, for example, as shown in FIG. 5, the cooling air is jetted toward the combustion gas flowing along the bottom surface of the recess, and is surrounded by a two-dot chain line in FIG. 5. In this portion, the combustion gas flowing along the bottom surface of the recess and the cooling air ejected from the cooling air supply hole collide (collision). At this time, since the cooling air and the combustion gas are sufficiently mixed, the temperature of the combustion gas that has entered can be sufficiently lowered.

The gas turbine according to the present invention can reduce the amount of combustion gas passing between the inner peripheral surface of the inner edge of the inner shroud and the outer peripheral surface of the front edge of the platform with a small amount of cooling air. The seal structure is provided.
According to such a turbine, since the amount of cooling air is small, consumption of the cooling air can be reduced and the performance of the gas turbine can be improved.

  According to the present invention, the amount of combustion gas passing between the inner peripheral surface of the rear edge portion of the inner shroud and the outer peripheral surface of the front edge portion of the platform can be reduced with a small amount of cooling air. Play.

Hereinafter, a gas turbine seal structure according to a first embodiment of the present invention will be described with reference to the drawings.
FIG. 1 is a schematic longitudinal sectional view of a turbine section 30 of a gas turbine 10 provided with a gas turbine seal structure (hereinafter referred to as “seal structure”) 40 according to the present embodiment.
As shown in FIG. 1, a gas turbine 10 includes a compressor (not shown) that compresses combustion air and sends it to a combustor 20, and combustion air that is sent from the compressor (compressor). A combustor 20 for injecting and burning fuel to generate a high-temperature combustion gas G, and a turbine unit 30 that is located downstream of the combustor 20 and driven by the combustion gas G exiting the combustor 20. It is configured as the main element.

The turbine section 30 includes, for example, a first stage (row 1) stationary blade 31, a first stage moving blade 32, a second stage (row 2) stationary blade 33, a second stage moving blade 34, a third stage (row 3) stationary blade 35, 3 from the upstream side. A step rotor blade 36, a four-stage (row 4) stationary blade 37, and a four-stage rotor blade 38 are arranged in order.
The first stage stationary blade 31, the second stage stationary blade 33, the third stage stationary blade 35, and the fourth stage stationary blade 37, respectively, are outer shrouds 31 a, 33 a, 35 a, 37 a, stationary blade blade bodies 31 b, 33 b, 35 b, 37 b having cross-sectional airfoils, And inner shrouds 31c, 33c, 35c, 37c, which are fixed to a stationary vehicle compartment (not shown) via outer shrouds 31a, 33a, 35a, 37a.
The first-stage stationary blade 31, the second-stage stationary blade 33, the third-stage stationary blade 35, and the fourth-stage stationary blade 37 and the fourth-stage stationary blade 37 expand and reduce the combustion gas G, and the outflow gas from the stationary blades 31, 33, 35, 37 is located downstream It functions to give the direction of flow so as to collide with moving blades 32, 34, 36, and 38 at an optimum angle.

  The first-stage moving blade 32, the second-stage moving blade 34, the third-stage moving blade 36, and the fourth-stage moving blade 38 are each provided with a moving blade blade body 32b, 34b, 36b, 38b. It is attached to the blade disks 32e, 34e, 36e, 38e via fixed platforms 32d, 34d, 36d, 38d.

Here, the flow of the combustion gas G will be described. The combustion gas G that has been heated by the combustor 20 flows into the first stage stationary blade 31 and expands in the process of flowing between the second to fourth stage blades. The rotor blades 32, 34, 36, and 38 are rotated, respectively, and the turbine rotor (turbine rotor) is given rotational power to be discharged.
In addition, the code | symbol R in FIG. 1 is a rotating shaft of the rotor blade disk 32e, 34e, 36e, 38e.

Next, the seal structure 40 according to the present embodiment will be described in detail with reference to FIGS. FIG. 2 and FIG. 3 are enlarged views showing in detail the portion surrounded by a circle A ′ in FIG.
As shown in FIGS. 2 and 3, the seal structure 40 includes an arm portion 41 provided on the front edge side (upstream side of the flow of the combustion gas G) of the platform 34d and the rear edge side (combustion gas) of the inner shroud 33c. And a recess 42 provided at a position facing the distal end of the arm 41 and downstream of the G flow.

The inner peripheral surface (inner peripheral surface facing the rotation axis R (see FIG. 1) of the rotor blade disks 32e, 34e, 36e, 38e) 41a located on the radially inner side of the arm portion 41 is a base end portion (the rotor blade). The inclined surface gradually increases in diameter from the root portion (on the blade body 34b side) to the tip surface 41b.
On the other hand, an outer peripheral surface (outer peripheral surface facing the inner peripheral surface 42a of the rear edge portion of the inner shroud 33c) 41c located on the radially outer side of the arm portion 41 is a base end portion (the root of the blade blade main body 34b). The inclined surface gradually decreases in diameter from the first portion toward the distal end and then gradually increases in diameter toward the distal end, that is, an inclined surface having a substantially V-shaped sectional view.
Further, at the tip of the arm portion 41, a convex portion (flange portion) 41d having a rectangular shape in cross section is provided that protrudes toward the inner peripheral surface 42a of the rear edge portion of the inner shroud 33c.
The tip surface 41b is formed so as to be substantially parallel to a surface orthogonal to the rotational axis R of the rotor blade disks 32e, 34e, 36e, 38e, and the outer peripheral surface 41e located on the radially outer side of the convex portion 41d. Is formed so as to be substantially parallel to the rotation axis R of the rotor blade disks 32e, 34e, 36e, 38e.

The recess 42 is formed on the rear edge side of the inner shroud 33c and on the radially outer side so as to surround (accommodate) the distal end portion of the arm portion 41, and a rear edge inner peripheral surface (wall surface) 42a; An inclined surface (wall surface) 42b, a bottom surface (wall surface) 42c, and a side surface (wall surface) 42d are provided.
The rear edge inner peripheral surface 42a faces the outer peripheral surface 41e of the convex portion 41d and is substantially parallel to the outer peripheral surface 41e of the convex portion 41d (that is, the rotating shafts of the rotor blade disks 32e, 34e, 36e, 38e). (So as to be substantially parallel to R).
The bottom surface 42c is positioned on the front edge side (upstream side with respect to the flow of the combustion gas G) with respect to the rear edge side end surface 33c ′ positioned on the radially inner side of the inner shroud 33c, and is substantially parallel to the front end surface 41b. That is, it is formed (so as to be substantially parallel to a plane perpendicular to the rotation axis R of the rotor blade disks 32e, 34e, 36e, 38e).
The inclined surface 42b connects one end on the front edge side of the inner peripheral surface 42a of the rear edge portion and one end on the outer side in the radial direction of the bottom surface 42c, and gradually increases in diameter from the inner side to the outer side in the radial direction. It is.
The side surface 42d is a plane that connects the other end on the radially inner side of the bottom surface 42c and the one end on the radially outer side of the rear edge side end surface 33c ′, and is substantially parallel to the inner peripheral surface 42a of the rear edge (that is, The rotor blade disks 32e, 34e, 36e, 38e are formed so as to be substantially parallel to the rotation axis R of the rotor blade disks 32e, 34e, 36e, 38e.
In addition, a gap t is formed between the rear edge side end surface 33c ′ located on the inner side in the radial direction of the inner shroud 33c and the front end surface 41b of the arm portion 41, and can be assembled in a gas turbine having a half crack structure. It has a structure.

According to the seal structure 40 according to the present embodiment, the cooling air C supplied (flowed) to the arm portion 41 side (radially outward) with the rotation of the rotor blade disk 34e is supplied to the arm portion 41. It will be smoothly guided along the inner peripheral surface 41a toward the tip end surface 41b (that is, toward the recess 42). On the other hand, the combustion gas G that has entered between the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d reaches the front end surface 41b along the inclined surface 42b, the bottom surface 42c, and the side surface 42d. Will be led to. Then, the cooling air C guided toward the front end surface 41b along the inner peripheral surface 41a of the arm portion 41 and the front end surface 41b along the inclined surface 42b, the bottom surface 42c, and the side surface 42d. Combustion gas G collides (collises) on the inner side in the radial direction of the front end surface 41b (that is, the portion surrounded by the broken line in FIG. 3). At this time, the cooling air C and the combustion gas G are sufficiently mixed, so that the temperature of the combustion gas G that has entered can be sufficiently lowered, and the inner side in the radial direction (that is, the blade disk). 34e side) can be made to flow the combustion gas G having a low temperature. In addition, the combustion gas G is blocked by the cooling air C on the radially inner side of the front end surface 41b (that is, the portion surrounded by the broken line in FIG. 3), and the state where the entry of the combustion gas G is hindered. Therefore, the amount of combustion gas G entering can be reduced. Furthermore, since the amount of the cooling air C is small, the consumption amount of the cooling air C can be reduced and the performance of the gas turbine 10 can be improved.
Moreover, since the seal structure 40 according to the present embodiment has a simple structure, the manufacturing cost can be reduced.

  As shown in FIG. 3, the combustion gas G that has flowed along the inclined surface 42b and the bottom surface 42c is radially inward of the tip surface 41b (that is, the portion surrounded by the broken line in FIG. 3) by the side surface 42d. The direction of travel will be changed. That is, the combustion gas G that has passed through the recess 42 (emitted from the recess 42) proceeds in a direction substantially parallel to the rotation axis R of the rotor blade disks 32e, 34e, 36e, and 38e. Thereby, the combustion gas G which has entered can be efficiently collided with the cooling air C, and the temperature of the combustion gas G can be reduced more efficiently.

  Further, since a gap t is formed between the rear edge side end surface 33c ′ located on the inner side in the radial direction of the inner shroud 33c and the front end surface 41b of the arm portion 41, the gas turbine 10 is half-finished. When the type is a crack type (type in which the upper compartment is mounted on the lower compartment), the upper compartment can be easily removed from the lower compartment and the upper compartment can be easily placed on the lower compartment. It can be mounted on.

  Furthermore, as shown in FIG. 2, the cooling air C flowing along the inner peripheral surface 41a is caused by the inclined surface 42b to the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d. During that time, the direction of travel will be changed. As a result, the cooling air C flowing along the inner peripheral surface 41a can smoothly flow out from between the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d. The inclined surface 42b, the rear edge inner peripheral surface 42a, and the tip of the arm portion 41 can be efficiently cooled.

  Furthermore, as shown in FIG. 2, a part of the cooling air C that flows out (leaks out) between the rear inner peripheral surface 42a of the inner shroud 33c and the outer peripheral surface 41e of the convex portion 41d is A vortex is created and retained on the surface 41c, and a part flows along the outer peripheral surface 41c toward the root portion of the blade main body 34b. On the other hand, as shown in FIG. 3, the combustion gas G flowing from the root portion of the moving blade blade body 34 b toward the convex portion 41 d along the outer peripheral surface 41 c changes its traveling direction outward in the radial direction at the convex portion 41 d. A part of which is swirled and stays, and part of the gas flows together with the main flow of the combustion gas G and flows into the recess 42 between the inner peripheral surface 42a and the outer peripheral surface 41e of the rear edge. Become. At this time, since the cooling air C and the combustion gas G are sufficiently mixed on the outer peripheral surface 41c, the temperature of the combustion gas G flowing along the outer peripheral surface 41c can be sufficiently lowered, and The arm part 41 can be efficiently cooled. Further, since the combustion gas G is blocked by the vortex generated on the outer peripheral surface 41c near the convex portion 41d and the ingress of the combustion gas G is hindered, the amount of the combustion gas G entering can be reduced. it can. Furthermore, since the amount of the cooling air C is small, the consumption amount of the cooling air C can be reduced and the performance of the gas turbine 10 can be improved.

A second embodiment of the seal structure according to the present invention will be described with reference to FIG.
The seal structure 50 according to the present embodiment is different from that of the first embodiment described above in that a side surface 52d is provided instead of the side surface 42d. Since other components are the same as those of the first embodiment described above, description of these components is omitted here.

  The side surface 52d is a plane that connects the other end on the radially inner side of the bottom surface 42c and one end on the radially outer side of the trailing edge side end surface 33c ′, and extends radially inward from the trailing edge side end surface 33c ′ side to the bottom surface 42c side. It is the inclined surface formed so that it might expand in diameter. That is, the side surface 52d is inclined with respect to the rear edge inner peripheral surface 42a (in other words, with respect to the rotational axis R of the rotor blade disks 32e, 34e, 36e, 38e, for example, approximately 20 degrees). ) Is formed.

According to the seal structure 50 according to the present embodiment, as shown in FIG. 4, the combustion gas G that has flowed along the inclined surface 42 b and the bottom surface 42 c is caused by the side surface 52 d by the front end surface 41 b (that is, in FIG. 4). The traveling direction can be changed to a portion surrounded by a one-dot chain line). That is, the main flow of the combustion gas G that has passed through the recess 42 (out of the recess 42) proceeds in a direction away from the rotation axis R of the rotor blade disks 32e, 34e, 36e, and 38e. A part thereof proceeds in a direction substantially parallel to the rotation axis R of the rotor blade disks 32e, 34e, 36e, 38e. As a result, the combustion gas G that has entered enters the vicinity of the tip surface 41b (that is, the portion surrounded by the one-dot chain line in FIG. 4) and the inside in the radial direction of the tip surface 41b (that is, the portion surrounded by the broken line in FIG. 4). , It is possible to efficiently collide with the cooling air C, and the temperature of the combustion gas G can be lowered more efficiently.
Further, the main flow of the combustion gas G that has passed through the recess 42 (emitted from the recess 42) proceeds in a direction away from the rotation axis R of the rotor blade disks 32e, 34e, 36e, 38e. Therefore, it is possible to reduce the amount of the combustion gas G that travels toward the blade disks 32e, 34e, 36e, and 38e.
Other operational effects are the same as those of the first embodiment described above, and therefore the description thereof is omitted here.

A third embodiment of the seal structure according to the present invention will be described with reference to FIG.
The seal structure 60 according to the present embodiment is the same as that of the second embodiment described above in that a cooling air supply hole 61 for supplying cooling air C is provided on the bottom surface 42c of the recess 42. Different. Since other components are the same as those of the second embodiment described above, description of these components is omitted here.

  The cooling air supply hole 61 is a hole for injecting the cooling air C toward the combustion gas G flowing along the bottom surface 42 c of the recess 42, and the cooling air C is supplied to the upstream side thereof. A source (not shown) is connected.

According to the seal structure 60 according to the present embodiment, as shown in FIG. 5, the cooling air C is ejected toward the combustion gas G flowing along the bottom surface 42c of the recess 42, and the two-dot chain line in FIG. The combustion gas G flowing along the bottom surface 42c of the recess 42 and the cooling air C ejected from the cooling air supply hole 61 collide (collision) with each other in the portion surrounded by. At this time, since the cooling air C and the combustion gas G are sufficiently mixed, the temperature of the combustion gas G that has entered can be sufficiently lowered.
The other operational effects are the same as those of the second embodiment described above, and the description thereof is omitted here.

  In the above-described embodiment, the gas turbine in which the stationary blades and the moving blades of the turbine section are each configured in four stages has been described. However, the seal structure according to the present invention is not applicable only to a gas turbine having such a configuration, and a gas turbine in which the stationary blades and the moving blades of the turbine section are each provided with three stages or less, or five stages or more. It can also be applied to.

  Moreover, in this embodiment mentioned above, the sealing structure by this invention is made into the front edge side (upstream with respect to the flow of the combustion gas G) of the platform 34d, and the rear edge side (with respect to the flow of the combustion gas G) of the inner shroud 33c. However, the present invention is not limited to this, and the sealing structure according to the present invention is applied to the rear edge side of the platform 34d and the front of the inner shroud 33c. It can also be applied between the edge side.

  Further, in the above-described embodiment, the seal structure according to the present invention has been described with the bottom applied to a half-crack type industrial gas turbine. However, the present invention is not limited to this, and the present invention is not limited thereto. The seal structure according to the invention can also be applied to a shaft type (a type in which a rotor is fitted from the axial direction) of an aircraft gas turbine.

It is a schematic longitudinal cross-sectional view of the turbine part of the gas turbine provided with the gas turbine seal structure of the present invention. FIG. 2 is a partially enlarged view showing in detail a portion surrounded by a circle A ′ in FIG. 1, for explaining the flow of cooling air. FIG. 2 is a partially enlarged view showing in detail an enlarged portion surrounded by a circle A ′ in FIG. 1, for explaining the flow of combustion gas. It is a figure for demonstrating 2nd Embodiment of the seal structure of the gas turbine of this invention, Comprising: It is a figure similar to FIG. It is a figure for demonstrating 3rd Embodiment of the seal structure of the gas turbine of this invention, Comprising: It is a figure similar to FIG. 3 and FIG.

Explanation of symbols

10 Gas turbine.
33b Stator blade blade body 33c Inner shroud 33c 'trailing edge side end surface 34b Blade blade body 34d Platform 34e Blade disk 40 Seal structure 41 Arm portion 41a Inner circumferential surface 41b Tip surface 42 Recess 42a Rear edge inner circumferential surface (wall surface)
42b Inclined surface (wall surface)
42c Bottom (wall surface)
42d Side surface (wall surface)
50 Seal structure 52d Side surface (wall surface)
60 Seal structure 61 Cooling air supply hole C Cooling air G Combustion gas R Rotating shaft

Claims (4)

A gas turbine seal structure disposed between an inner shroud disposed radially inward of a stationary blade body and a platform disposed radially inward of the blade body;
An arm portion provided on at least one side of the front edge side or the rear edge side of the platform, and at least one side of the rear edge side or the front edge side of the inner shroud, and so as to surround the front end portion of the arm portion. With a recess,
The inner peripheral surface located on the radially inner side of the arm portion is an inclined surface extending from the base end portion to the distal end surface and extending radially outward and toward the recess ,
And a side in which the recess, with located radially inward from the distal end surface, extending in the axial direction as before Symbol platform is on the rotary shaft and the flat row of rotor blade disks attached,
A bottom surface extending from one end of the side surface radially outward and parallel to the tip surface;
A gas turbine seal structure comprising: an inclined surface extending outward in the radial direction and extending toward the arm portion from one end located radially outward of the bottom surface . A gas turbine seal structure disposed between an inner shroud disposed radially inward of a stationary blade body and a platform disposed radially inward of the blade body;
An arm portion provided on at least one side of the front edge side or the rear edge side of the platform, and at least one side of the rear edge side or the front edge side of the inner shroud, and so as to surround the front end portion of the arm portion. With a recess,
The inner peripheral surface located on the radially inner side of the arm portion is an inclined surface extending from the base end portion to the distal end surface and extending radially outward and toward the recess ,
The recess is located radially inward from the tip surface, and extends from one end to the other end in the radial direction and extending toward the tip surface.
A bottom surface extending from one end of the side surface radially outward and parallel to the tip surface;
A gas turbine seal structure comprising: an inclined surface extending outward in the radial direction and extending toward the arm portion from one end located radially outward of the bottom surface .   The cooling air supply hole which injects the air for cooling toward the combustion gas which flows along the wall surface which forms the said recess in the said inner shroud is provided. Gas turbine seal structure.   A gas turbine comprising the gas turbine seal structure according to any one of claims 1 to 3.

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