JPH0953409A - Ceramic stationary blade for gas turbine - Google Patents

Abstract

PROBLEM TO BE SOLVED: To resolve the occurrence of thermal stress at the rear end section of a mandrel insertion hole by providing multiple circular holes in the vertical direction at the rear half section of a ceramic blade section, i.e., the portion where no mandrel insertion hole is provided. SOLUTION: A ceramic stationary blade for a gas turbine is provided with ceramic blade sections 1..., a mandrel insertion hole 12 penetrating in the vertical direction is formed at the front half section of each blade section 1, a mandrel 4 is inserted into the mandrel insertion hole 12 via a heat insulating material 10, and cooling air holes 11... penetrating in the vertical direction are formed in the mandrel 4. Multiple circular holes 13 penetrating in the vertical direction are formed at a thick section behind the rear end section A of the mandrel insertion hole 12 of the blade section 1, cavities are formed in the thick section behind the section A, the thickness is practically reduced, and the heat capacity is reduced. The cooling speeds at the front and rear of the section A are made nearly equal at the time of a gas turbine trip, and the occurrence of cracks based on the thermal stress generated at the section A due to a temperature difference is prevented.

Description

Detailed Description of the Invention

[0001]

TECHNICAL FIELD The present invention relates to a ceramic vane for a gas turbine.

[0002]

2. Description of the Related Art FIG. 3 is a cross-sectional view of a conventional ceramic vane for a gas turbine (III-III cross-sectional view of FIG. 4), and FIG. 4 is a vertical cross-sectional view of the same vane (IV-IV cross-sectional view of FIG. 3). ). In these figures, the front-back direction and the vertical direction are defined by arrows. In FIG. 3, reference numeral 1 is a ceramic wing portion, 12 is a mandrel insertion hole vertically extending through the front half of the wing portion, 4 is a mandrel inserted in the insertion hole, and 10 is a concentric core. A heat insulating material filled between the gold and the blade portion, and 11 are a plurality of cooling air holes penetrating the core metal in the longitudinal direction.
In FIG. 4, the ceramic vane is attached to the gas turbine with the lower part of the figure inside the gas turbine and the upper part of the figure outside the gas turbine. The terms "inside" and "outside" in the names of the following members correspond to "inside" and "outside" at the installation position in the gas turbine. Two
Is a ceramic inner shroud provided inside the blade portion, 3 is a ceramic outer shroud provided outside the blade portion, 5 is a metal inner shroud provided adjacent to the shroud 2, and 6 Is a metal outer shroud provided adjacent to the shroud 3, 7 is a spacer, 9 is a spring, and 8 is a nut.

In order to relieve the thermal stress at the root of the ceramic vane, the vane 1 is made of ceramic and the shrouds 2 and 3 made of ceramics inside and outside are divided into three parts. The cored bar 4 made of metal
It has a structure in which it is tightened with a nut 8 through inner and outer metal shrouds 5 and 6 attached to both ends thereof, a spacer 7, and a spring 9 to be integrated.

[0004]

FIG. 3A shows the rear end portion of the cored bar insertion hole. In the blade cross section, the thickness of the portion behind the A portion is larger than the thickness of the front portion of the A portion.
When the gas turbine trips, about 130 ceramic vanes
Since it is rapidly cooled from 0 ℃ to about 400 ℃, the difference in heat capacity due to the difference in wall thickness causes the difference in cooling rate,
A state in which the temperature of the rear side is higher than the temperature of the front side with respect to the portion A as a boundary occurs. As a result, thermal stress is concentrated on the A portion. If the thermal stress exceeds 40 to 50 kg / mm ² , cracks may occur and cause an accident.

The present invention is intended to solve the above-mentioned drawbacks of the prior art, eliminate the occurrence of thermal stress at the rear end portion A of the cored bar insertion hole, and prevent the occurrence of cracks.

[0006]

DISCLOSURE OF THE INVENTION The present invention has been made to solve the above-mentioned problems, and has a ceramic wing portion having a vertical mandrel insertion hole in the front half thereof and ceramics provided at both ends of the wing portion. In a ceramic turbine vane for a gas turbine, which is configured by inserting a core shroud into the core bar insertion hole and tightening the metal shrouds through metal shrouds provided at both ends of the core bar, the ceramic shroud The present invention relates to a ceramic turbine vane for a gas turbine, characterized in that a plurality of longitudinal circular holes are provided in the latter half of the blade portion, that is, in the portion where the core metal insertion hole is not provided.

[0007]

1 is a cross-sectional view of a ceramic turbine vane for a gas turbine according to an embodiment of the present invention (I in FIG. 2).
-I sectional view), and Fig. 2 is a vertical sectional view of the same vane (II-II in Fig. 1).
FIG. In the figure, 13 is a ceramic wing 1.
Is a plurality of circular holes provided in the thick portion rearward of the rear end portion A of the core metal insertion hole so as to penetrate in the longitudinal direction. The configuration other than the above is the same as that of the conventional technique, and therefore the description of the configuration is omitted.

The stator vane has a cavity formed inside the thick portion rearward of the portion A to substantially reduce the thickness and reduce the heat capacity. The cooling rate before and after the part can be made almost equal, and the thermal stress generated in part A due to the temperature difference is 20 kg.
/ mm ² or less, and it is possible to prevent the occurrence of cracks.

[0009]

In the ceramic vane for a gas turbine of the present invention, since a plurality of longitudinal circular holes are provided in the latter half of the ceramic blade, that is, in the portion where the core metal insertion hole is not provided, It is possible to eliminate the occurrence of thermal stress at the rear end of the cored bar insertion hole and prevent the occurrence of cracks.

[Brief description of drawings]

FIG. 1 is a cross-sectional view (cross-sectional view taken along the line I-I of FIG. 2) of a ceramic vane for a gas turbine according to an embodiment of the present invention.

FIG. 2 is a longitudinal sectional view of the stationary vane (II-II sectional view of FIG. 1).

FIG. 3 is a transverse cross-sectional view of a conventional ceramic vane for a gas turbine (III-III cross-sectional view of FIG. 4).

FIG. 4 is a vertical sectional view of the same vane (IV-IV sectional view of FIG. 3).

[Explanation of symbols]

 1 Ceramic Blade 2 Ceramic Inner Shroud 3 Ceramic Outer Shroud 4 Core Bar 5 Metal Inner Shroud 6 Metal Outer Shroud 7 Spacer 8 Nut 9 Spring 10 Heat Insulation Material 11 Cooling Air Hole 12 Core Bar Insertion Hole 13 Circular Hole A Rear end of core metal insertion hole

Claims (1)

[Claims] 1. A ceramic wing part having a vertical mandrel insertion hole in the front half thereof and ceramic shrouds provided at both ends of the wing part, and a cored bar inserted in the mandrel insertion hole. In a ceramic turbine vane for a gas turbine, which is integrally tightened through metal shrouds provided at both ends of the core metal, in the latter half of the ceramic blade portion, that is, in a portion where the core metal insertion hole is not provided. A ceramic turbine vane for a gas turbine, which is provided with a plurality of vertical holes.