JPS63223302A - Ceramics stationary blade for gas turbine - Google Patents
Ceramics stationary blade for gas turbineInfo
- Publication number
- JPS63223302A JPS63223302A JP5659287A JP5659287A JPS63223302A JP S63223302 A JPS63223302 A JP S63223302A JP 5659287 A JP5659287 A JP 5659287A JP 5659287 A JP5659287 A JP 5659287A JP S63223302 A JPS63223302 A JP S63223302A
- Authority
- JP
- Japan
- Prior art keywords
- blade
- lower side
- ceramic
- side wall
- 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.)
- Pending
Links
- 239000000919 ceramic Substances 0.000 title claims abstract description 21
- 238000006073 displacement reaction Methods 0.000 claims abstract description 5
- 239000000567 combustion gas Substances 0.000 claims description 8
- 239000007789 gas Substances 0.000 claims description 7
- 238000010276 construction Methods 0.000 abstract 1
- 230000008642 heat stress Effects 0.000 abstract 1
- 239000002184 metal Substances 0.000 description 11
- 229910052751 metal Inorganic materials 0.000 description 11
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical compound [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 description 10
- 229910010271 silicon carbide Inorganic materials 0.000 description 10
- 238000001816 cooling Methods 0.000 description 8
- 239000011810 insulating material Substances 0.000 description 6
- 230000008646 thermal stress Effects 0.000 description 4
- 230000000694 effects Effects 0.000 description 3
- 230000035882 stress Effects 0.000 description 3
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 2
- 238000010438 heat treatment Methods 0.000 description 2
- 230000003647 oxidation Effects 0.000 description 2
- 238000007254 oxidation reaction Methods 0.000 description 2
- 239000000956 alloy Substances 0.000 description 1
- 229910045601 alloy Inorganic materials 0.000 description 1
- 230000015572 biosynthetic process Effects 0.000 description 1
- 230000007797 corrosion Effects 0.000 description 1
- 238000005260 corrosion Methods 0.000 description 1
- 230000006378 damage Effects 0.000 description 1
- 239000000835 fiber Substances 0.000 description 1
- 238000000227 grinding Methods 0.000 description 1
- 238000009413 insulation Methods 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 229910052759 nickel Inorganic materials 0.000 description 1
- 230000000630 rising effect Effects 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- 238000005245 sintering Methods 0.000 description 1
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【発明の詳細な説明】 〔産業上の利用分野〕 本発明はガスタービン用セラミック静翼に係り。[Detailed description of the invention] [Industrial application field] The present invention relates to a ceramic stator blade for a gas turbine.
特に構造の簡単なセラミック静翼に関する。In particular, it relates to ceramic stator blades with a simple structure.
耐熱ｌ９１１製のケーシングに組込んで用いられる従来
のセラミック静翼は、特開昭５４−１０６７１４号に記
載のように、中央翼の内部を貫通して金属コアを設置し
、この金属コアによって翼とその上下に配置したサイド
ウオールを一体化する構成となしていた。この構成の特
徴は、耐熱性のセラミック翼とサイドウオールで断熱と
流路形成を行い、金属コアで負荷を分担するように機能
を分けた点にある。Conventional ceramic stator blades that are used by being assembled into a casing made of heat-resistant l911 have a metal core installed through the center blade, as described in Japanese Patent Application Laid-open No. 54-106714. It was designed to integrate the sidewalls placed above and below it. The feature of this configuration is that the heat-resistant ceramic blades and sidewalls provide insulation and flow path formation, while the metal core divides the functions to share the load.
この従来技術においては、高温の燃焼ガスの下で金属コ
アの強度や耐食性を確保するため、金属コア内を空冷し
たり、翼と金属コアの間に断熱材を設置する必要がある
。金属コア内の冷却空気は燃焼ガス中に放出されるため
ガスタービンの効率を低下させる要因となる。また、こ
の冷却空気によって翼内面やサイドウオールの片面を冷
却すると、燃焼ガスに接する他の部分との間に温度差を
生じ、これらセラミック部材内に熱応力を発生させるこ
とになる。さらに、この空気の流路が閉塞するなどして
十分な冷却が行なわれない場合には、金属コアの変形や
酸化が進むことが考えられる。In this conventional technology, in order to ensure the strength and corrosion resistance of the metal core under high-temperature combustion gas, it is necessary to air-cool the inside of the metal core or install a heat insulating material between the blades and the metal core. The cooling air within the metal core is released into the combustion gases and is a factor that reduces the efficiency of the gas turbine. Furthermore, when the inner surface of the blade or one side of the sidewall is cooled by this cooling air, a temperature difference is generated between the inner surface of the blade and one side of the sidewall that is in contact with the combustion gas, and thermal stress is generated within these ceramic members. Further, if sufficient cooling is not performed due to blockage of the air flow path, deformation and oxidation of the metal core may proceed.
本発明の目的は、金属コアとその冷却手段をなくして構
造の簡単なガスタービン用セラミック静翼を提供するこ
とにある。An object of the present invention is to provide a ceramic stator vane for a gas turbine that has a simple structure by eliminating a metal core and its cooling means.
上記目的は、耐熱セラミック製ボルトを翼内とサイドウ
オールの一つを貫通させて、耐熱セラミックばねを介し
ナツト締結することにより、達成される。The above object is achieved by passing a heat-resistant ceramic bolt through the wing and one of the sidewalls and fastening it with a nut via a heat-resistant ceramic spring.
燃焼ガスによって翼が加熱膨張することによる半径方向
の相対変化はセラミック製ばねにより吸収されるため、
翼とサイドウオール間に大きな熱応力が発生することは
防止でき、耐熱セラミック製ボルト、ナツト、ばねは冷
却を必要としない。The relative change in the radial direction due to the heating and expansion of the blades due to the combustion gases is absorbed by the ceramic springs.
Large thermal stresses between the wings and sidewalls are avoided, and the high-temperature ceramic bolts, nuts, and springs do not require cooling.
以下、本発明の一実施例を第１図及び第２図により説明
する。翼１は耐熱セラミックの炭化珪素製で、軸方向の
流路を形成し、内部が中空の柱状をなしている。上、下
サイドウオール板２，３は、それぞれ前記翼１の上下を
保持して半径方向の流路を形成するもので、同じく炭化
珪素製で中央に円孔を有する平行四辺形の板状をなして
いる。上。An embodiment of the present invention will be described below with reference to FIGS. 1 and 2. The blade 1 is made of silicon carbide, a heat-resistant ceramic, and has a columnar shape with an axial flow path and a hollow interior. The upper and lower sidewall plates 2 and 3 respectively hold the upper and lower sides of the blade 1 to form a radial flow path, and are also made of silicon carbide and have a parallelogram plate shape with a circular hole in the center. I am doing it. Up.
下サイドウオール４，５は同様の平行四辺形の縁形状を
有し中央に円孔を有するニッケル基耐熱合金製で、それ
ぞれ炭化珪素製シールリング６．７を介して前記上、下
サイドウオール板２，３の上。The lower sidewalls 4 and 5 are made of a nickel-based heat-resistant alloy with similar parallelogram edge shapes and a circular hole in the center, and are connected to the upper and lower sidewall plates through silicon carbide seal rings 6 and 7, respectively. Above 2 and 3.
下面に接している。これら上、下サイドウオール４．５
の上、下面には突起８，９が形成され、それらはそれぞ
れガスタービンケーシングのリテーナリング１０とサポ
ートリング１１に係合して翼体を固定する。ボルト１２
は炭化珪素製で焼結後にねじ部を精密研削され、下方の
下サイドウオール５外面に頭部が固定され、下サイドウ
オール５゜下サイドウオール板３．翼１．上サイドウオ
ール板２および上サイドウオール６の中央部および板ば
ね１３の中央部を貫通している。板ばね１３は炭化珪素
製で下面は精密研削で平坦に仕上げられ、同じく炭化珪
素製の平板座１４を介して上サイドウオール４の上面に
固定される。ナツト１５は炭化珪素製で、ボルト１２に
嵌合しばね１３を介して翼体を組立てる。断熱材１６は
炭化珪素繊維製で、翼１内及び上、下サイドウオール板
２，３と上、下サイドウオール４，５間の空間に配置さ
れる。耐熱シール１７．１８は上、下サイドウオール板
２，３及び上、下サイドウオール４，５の縁部に設けら
れた溝に挿入されて隣接翼間における燃焼ガスのシール
を行う。It touches the bottom surface. These upper and lower side walls 4.5
Protrusions 8 and 9 are formed on the upper and lower surfaces of the blade body, and these protrusions 8 and 9 engage with a retainer ring 10 and a support ring 11 of the gas turbine casing, respectively, to fix the blade body. bolt 12
is made of silicon carbide, the threaded part is precisely ground after sintering, and the head is fixed to the outer surface of the lower sidewall 5 below, and the lower sidewall plate 3. Wings 1. It passes through the center portions of the upper sidewall plate 2 and the upper sidewall 6 and the center portion of the leaf spring 13. The leaf spring 13 is made of silicon carbide, the lower surface of which is finished flat by precision grinding, and is fixed to the upper surface of the upper sidewall 4 via a flat seat 14 also made of silicon carbide. The nut 15 is made of silicon carbide, and is fitted onto the bolt 12 to assemble the wing body via the spring 13. The heat insulating material 16 is made of silicon carbide fiber and is disposed within the blade 1 and in the space between the upper and lower sidewall plates 2 and 3 and the upper and lower sidewalls 4 and 5. The heat-resistant seals 17 and 18 are inserted into grooves provided at the edges of the upper and lower sidewall plates 2 and 3 and the upper and lower sidewalls 4 and 5 to seal combustion gas between adjacent blades.
以上の構成で翼体を組立てるにあたっては、下サイドウ
オール５にボルト１２を貫通させると共にその上面にシ
ールリング７と断熱材１６を配置してその上に下サイド
ウオール板３と翼１を載置する。次に、翼１内に断熱材
１６を充填した後、その上に上サイドウオール板２．シ
ールリング６゜断熱材１６及び上サイドウオール４を載
置し、さらにその上に平板座１４とばね１３を設置する
。When assembling the wing body with the above configuration, the bolts 12 are passed through the lower sidewall 5, the seal ring 7 and the heat insulating material 16 are arranged on the upper surface, and the lower sidewall plate 3 and the wing 1 are placed on top of the bolts 12. do. Next, after filling the inside of the blade 1 with a heat insulating material 16, an upper sidewall plate 2. A seal ring 6°, a heat insulating material 16 and an upper side wall 4 are placed, and a flat plate seat 14 and a spring 13 are further installed thereon.
その後、ボルト１２の上端からナツト１５を挿入して締
付けることにより、翼体は組立てられる。Thereafter, the wing body is assembled by inserting and tightening the nut 15 from the upper end of the bolt 12.
このとき、翼１には半径方向に圧縮応力が加えられ、熱
衝撃時などにおける翼１の破壊を防止する作用がある。At this time, compressive stress is applied to the blade 1 in the radial direction, which has the effect of preventing destruction of the blade 1 during thermal shock or the like.
これら翼体をケーシングに設置する方法は、従来の金属
製翼体の場合と同様で、円周方向に一部切欠いた部分か
ら挿入してケーシングのリテーナリング１０とサポート
リング１１に突起８，９を係合させて周方向に送り、隣
接する翼体の両者間にシール１７．１８を設置する。サ
ポートリング１１と突起９の間には半径方向の熱変形を
吸収する間隙２１を存在させである。The method for installing these wing bodies in the casing is the same as in the case of conventional metal wing bodies, by inserting them from a partially cut out part in the circumferential direction and attaching the protrusions 8 and 9 to the retainer ring 10 and support ring 11 of the casing. are engaged and sent circumferentially to establish seals 17, 18 between adjacent airfoils. A gap 21 is provided between the support ring 11 and the protrusion 9 to absorb thermal deformation in the radial direction.
約１５００℃の燃焼ガスが流入した場合も、流路を形成
してこの燃焼ガスと接触する翼１．上。Even when combustion gas of about 1500°C flows in, the blades 1 form a flow path and come into contact with the combustion gas. Up.
下サイドウオール板２，３及び伝熱によって昇温するボ
ルト１２．ばね１３．ナツト１５はいずれも十分な耐熱
性を有する炭化珪素製であるから、これらは空冷しなく
ても十分な強度と耐酸化性を有する。また、翼１が昇降
温によって膨張収縮し、ボルトに対して半径方向に相対
変位を生じても、そればばね１３の変形によって吸収し
、翼体に無理な応力が発生することは防止される。なお
、リテーナリング１０と上サイドウオール６の間の空間
１９及び下サイドウォール５下部の空間２０には、従来
の金属製翼体の場合と同様、冷却空気が導入されている
から、金属製のこれら部材１０゜６．５は高温にさらさ
れることはない。Lower sidewall plates 2, 3 and bolts 12 whose temperature increases due to heat transfer. Spring 13. Since the nuts 15 are all made of silicon carbide which has sufficient heat resistance, they have sufficient strength and oxidation resistance even without air cooling. Furthermore, even if the blade 1 expands and contracts due to rising and falling temperatures and causes a relative displacement in the radial direction with respect to the bolt, this is absorbed by the deformation of the spring 13, and unreasonable stress is prevented from being generated in the blade body. . Note that cooling air is introduced into the space 19 between the retainer ring 10 and the upper sidewall 6 and the space 20 below the lower sidewall 5, as in the case of conventional metal wing bodies. These parts 10°6.5 are not exposed to high temperatures.
第３図に示す他の実施例では、上、下サイドウオール４
，５そのものを炭化珪素製とし、ボルト１２も下サイド
ウオール５と一体としたもので、第１図に示した実施例
における上、下サイドウオール板２，３．シールリング
６．７．平板座１４及び断熱材１６を排除することが出
来る。この構造では熱応力が発生し易いため、温度変化
のゆるやかな場合に適している。In another embodiment shown in FIG.
, 5 themselves are made of silicon carbide, and the bolts 12 are also integrated with the lower sidewall 5, so that the upper and lower sidewall plates 2, 3, . Seal ring 6.7. The flat seat 14 and the heat insulating material 16 can be eliminated. Since thermal stress is likely to occur in this structure, it is suitable for cases where temperature changes are gradual.
第４図に示すさらに他の実施例は、第１図に示した実施
例における翼１．上、下サイドウオール板２，３を一体
に成形し、上、下サイドウオール板２．３と上、下サイ
ドウオール４，５間に板状のばね１３，１３を配置した
もので、シールリング６．７．ボルト１２．ナツト１５
及び平板座１４を排除できる。翼体の組立ては、リテー
ナリング１０．サポートリング１１へ挿入時に同時に行
なわれ、ばね１３，１３により突起８，９はそれぞれリ
テーナリング１０．サポートリング１１の内面へ半径方
向に押付けられる。Still another embodiment shown in FIG. 4 is based on the blade 1 in the embodiment shown in FIG. The upper and lower sidewall plates 2, 3 are integrally molded, and plate-shaped springs 13, 13 are arranged between the upper and lower sidewall plates 2.3 and the upper and lower sidewalls 4, 5, and the seal ring 6.7. Bolt 12. Natsu 15
And the flat plate seat 14 can be eliminated. To assemble the wing body, use the retainer ring 10. When inserted into the support ring 11, the springs 13, 13 cause the protrusions 8, 9 to move, respectively, into the retainer ring 10. It is pressed radially onto the inner surface of the support ring 11.
このように、本発明によれば翼に半径方向の圧縮応力が
付与されるとともに翼が加熱膨張することによる半径方
向変位はセラミック製ばねにより吸収されるため、過大
な熱応力が発生することは防止でき、さらに従来用いら
れていた金属コアとその冷却手段をなくして、構造簡単
なガスタービン用セラミック静翼が得られる。As described above, according to the present invention, radial compressive stress is applied to the blade and the radial displacement caused by heating and expansion of the blade is absorbed by the ceramic spring, so that excessive thermal stress is prevented from occurring. Furthermore, by eliminating the conventionally used metal core and its cooling means, a ceramic stator vane for a gas turbine with a simple structure can be obtained.
第１図は本発明の一実施例の縦断面図、第２図はその横
断面図、第３図は本発明の他の実施例の縦断面図、第４
図は本発明の更に他の実施例の断面図である。
１・・・翼、２・・・上サイドウオール板、３・・・下
サイドウオール板、４・・・上サイドウオール、５・・
・下サイドウオール、６，７・・・シールリング、１０
・・・リテーナリング、１１・・・サポートリング、１
２・・・ボルト、１３・・・ばね。FIG. 1 is a longitudinal cross-sectional view of one embodiment of the present invention, FIG. 2 is a cross-sectional view thereof, FIG. 3 is a longitudinal cross-sectional view of another embodiment of the present invention, and FIG.
The figure is a sectional view of yet another embodiment of the present invention. 1... Wing, 2... Upper side wall plate, 3... Lower side wall plate, 4... Upper side wall, 5...
・Lower side wall, 6, 7... Seal ring, 10
... Retainer ring, 11 ... Support ring, 1
2...Bolt, 13...Spring.
Claims (1)
それぞれケーシング外周部のリテーナリングと内周部の
サポートリングに係合すると共に前記翼の半径方向両端
を保持する上、下サイドウオールとからなり、これら翼
、上、下サイドウオールの間の少なくとも１ケ所にセラ
ミツク製ばねを介在させて半径方向のこれら間の相対変
位を吸収するようにしたガスタービン用セラミツク静翼
。 ２、特許請求の範囲第１項のものにおいて、翼を中空と
なすと共に該中空部を貫通してセラミツク製ボルトを設
置し、該ボルトは一端を上、下サイドウオールの一方に
固定され他端は他方を貫通してその外部に設置したばね
を介してナツトにより固定され、上、下サイドウオール
と翼を可撓性を有して半径方向に締結してなるガスター
ビン用セラミツク静翼。[Claims] 1. Ceramic blades forming an axial flow path for combustion gas;
The upper and lower sidewalls each engage a retainer ring on the outer periphery of the casing and a support ring on the inner periphery of the casing and hold both ends of the blade in the radial direction. A ceramic stator vane for a gas turbine in which ceramic springs are interposed at two positions to absorb relative displacement between these parts in the radial direction. 2. In claim 1, the wing is hollow and a ceramic bolt is installed through the hollow part, one end of the bolt is fixed to one of the upper and lower side walls, and the other end is fixed to one of the upper and lower side walls. A ceramic stationary vane for a gas turbine, which is fixed by a nut through a spring installed on the outside of the second side wall, and the upper and lower sidewalls and the blade are flexibly fastened in the radial direction.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP5659287A JPS63223302A (en) | 1987-03-13 | 1987-03-13 | Ceramics stationary blade for gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP5659287A JPS63223302A (en) | 1987-03-13 | 1987-03-13 | Ceramics stationary blade for gas turbine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPS63223302A true JPS63223302A (en) | 1988-09-16 |
Family
ID=13031466
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP5659287A Pending JPS63223302A (en) | 1987-03-13 | 1987-03-13 | Ceramics stationary blade for gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS63223302A (en) |
Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH02218824A (en) * | 1988-12-14 | 1990-08-31 | General Electric Co <Ge> | Frame assembly of gas turbine engine |
| US6081379A (en) * | 1998-10-28 | 2000-06-27 | Coherent, Inc. | Multiple coupled Gires-Tournois interferometers for group-delay-dispersion control |
| US6154318A (en) * | 1998-08-18 | 2000-11-28 | Coherent, Inc. | Group-delay-dispersive multilayer-mirror structures and method for designing same |
| US6164903A (en) * | 1998-12-22 | 2000-12-26 | United Technologies Corporation | Turbine vane mounting arrangement |
| JP2001041003A (en) * | 1999-07-16 | 2001-02-13 | General Electric Co <Ge> | Prestressed gas turbine nozzle |
| GB2409004A (en) * | 2003-12-08 | 2005-06-15 | Ingersoll Rand Energy Systems | Radial flow turbine |
| JP2007182881A (en) * | 2006-01-03 | 2007-07-19 | General Electric Co <Ge> | Gas turbine stator and turbine blade assembly |
| US7326030B2 (en) * | 2005-02-02 | 2008-02-05 | Siemens Power Generation, Inc. | Support system for a composite airfoil in a turbine engine |
| US10072516B2 (en) | 2014-09-24 | 2018-09-11 | United Technologies Corporation | Clamped vane arc segment having load-transmitting features |
| EP3670837A1 (en) * | 2018-12-21 | 2020-06-24 | Rolls-Royce plc | Turbine section of a gas turbine engine with ceramic matrix composite vanes |
| EP4030036A1 (en) * | 2021-01-15 | 2022-07-20 | Raytheon Technologies Corporation | Vane arc segment support platform with curved radial channel |
-
1987
- 1987-03-13 JP JP5659287A patent/JPS63223302A/en active Pending
Cited By (16)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH02218824A (en) * | 1988-12-14 | 1990-08-31 | General Electric Co <Ge> | Frame assembly of gas turbine engine |
| US6154318A (en) * | 1998-08-18 | 2000-11-28 | Coherent, Inc. | Group-delay-dispersive multilayer-mirror structures and method for designing same |
| US6081379A (en) * | 1998-10-28 | 2000-06-27 | Coherent, Inc. | Multiple coupled Gires-Tournois interferometers for group-delay-dispersion control |
| US6164903A (en) * | 1998-12-22 | 2000-12-26 | United Technologies Corporation | Turbine vane mounting arrangement |
| EP1013885A3 (en) * | 1998-12-22 | 2001-08-01 | United Technologies Corporation | Turbine vane mounting arrangement |
| JP2001041003A (en) * | 1999-07-16 | 2001-02-13 | General Electric Co <Ge> | Prestressed gas turbine nozzle |
| GB2409004A (en) * | 2003-12-08 | 2005-06-15 | Ingersoll Rand Energy Systems | Radial flow turbine |
| GB2409004B (en) * | 2003-12-08 | 2007-07-04 | Ingersoll Rand Energy Systems | Nozzle bolting arrangement for a radial flow turbine |
| US7326030B2 (en) * | 2005-02-02 | 2008-02-05 | Siemens Power Generation, Inc. | Support system for a composite airfoil in a turbine engine |
| JP2007182881A (en) * | 2006-01-03 | 2007-07-19 | General Electric Co <Ge> | Gas turbine stator and turbine blade assembly |
| US10072516B2 (en) | 2014-09-24 | 2018-09-11 | United Technologies Corporation | Clamped vane arc segment having load-transmitting features |
| EP3000979B1 (en) * | 2014-09-24 | 2018-10-31 | United Technologies Corporation | Clamped vane arc segment having load-transmitting features |
| EP3670837A1 (en) * | 2018-12-21 | 2020-06-24 | Rolls-Royce plc | Turbine section of a gas turbine engine with ceramic matrix composite vanes |
| US11047247B2 (en) | 2018-12-21 | 2021-06-29 | Rolls-Royce Plc | Turbine section of a gas turbine engine with ceramic matrix composite vanes |
| EP4030036A1 (en) * | 2021-01-15 | 2022-07-20 | Raytheon Technologies Corporation | Vane arc segment support platform with curved radial channel |
| US11448096B2 (en) | 2021-01-15 | 2022-09-20 | Raytheon Technologies Corporation | Vane arc segment support platform with curved radial channel |
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