JP2004316542A - Turbine nozzle cooling structure for gas turbine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a turbine nozzle cooling structure capable of effectively cooling a band part without giving the part a double plate structure and thereby, attaining cost reduction and improvement of reliability, and improving engine performance by reducing weight, in a gas turbine in which turbine inlet temperature is high. <P>SOLUTION: A turbine nozzle 10 which is positioned between a combustor 2 and a turbine moving blade 4 and introduces high temperature gas from the combustor 2 to a turbine in a downstream side and a seal plate 8 for sealing a space between a front end part of the turbine nozzle 10 and the combustor 2 are provided. A flange part 2a of the combustor 2 contacted with the seal plate 8 has a plurality of notch grooves 20 communicating the inner and outer faces. Through these notch grooves 20, cooling air flows in and along a band surface of the turbine nozzle 10, performing film cooling of the surface. <P>COPYRIGHT: (C)2005,JPO&NCIPI

Description

[0001]
TECHNICAL FIELD OF THE INVENTION
The present invention relates to a turbine nozzle cooling structure for a gas turbine.
[0002]
[Prior art]
As shown in FIG. 4A of [Non-Patent Document 1], the turbine nozzle is formed by arranging nozzle guide vanes (stationary vanes) in a ring shape, and includes a vane and a supporting structure thereof. An air-cooled turbine vane shown in FIG. 4B is also known.
The cooling of the turbine nozzle is described in, for example, [Non-Patent Document 2].
[0003]
[Non-patent document 1]
New Aviation Engineering Course 8, Jet Engine (Structure), Japan Aviation Technology Association [Non-Patent Document 2]
Kenichiro Takeishi, Turbine Loss Generation Mechanism and Thermodynamic Considerations, Proceedings of the TED-COF. '01, JSME
[0004]
The turbine nozzle has a function of accelerating high-temperature gas from the combustor and introducing the gas to a downstream turbine. Therefore, cooling is indispensable because the turbine nozzle is exposed to the hot gas in the engine flow path.
[0005]
In order to satisfy this demand, for example, a “gas turbine turbine cooling vane” of Patent Document 1 has been proposed.
[0006]
[Patent Document 1]
JP 2001-254604 A
As shown in FIG. 5, the “gas turbine turbine cooling vane” of [Patent Document 1] has a lattice-like waffle pattern 101 formed on the outside wall of the outside shroud 121 of the first stage vane 120 and on the inside wall surface of the inside shroud. The cooling holes in the last row of the trailing edge of the one-stage stationary blade 120 are enlarged cooling holes 106 larger than other hole diameters to increase the cooling efficiency of the trailing edge, and the ribs 102 are formed inside the wall to reduce the wall thickness. Cooling passages 103 are provided inside the rear and ventral side ends of the inner shroud 122, and a plurality of cooling holes 105 penetrating from the inside and opening to the surface end are provided to enhance the cooling effect of the shroud. These improvements prevent cracking and deformation of the trailing edge of the blade and the shroud.
[0008]
[Problems to be solved by the invention]
As described above, in a conventional gas turbine having a high turbine inlet temperature, cooling of an inner shroud (also referred to as a “band portion”) of a turbine nozzle (first-stage stationary blade) is indispensable, and the band portion is formed as a double plate structure inside. It was cooled by passing cooling air.
[0009]
However, the two-plate structure for cooling the band portion has a problem that not only the structure becomes complicated and the cost is increased, but also the reliability is reduced, and the engine performance is reduced due to the increased weight.
[0010]
The present invention has been made to solve the above problems. In other words, an object of the present invention is to provide a gas turbine having a high turbine inlet temperature, in which the band portion can be effectively cooled without forming the band portion with a double plate structure. It is an object of the present invention to provide a turbine nozzle cooling structure of a gas turbine that can achieve down and reliability improvement, and can improve engine performance by weight reduction.
[0011]
[Means for Solving the Problems]
ADVANTAGE OF THE INVENTION According to this invention, the turbine nozzle located between a combustor and a turbine blade, and introduce | transducing hot gas from a combustor to a downstream turbine, and sealing between the front end part of this turbine nozzle and a combustor A plurality of notch grooves communicating the inner and outer surfaces of the flange portion of the combustor in contact with the seal plate, and cooling air flows through the notch grooves and flows along the band surface of the turbine nozzle. And cooling the surface thereof by film cooling.
[0012]
According to the configuration of the present invention, since the plurality of notch grooves communicating the inner and outer surfaces are provided in the flange portion of the combustor in contact with the seal plate, the cooling air flows in from the outside through the notch grooves and the turbine It can flow along the band surface of the nozzle and the surface can be film cooled.
Therefore, in a gas turbine having a high turbine inlet temperature, the band portion can be effectively cooled without forming the band portion with a double plate structure, thereby reducing costs and improving reliability by simplifying the structure. And engine performance can be improved by weight reduction.
[0013]
BEST MODE FOR CARRYING OUT THE INVENTION
Hereinafter, preferred embodiments of the present invention will be described with reference to the drawings. In each of the drawings, common portions are denoted by the same reference numerals, and redundant description is omitted.
[0014]
FIG. 1 is an overall cross-sectional view of a turbine nozzle portion according to the cooling structure of the present invention. As shown in this figure, the turbine nozzle cooling structure of the present invention includes a combustor 2 and a turbine nozzle 10.
[0015]
The turbine nozzle 10 is located between the combustor 2 and the turbine blade 4 and has a function of introducing high-temperature gas from the combustor 2 to a downstream turbine.
[0016]
In FIG. 1, the turbine nozzle cooling structure of the present invention further includes a nozzle support member 14 and a stopper member 16.
The nozzle support member 14 is a rotating body centered on the center of the engine fixed inside the turbine nozzle 10 and has a radially extending groove 14a around which the turbine nozzle 10 is fitted. In addition, a plurality of axial through holes 14b are provided in the front and rear flange portions sandwiching the groove portion 14a at regular intervals in the circumferential direction.
[0017]
The stopper member 16 is located outside the turbine nozzle 10 and has an outer end fixed to the casing 6 of the engine.
[0018]
FIG. 2 is a perspective view of a nozzle segment constituting the present invention. As shown in FIG. 1, the turbine nozzle 10 includes a plurality of nozzle segments 12 divided in a circumferential direction. The circumferential end face of each nozzle segment 12 is in close contact with the end face of each adjacent nozzle segment 12, and the space therebetween is sealed by a seal plate (not shown) fitted into a groove 11 provided on the end face.
[0019]
The nozzle segment 12 includes an inner shroud 12a, an outer shroud 12b, and a wing 12c. The inner shroud 12a and the outer shroud 12b are also referred to as an inner band portion and an outer band portion, respectively, in the present invention.
[0020]
The nozzle segment 12 has an inner flange 15a extending radially inward from an inner shroud 12a (inner band portion) and an outer flange 15b extending radially outward from an outer shroud 12b (outer band portion). . The outer flange 15b is fitted in the groove 14a of the nozzle support member 14 so as not to move in the front-rear direction.
[0021]
Further, a single axial through hole 13 is provided in the inner flange 15 a, and the mounting segment 18 passes through the axial through hole 13 together with the groove 14 a of the nozzle support member 14, so that the nozzle segment 12 passes through the nozzle support member 14. Mounted.
[0022]
In FIG. 1, the stopper member 16 is a ring-shaped member centered on the center of the engine. In addition, fitting portions 16b for positioning in the circumferential direction with the casing 6 are provided at a plurality of locations on the outer peripheral edge. The fitting portion 16b is a rectangular member projecting in the axial direction, and positions the stopper member 16 in the circumferential direction by fitting with a concave portion provided in the casing 6.
Further, the stopper member 16 has a plurality of recesses 16a provided with a step on the inner peripheral edge and provided in the circumferential direction along the inner peripheral edge. The recess 16 a is provided at a location corresponding to each of the plurality of divided nozzle segments 12.
[0023]
In FIG. 1, in the turbine nozzle cooling structure of the present invention, a seal plate 8 that seals between the front end of the nozzle segment 12 and the combustor 2, and a space between the rear end of the nozzle segment 12 and the casing 6. An E-seal 9 for sealing is provided, and by sealing the front and rear of the nozzle segment 12, the sealing performance between the inside and the outside of the nozzle is maintained.
[0024]
FIG. 3 is a view taken in the direction of arrows AA in FIG. 1. As shown in this figure, in the present invention, a plurality of notch grooves 20 are provided in the flange portion 2a of the combustor 2 that comes into contact with the seal plate 8 and communicate the inner and outer surfaces. The notch grooves 20 allow cooling air to flow in from the outside of the combustor 2 through the grooves, flow along the band surface of the inner shroud 12 a (inner band portion) of the turbine nozzle 12, and cool the surface by film cooling. Has become.
[0025]
The cross-sectional shape and direction of the notch groove 20 are arbitrary, may be semicircular or rectangular, and may be radial or oblique. Further, the size of the notch groove 20 is small enough to form a film cooling film flowing along the band surface of the inner band portion and prevent overheating of the inner band portion, so as not to affect the turbine inlet temperature of the mainstream gas. Set to.
[0026]
According to the configuration of the present invention, since the plurality of notch grooves (20) communicating the inner and outer surfaces are provided in the flange portion of the combustor (2) in contact with the seal plate (8), the notch grooves are provided. Cooling air flows in from outside and flows along the band surface of the turbine nozzle to cool the surface.
Therefore, in a gas turbine having a high turbine inlet temperature, the band can be effectively cooled without having a double plate structure in the band, whereby the structure is simplified and the cost is reduced and the reliability is improved. And engine performance can be improved by weight reduction.
[0027]
It should be noted that the present invention is not limited to the above-described embodiment, and can be freely modified without departing from the gist of the present invention. For example, the present invention can be widely applied to aviation, marine, and land gas turbines.
[0028]
【The invention's effect】
According to the present invention described above, the band cooling structure of the stationary blade itself becomes unnecessary, and the following effects can be expected.
(1) Improvement of reliability by simplification of structure (2) Cost reduction by simplification of structure (3) Improvement of engine performance by weight reduction
Therefore, the turbine nozzle cooling structure of the present invention can effectively cool the band portion of the gas turbine having a high turbine inlet temperature without having the band portion have a double plate structure. The present invention has excellent effects such as achieving cost reduction and improvement in reliability due to the reduction of the weight, and improving the engine performance by reducing the weight.
[Brief description of the drawings]
FIG. 1 is an overall cross-sectional view of a turbine nozzle portion according to a cooling structure of the present invention.
FIG. 2 is a perspective view of a nozzle segment constituting the present invention.
FIG. 3 is a view as viewed in the direction of arrows AA in FIG. 1;
FIG. 4 is a perspective view of a conventional turbine nozzle structure.
FIG. 5 is an overall sectional view showing an example of a conventional turbine nozzle cooling structure.
[Explanation of symbols]
2 combustor, 2a flange part, 4 turbine rotor blade,
6 casing, 8 seal plate, 9 E seal,
10 turbine nozzles, 12 nozzle segments,
12a inner shroud (inner band part),
12b outer shroud (outer band part),
12c wing, 13 through hole,
14 nozzle support member, 15a inner flange,
15b outer flange, 15c tab,
16 stopper member,
16a recess, 16b fitting part,
18 mounting pins, 20 notch grooves

Claims (1)

A turbine nozzle positioned between the combustor and the turbine blade to introduce hot gas from the combustor to a downstream turbine; and a seal plate for sealing between a front end of the turbine nozzle and the combustor. ,
A plurality of notch grooves communicating the inner and outer surfaces are provided in a flange portion of the combustor in contact with the seal plate, and cooling air flows in through the notch grooves, flows along the band surface of the turbine nozzle, and cools the surface. A turbine nozzle cooling structure for a gas turbine.

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