JPH09242563A - Gas turbine cooling system - Google Patents

Abstract

PROBLEM TO BE SOLVED: To lower the temperature gradient inside of a disc, and to lower the thermal stress to be generated inside of the disc without generating a contact of the surface of the disc with the dynamic vane cooling medium before and after cooling in the inner peripheral side from the abutment surface with an adjacent spacer. SOLUTION: A gas turbine is provided with a dynamic vane 5, a static vane 4, a gas pressure sealing shroud 5, and a disc 1 provided with a fur tree groove, in which a fur tree provided in a root part of the dynamic vane 2 is planted for fixation. The disc 1 is provided with a hub part 15 in an abutment surface 7 with an adjacent spacer 6 and provided with a faucet joint part 16 as a constraint in the radial direction, and the abutment surface 7 is provided with bolt holes 8, of which radius position from a rotary center shaft 17 is equal to each other. Supply and recovery of the cooling medium to/from the dynamic vane is performed by using a hollow shaft 22 provided at a central part and a spacer 26 having a hole 23 opened in the radial direction from the surface of the center hole.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine facility, and more particularly to a turbine cooling system.

[0002]

2. Description of the Related Art A conventional turbine blade cooling system introduces air extracted from an arbitrary stage of a compressor into the turbine to warm or cool the disk to reduce the temperature gradient generated in the disk. In general, an open cooling system is used in which after cooling, it is introduced into the moving blade to reduce the temperature of the moving blade metal, and the cooled air is discharged into the turbine gas passage.

However, in recent years, gas turbine equipment has been required to have high system efficiency for the purpose of energy saving and environmental protection. A closed cooling system that collects a moving blade cooling medium after cooling the moving blade without discharging it into a turbine gas flow channel is being adopted as a means of improving efficiency.

When a closed cooling system is adopted, it is necessary to separately collect the cooling medium that has become hot after cooling the moving blades. In the conventional open cooling system, the cooling medium, which became high in temperature after cooling the moving blades, could be discharged to the turbine gas flow path, so such a recovery method in the turbine cooling system did not pose a particular problem. However, if a closed cooling system is adopted, a system that supplies this cooling medium to the moving blades, cools it, and then collects it without contacting the cooling medium before and after cooling is required from the viewpoint of cooling efficiency. Become.

As an example of such a cooling medium supply / recovery system, the Yokohama International Gas Turbine Society (1995
Year) proposed steam closed cooling system. FIG. 1 is a structural cross-sectional view of a turbine portion of a known example.
FIG. 1 is a cross-sectional view of a stacked rotor in which disks to which a plurality of moving blades are fitted are stacked in a multi-stage unit in the rotational axis direction. The left side is the upstream front stage side, and the right side is the downstream rear stage side. In the figure, 1 is a disk, 2 is a moving blade, 3 is a stacking bolt,
Reference numeral 4 is a vane, 5 is a shroud, and 6 is a spacer. Structurally, the disks 1 are fastened by stacking bolts 3 through the bolt holes 8 formed in the contact surface 7 with the adjacent spacer 6 and stacking from both ends.

In the rotor disk 1 of FIG. 1, a hole 9 through which a cooling medium flows is provided in the contact surface 7 with the adjacent spacer 6 in addition to the hole 8 through which the stacking bolt 3 penetrates to cool the moving blade 2. Realizes a closed cooling system which is isolated from each other by separately passing through the holes 9 for the cooling medium before and after cooling.

[0007]

In the gas turbine, it is expected that the efficiency will be further improved and the turbine inlet temperature will be increased as one means thereof in the future. In order to meet this demand, by extending the conventional technology, if a method is adopted in which the cooling medium is simply passed through another path before and after cooling so that they do not come into contact with each other, the cooling of the blades with the rise in turbine inlet temperature Due to the subsequent increase in medium temperature, the temperature difference between before and after cooling of the cooling medium becomes large. Therefore, when a closed cooling system is adopted, a large thermal stress due to the temperature difference between before and after cooling of the cooling medium is further superposed inside the disk at a high level even with the centrifugal stress of a single unit. Therefore, there is a problem in that the mutual isolation of the cooling medium before and after the cooling is not sufficient to cope with the higher efficiency and higher temperature in terms of the stress generated in the disk.

[0008]

The above-mentioned problem is solved by providing a hollow shaft provided at the center and a spacer having a hole formed in the radial direction from the surface of the central hole for supplying or recovering the cooling medium to the moving blade. It is solved by using it.

Since the turbine cooling system of the present invention has the structure of the above-mentioned solving means, the surface of the disk on the inner peripheral side of the contact surface with the adjacent spacer is for cooling the moving blades before and after cooling. The medium does not touch either. Therefore, the temperature gradient inside the disk is reduced as much as possible. As a result, the thermal stress generated in the disk can be reduced.

[0010]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS A rotor disk according to the present invention comprises a plurality of blades for receiving work from a fluid, a rotating disk in which the blades are implanted, and a stacking bolt for fastening the disk through a plurality of stages. Have. The rotor blade has a fir tree to be implanted in the disc at its root, and a groove is cut on the outer peripheral surface of the disc at an arbitrary angle in the rotational axis direction or the circumferential direction to implant the fir tree of the rotor blade. It is supposed to be. In addition, on both sides of the disk, a hub portion with a constant thickness is provided as a contact surface with the adjacent spacer, and the hub portion has a group of holes equally spaced in the circumferential direction with the same radial position from the rotation center axis. A plurality of stacking bolts are passed through and fastened by a plurality of stacking bolts so that a plurality of stages of disks are stacked and fixed in the axial direction. Further, the disk has a spigot portion at a contact portion with an adjacent spacer and is fixed in the radial direction by meshing with the adjacent spacer. Hereinafter, embodiments of the present invention will be described with reference to the drawings.

2 and 3 are cross-sectional views of the turbine cooling system of the present invention. 2 is a sectional view of the cooling system in the rotation axis direction, and FIG. 3 is a structural schematic view of the four-stage disc 10 in FIG. In FIG. 2, the chain double-dashed line shows the flow 11 of the medium before cooling the moving blade, and the broken line shows the flow 12 of the medium after cooling the moving blade. 2 and 3, the rotor blades 2 and the inter-blade gas pressure sealing shrouds 5 are alternately arranged from the upstream side to the downstream side of the turbine. The rotor blade 2 is fixed by implanting a fir tree 13 provided at the root part in a fir tree groove 14 formed in the outer peripheral part of the disk 1. The turbine channel cross-sectional area tends to increase from upstream to downstream. The disk 1 has a hub portion 15 and a spigot portion 1 on a contact surface 7 with an adjacent spacer 6.
6 are provided, and they are constrained in the radial direction by meshing with the adjacent spacers 6. Further, on the contact surface 7 with the spacer 6 adjacent to the disk 1, the rotation center axis 1
A bolt hole 8 having an equal radial position from 7 is drilled, through which the stacking bolt 3 penetrates, and nuts 18 and 1
8'by distant piece 19, stub shaft 2
All the stages are fastened by tightening both ends through 0, and the disc 1 is restrained in the rotation axis direction. A labyrinth seal surface 21 is provided on the outer peripheral portion of the spacer 6 between the shroud 5 provided at the tip of the vane 4.

Further, in the turbine cooling system of the present invention, as shown in FIG. 2, the moving blade cooling medium supplied from the outside of the turbine structure is not cooled in the hollow shaft 22 arranged in the central hole of the disk 1. Pass in the state of spacer 6
First, the rotor blade 2 is cooled by being guided to the outer peripheral side of the disk 1 through the radial hole 23 provided in the. The medium after cooling passes through the slit 24 provided on the contact surface 7 between the spacer 1 and the distant piece 19 adjacent to the disk 1, and the radial positions from the rotation center axis 17 provided on the hub portion 15 are the same. It is guided to the axial hole 25. The medium after cooling is then guided to the stub shaft 20 and collected outside the turbine structure, so that the surface of the disk 1 on the inner peripheral side of the contact surface 7 is cooled by blades before and after cooling. The medium for use does not touch either.
Therefore, the thermal stress generated in the disk 1 is reduced by reducing the temperature gradient in the disk 1 as much as possible.

Further, FIG. 4 is a structural schematic view of the spacer 26 between the first and second stages in FIG. Two types of hole groups are formed in the spacer between the one- and two-stage disks: a radial hole 23 through which the medium before cooling the moving blade passes and an axial hole 27 through which the medium after cooling the moving blade passes. However, as a premise of the closed cooling system, since the media before and after cooling cannot contact each other, the radial holes 23 and the axial holes 27 are alternately arranged as shown in FIG.

In the embodiment shown in FIG. 5, the flow of the cooling medium is the same as that in the embodiment shown in FIG. 2, but the wheel 28 is used to convey the cooling medium between the hollow shaft 22 and the spacer 6.
Is provided. However, the wheel 28 may be a pipe arranged in the radial direction between the hollow shaft 22 and the spacer 6.

In the embodiment shown in FIG. 6, the blade cooling medium supplied from the outside of the turbine structure passes through the hollow shaft 22 arranged in the central hole of the disk 1 in a state before cooling, and the spacer 6 First, the moving blade 2 is cooled by being guided to the outer peripheral side of the disk 1 through the radial hole 23 provided in the disk and the radial hole 29 provided in the distant piece 19. After cooling, the medium is the axial hole 2 provided in the spacer 6.
7 and the radial hole 23, and is guided to the axial hole 25 having the same radial position from the rotation center axis 17 provided on the contact surface 7. The medium after cooling is then guided to the stub shaft 20 and collected outside the turbine structure, so that the surface of the disk 1 on the inner peripheral side of the contact surface 7 is cooled by blades before and after cooling. The medium for use does not touch either. Therefore, the thermal stress generated in the disk 1 is reduced by reducing the temperature gradient in the disk 1 as much as possible.

In the embodiment shown in FIG. 7, the flow of the cooling medium is the same as that in the embodiment shown in FIG. 6, but the wheel 28 is used to convey the cooling medium between the hollow shaft 22 and the spacer 7 and the distant piece 19. Is provided. However, the wheel 28 may be a pipe arranged in the radial direction between the hollow shaft 22 and the spacer 6.

In the embodiment shown in FIG. 8, the blade cooling medium supplied from the outside of the turbine structure passes through the hollow shaft 22 arranged in the center hole of the disk 1 in a state before cooling and passes through the spacer 6 First, the moving blade 2 is cooled by being guided to the outer peripheral side of the disk 1 through the radial hole 23 provided in the disk and the radial hole 29 provided in the distant piece 19. After cooling, the medium is the axial hole 2 provided in the spacer 6.
7 and the radial hole 23, is guided to the hollow shaft 22 provided in the central portion, and is recovered to the outside of the turbine structure.
The surface of the blade is not touched by the blade cooling medium before and after cooling. Therefore, the thermal stress generated in the disk 1 is reduced by reducing the temperature gradient in the disk 1 as much as possible.

In the embodiment shown in FIG. 9, the flow of the cooling medium is the same as that in the embodiment shown in FIG. 8, but the wheel 28 is used to convey the cooling medium between the hollow shaft 22 and the spacer 6 and the distant piece 19. Is provided. However, the wheel 28 may be a pipe arranged in the radial direction between the hollow shaft 22 and the spacer 6.

In the embodiment shown in FIGS. 8 and 9, the moving blade cooling medium passes through the hollow shaft 22 disposed in the disk center hole both before and after cooling, but the medium before and after cooling does not touch each other. In addition, as shown in FIG. 10, it is possible to adopt a structure in which a partition plate 30 is provided inside the hollow shaft 22 or several dividing pipes 31 are put together.

However, in each of the above embodiments, the flow of the cooling medium may be in the opposite direction to the arrow in the drawing, that is, the dotted line before cooling the moving blade and the two-dot chain line after cooling the moving blade.

[0021]

According to the present invention, since the cooling medium is supplied to or collected from the moving blades by using the hollow shaft provided at the center and the spacer having the hole formed in the radial direction from the surface of the center hole, Neither the cooling medium before cooling nor the cooling medium for the moving blade touches the surface of the disk on the inner peripheral side of the contact surface with the spacer. Therefore, the temperature gradient inside the disk is reduced.
As a result, it is possible to reduce the thermal stress generated in the disk.

[Brief description of drawings]

FIG. 1 is a cross-sectional view of a turbine that employs a conventional closed cooling system.

FIG. 2 is a cross-sectional view of a turbine that best represents the present invention.

FIG. 3 is an explanatory view of a fir tree part for implanting a moving blade.

FIG. 4 is a layout view of various holes provided in the spacer between the first and second stages.

FIG. 5 is a cross-sectional view of a structure in which a cooling medium flows through a hollow shaft, a wheel and a spacer.

FIG. 6 is a cross-sectional view of a structure in which a cooling medium flows through a hollow shaft, a spacer and a distant piece.

FIG. 7 is a cross-sectional view of a structure in which a cooling medium flows through hollow shafts, wheels, spacers, and distant pieces.

FIG. 8 is a cross-sectional view of a structure in which a cooling medium flows through a hollow shaft and a spacer before and after cooling.

FIG. 9 is a sectional view of a structure in which a cooling medium flows through a hollow shaft, a wheel, and a spacer before and after cooling.

FIG. 10 is an explanatory diagram of a region inside the hollow shaft at the central portion.

[Explanation of symbols]

DESCRIPTION OF SYMBOLS 1 ... Disk, 2 ... Blade, 3 ... Stacking bolt, 4
... Static vane, 5 ... Shroud, 6 ... Spacer, 7 ... Abutting surface, 8 ... Bolt hole, 10 ... Four-stage disk, 11 ... Medium flow before cooling, 12 ... Medium flow after cooling, 15 ... Hub part, 16
... Inlay part, 17 ... Rotation center shaft, 18 ... Nut, 19
… Distant piece, 20… Stub shaft, 21…
Labyrinth seal surface, 22 ... Hollow shaft, 23 ... Spacer radial hole, 24 ... Slit, 25 ... Disc axial hole, 26 ... Spacer between 1 and 2 discs, 27 ... Spacer axial hole.

Claims (5)

[Claims] 1. A combination of a rotating disk and a plurality of moving blades fitted to the rotating disk is fastened by a multi-stage stacking bolt, and a medium for cooling the moving blade is provided with a turbine gas flow after cooling the moving blade. In a gas turbine that collects gas without discharging it to a passage, the cooling medium is supplied to or collected from the moving blades by using a hollow shaft provided at the center and a spacer having a hole formed in the radial direction from the center hole surface. A gas turbine cooling system characterized by being used. 2. The gas turbine cooling system according to claim 1, wherein the cooling medium is supplied to or collected from the moving blade by using a distant piece having a hole formed in a radial direction from a surface of the central hole. 3. The cooling medium according to claim 1, wherein the cooling medium is collected and supplied to the moving blade via a wheel having a hole in the radial direction between the hollow shaft and the spacer or the distant piece. Pass through gas turbine cooling system. 4. The method according to claim 1, wherein the cooling medium is collected and supplied to the moving blade via a pipe provided in a radial direction between the hollow shaft and a spacer or a distant piece. Gas turbine cooling system. 5. The cooling medium according to claim 1, 2, 3 or 4, wherein the hollow shaft is divided into a plurality of regions or a plurality of pipes are arranged in the hollow shaft. A gas turbine cooling system that has the functions of collecting and supplying simultaneously.