JPH0953409A - Ceramic stationary blade for gas turbine - Google Patents

Ceramic stationary blade for gas turbine

Info

Publication number
JPH0953409A
JPH0953409A JP20807795A JP20807795A JPH0953409A JP H0953409 A JPH0953409 A JP H0953409A JP 20807795 A JP20807795 A JP 20807795A JP 20807795 A JP20807795 A JP 20807795A JP H0953409 A JPH0953409 A JP H0953409A
Authority
JP
Japan
Prior art keywords
section
ceramic
insertion hole
blade
gas turbine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
JP20807795A
Other languages
Japanese (ja)
Inventor
Yoichiro Iritani
Takakuni Kasai
陽一郎 入谷
剛州 笠井
Original Assignee
Mitsubishi Heavy Ind Ltd
三菱重工業株式会社
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Mitsubishi Heavy Ind Ltd, 三菱重工業株式会社 filed Critical Mitsubishi Heavy Ind Ltd
Priority to JP20807795A priority Critical patent/JPH0953409A/en
Publication of JPH0953409A publication Critical patent/JPH0953409A/en
Application status is Withdrawn legal-status Critical

Links

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

【発明の詳細な説明】 DETAILED DESCRIPTION OF THE INVENTION

【0001】 [0001]

【発明の属する技術分野】本発明はガスタービン用セラミック静翼に関するものである。 BACKGROUND OF THE INVENTION The present invention relates to a ceramic stationary blade for a gas turbine.

【0002】 [0002]

【従来の技術】図3は従来のガスタービン用セラミック静翼の横断面図(図4の III−III 断面図)、図4は同静翼の縦断面図(図3のIV−IV断面図)である。 BACKGROUND ART FIG. 3 is a cross-sectional view of the ceramic stator vanes for conventional gas turbines (III-III sectional view of FIG. 4), FIG. 4 is a longitudinal sectional view of the vane (IV-IV cross-sectional view of FIG. 3 ) it is. これらの図において、前後方向および縦方向は矢印で定義してある。 In these drawings, the front-rear direction and the vertical direction are defined by arrows. 図3において、1はセラミック製の翼部、12は同翼部の前半部において縦方向に貫通している芯金挿通孔、4は同挿通孔に挿通されている芯金、10は同芯金と前記翼部との間に充填されている断熱材、11は前記芯金に縦方向に貫通している複数の冷却空気孔である。 3, 1 wings made of ceramic, 12 core metal insertion hole extending through in the longitudinal direction in the first half portion of the blade portion, 4 is inserted into the insertion hole core metal 10 coaxially heat insulating material filled in between the gold and the wings, 11 is a plurality of cooling air holes extending through the longitudinal direction to the metal core.
図4において、このセラミック静翼は図の下部をガスタービンの内側に、図の上部をガスタービンの外側にしてガスタービンに取付けられる。 4, the ceramic stator vanes inside the lower gas turbine of FIG attached to the gas turbine to the top of the figure to the outside of the gas turbine. 以下の部材の名称における「内側」「外側」というのは上記ガスタービン内での設置位置における「内側」「外側」に対応している。 It is because "inner", "outer" in the name of the following member corresponds to the "inner", "outer" in the installation position in the gas turbine. 2
は翼部の内側に設けられているセラミック製内側シュラウド、3は翼部の外側に設けられているセラミック製外側シュラウド、5は前記シュラウド2に隣接して設けられている金属製内側シュラウド、6はシュラウド3に隣接して設けられている金属製外側シュラウド、7はスペーサ、9はバネ、8はナットである。 Ceramic inner shroud is provided on the inner side of the wings, a ceramic outer shroud 3 is provided outside of the wings, 5 metallic inner shroud is provided adjacent to the shroud 2, 6 the metallic outer shroud is provided adjacent to the shroud 3, 7 spacer, 9 a spring, 8 is a nut.

【0003】セラミック静翼においては、翼根の熱応力を緩和するために、セラミック製翼部1と内外のセラミック製シュラウド2,3とに3分割して作られ、それを組立てるために、金属製の芯金4を前記芯金挿通孔12 [0003] In the ceramic vanes, in order to mitigate the thermal stress of the blade root is made in 3 divided into a ceramic blade portion 1 and the inside and outside of the ceramic shroud 2, to assemble it, metal wherein the manufacturing of the core metal 4 metal core insertion hole 12
に挿通し、その両端に取り付けられた内外の金属製シュラウド5,6、およびスペーサ7、バネ9を介してナット8で締付け一体化する構造となっている。 Inserted into, and out of metal shroud 5 and 6 mounted on opposite ends thereof, and the spacer 7, through the spring 9 has a structure of integrating tightening a nut 8.

【0004】 [0004]

【発明が解決しようとする課題】図3のAは芯金挿通孔の後端部である。 A of FIG. 3 [SUMMARY OF THE INVENTION] is a rear end of the metal core insertion hole. 翼断面において、このA部より後方の部分の肉厚は、A部より前方の部分の肉厚より大きい。 In blade section, the thickness of the rear portion from the part A is greater than the thickness of the forward section from A unit.
ガスタービントリップ時にはセラミック静翼が約130 At the time of gas turbine trip ceramic stator blade is about 130
0℃から約400℃へ急冷されるので、この肉厚の違いによる熱容量の違いによって、冷却速度の違いが生じ、 0 because it is quenched from ° C. to about 400 ° C., the difference in heat capacity due to the thick differences, resulting difference in cooling rate,
A部を境にして、前方の温度に比して後方の温度が高いという状態が生じる。 And the A unit as a boundary, the state occurs that the temperature of the rear is higher than the front of the temperature. これによって、A部に熱応力の集中が生じる。 Thus, concentration of thermal stress in the A section. 熱応力が40〜50kg/mm 2以上になるとクラックが発生し事故の原因となる。 Crack when a thermal stress is 40~50kg / mm 2 or more causes of occurred accident.

【0005】本発明は上記従来技術の欠点を解消し、芯金挿通孔の後端部Aにおける熱応力の発生を無くし、クラックの発生を防止しようとするものである。 The present invention eliminates the disadvantages of the prior art, eliminate the occurrence of thermal stress in the rear end portion A of the core insertion hole, it is intended to prevent the occurrence of cracks.

【0006】 [0006]

【課題を解決するための手段】本発明は上記課題を解決したものであって、その前半部に縦方向の芯金挿通孔を有するセラミック製翼部と同翼部の両端部に設けられるセラミック製シュラウドとを、上記芯金挿通孔に芯金を挿通し同芯金の両端部に設けられた金属製シュラウドを介して締め付け一体化して構成されるガスタービン用セラミック静翼において、上記セラミック製翼部の後半部即ち芯金挿通孔の設けられていない部分に複数個の縦方向の円孔を設けたことを特徴とするガスタービン用セラミック静翼に関するものである。 The present invention SUMMARY OF] is a solves the above problems, a ceramic which is provided at both ends of the ceramic blade portion and the wing portion having a longitudinal metal core insertion hole in its front half portion and manufacturing the shroud, in the ceramic stationary blade for formed a gas turbine integrated by tightening via the metal shroud provided at both ends of the Duasyn gold inserted through the core metal to the metal core insertion hole, the ceramic it relates ceramic stationary blade for a gas turbine, characterized in that in the latter half portion or portions provided with no core metal insertion hole of the blade portion is provided a plurality of longitudinal circular hole.

【0007】 [0007]

【発明の実施の形態】図1は本発明の実施の一形態に係るガスタービン用セラミック静翼の横断面図(図2のI DESCRIPTION OF THE PREFERRED EMBODIMENTS Figure 1 is a cross-sectional view of the ceramic stationary blade for a gas turbine according to an embodiment of the present invention (in FIG. 2 I
−I断面図)、図2は同静翼の縦断面図(図1のII−II -I sectional view), FIG. 2 is a longitudinal sectional view of the vane (FIG. 1 II-II
断面図)である。 It is a cross-sectional view). 図において13はセラミック製翼部1 13 In figure ceramic blade portion 1
の芯金挿通孔の後端部Aより後方の肉厚部に縦方向に貫通して設けられた複数個の円孔である。 A metal core insertion hole plurality of circular holes of the rear end A were provided through longitudinally thick portion of the rear. 上記以外の部分の構成は従来技術と同じであるから構成の説明を省略する。 Structure of a portion other than the above will be omitted in the configuration because it is the same as the prior art.

【0008】上記静翼は、A部より後方の厚肉部の内部に空洞を形成して実質的に肉厚を減少させ、熱容量を低減させたものであり、これによってガスタービントリップ時におけるA部の前後の冷却速度をほぼ等しくすることができ、温度差によってA部に生じる熱応力を20kg [0008] The vanes, substantially reduces the wall thickness by forming a cavity inside the thick portion of the rear of the part A, which reduced the heat capacity, whereby A when the gas turbine trip before and after the cooling rate of the part can be a substantially equal, the heat stress generated in the part a by the temperature difference 20kg
/mm 2以下に低減させることができ、クラックの発生を防止することができる。 / mm 2 can be reduced below, it is possible to prevent the occurrence of cracks.

【0009】 [0009]

【発明の効果】本発明のガスタービン用セラミック静翼においては、セラミック製翼部の後半部即ち芯金挿通孔の設けられていない部分に複数個の縦方向の円孔を設けてあるので、芯金挿通孔の後端部における熱応力の発生を解消し、クラックの発生を防止することができる。 In the ceramic stationary blade for a gas turbine of the present invention, since is provided a plurality of longitudinal circular holes in the second half portion or portions provided with no core metal insertion hole of the ceramic blade portion, eliminating the occurrence of thermal stress in the rear end portion of the core insertion hole, it is possible to prevent the occurrence of cracks.

【図面の簡単な説明】 BRIEF DESCRIPTION OF THE DRAWINGS

【図1】本発明の実施の一形態に係るガスタービン用セラミック静翼の横断面図(図2のI−I断面図)。 Figure 1 is a cross-sectional view of the ceramic stationary blade for a gas turbine according to an embodiment of the present invention (I-I section view of FIG. 2).

【図2】同静翼の縦断面図(図1のII−II断面図)。 Figure 2 is a longitudinal sectional view of the vane (II-II cross-sectional view of FIG. 1).

【図3】従来のガスタービン用セラミック静翼の横断面図(図4の III−III 断面図)。 Figure 3 is a cross-sectional view of a conventional ceramic stationary blade for a gas turbine (III-III sectional view of FIG. 4).

【図4】同静翼の縦断面図(図3のIV−IV断面図)。 Figure 4 is a longitudinal sectional view of the vane (IV-IV cross-sectional view of FIG. 3).

【符号の説明】 DESCRIPTION OF SYMBOLS

1 セラミック製翼部 2 セラミック製内側シュラウド 3 セラミック製外側シュラウド 4 芯金 5 金属製内側シュラウド 6 金属製外側シュラウド 7 スペーサ 8 ナット 9 バネ 10 断熱材 11 冷却空気孔 12 芯金挿通孔 13 円孔 A 芯金挿通孔の後端部 1 ceramic blade portion 2 ceramic inner shroud 3 ceramic outer shroud 4 core metal 5 metal inner shroud 6 metal outer shroud 7 spacer 8 Nut 9 spring 10 heat insulating material 11 cooling air holes 12 metal core insertion hole 13 circular hole A the rear end portion of the core insertion hole

Claims (1)

    【特許請求の範囲】 [The claims]
  1. 【請求項1】 その前半部に縦方向の芯金挿通孔を有するセラミック製翼部と同翼部の両端部に設けられるセラミック製シュラウドとを、上記芯金挿通孔に芯金を挿通し同芯金の両端部に設けられた金属製シュラウドを介して締め付け一体化して構成されるガスタービン用セラミック静翼において、上記セラミック製翼部の後半部即ち芯金挿通孔の設けられていない部分に複数個の縦方向の円孔を設けたことを特徴とするガスタービン用セラミック静翼。 1. A a ceramic shroud provided at both end portions of the ceramic blade portion and the wing portion having a longitudinal metal core insertion hole in its front half portion, the inserted core metal on the core metal insertion hole in the ceramic stationary blade for formed a gas turbine integrated by tightening via the metal shroud provided at both ends of the core, in the latter half portion or portions provided with no core metal insertion hole of the ceramic blade portion ceramic stationary blade for a gas turbine, characterized in that a plurality of longitudinal circular hole.
JP20807795A 1995-08-15 1995-08-15 Ceramic stationary blade for gas turbine Withdrawn JPH0953409A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
JP20807795A JPH0953409A (en) 1995-08-15 1995-08-15 Ceramic stationary blade for gas turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
JP20807795A JPH0953409A (en) 1995-08-15 1995-08-15 Ceramic stationary blade for gas turbine

Publications (1)

Publication Number Publication Date
JPH0953409A true JPH0953409A (en) 1997-02-25

Family

ID=16550267

Family Applications (1)

Application Number Title Priority Date Filing Date
JP20807795A Withdrawn JPH0953409A (en) 1995-08-15 1995-08-15 Ceramic stationary blade for gas turbine

Country Status (1)

Country Link
JP (1) JPH0953409A (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2361921A2 (en) 1997-03-07 2011-08-31 Exiqon A/S Novel bicyclonucleoside and oligonnucleotide analogue

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2361921A2 (en) 1997-03-07 2011-08-31 Exiqon A/S Novel bicyclonucleoside and oligonnucleotide analogue
EP1013661B2 (en) 1997-03-07 2018-10-24 Exiqon A/S 2'-O,4'-C-methylene bicyclonucleosides

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A300 Withdrawal of application because of no request for examination

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Effective date: 20021105