JPH0425404B2 - - Google Patents
Info
- Publication number
- JPH0425404B2 JPH0425404B2 JP18817284A JP18817284A JPH0425404B2 JP H0425404 B2 JPH0425404 B2 JP H0425404B2 JP 18817284 A JP18817284 A JP 18817284A JP 18817284 A JP18817284 A JP 18817284A JP H0425404 B2 JPH0425404 B2 JP H0425404B2
- Authority
- JP
- Japan
- Prior art keywords
- cooling
- ceramic
- same material
- turbine
- cooling air
- 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.)
- Expired
Links
- 238000001816 cooling Methods 0.000 claims description 30
- 239000000463 material Substances 0.000 claims description 19
- 239000000919 ceramic Substances 0.000 claims description 7
- 239000000945 filler Substances 0.000 claims description 6
- 239000000956 alloy Substances 0.000 claims description 5
- 229910045601 alloy Inorganic materials 0.000 claims description 4
- 229910010293 ceramic material Inorganic materials 0.000 description 11
- 239000007769 metal material Substances 0.000 description 4
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 2
- 230000007423 decrease Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 238000005516 engineering process Methods 0.000 description 2
- 238000004519 manufacturing process Methods 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 230000008646 thermal stress Effects 0.000 description 2
- 229910052581 Si3N4 Inorganic materials 0.000 description 1
- 230000003247 decreasing effect Effects 0.000 description 1
- 238000007599 discharging Methods 0.000 description 1
- 229910052759 nickel Inorganic materials 0.000 description 1
- 230000035939 shock Effects 0.000 description 1
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical compound [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 description 1
- 229910010271 silicon carbide Inorganic materials 0.000 description 1
- HQVNEWCFYHHQES-UHFFFAOYSA-N silicon nitride Chemical compound N12[Si]34N5[Si]62N3[Si]51N64 HQVNEWCFYHHQES-UHFFFAOYSA-N 0.000 description 1
- 238000005245 sintering Methods 0.000 description 1
- 229910000601 superalloy Inorganic materials 0.000 description 1
- 230000003313 weakening effect Effects 0.000 description 1
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
- F01D5/3084—Fixing blades to rotors; Blade roots ; Blade spacers the blades being made of ceramics
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators
Description
【発明の詳細な説明】
産業上の利用分野
この発明は、ガスタービンのタービン翼に関す
る。DETAILED DESCRIPTION OF THE INVENTION Field of Industrial Application This invention relates to turbine blades for gas turbines.
従来の技術
ガスタービンのタービン入口ガス温度は、熱効
率の向上および出力増大に応じて増大する傾向に
あり、そのためにタービン動静翼への熱負荷も増
大する一方であるが、いうまでもなく金属材料は
高温ではその強度の低下が著しく、そのために前
記ガス温度を上げても該金属材料の温度を低く保
持させる必要がある。その対策としては、従来か
ら材質向上および空気冷却の二方面からの処方が
行なわれているが、前者は原価の増大となり、後
者は性能の低下がともなうとされている。いま、
耐熱合金を使用し、空気による冷却を施した従来
タービン翼についてその静翼を挙例して説明する
と、翼材質には超耐熱合金を使用して精密鋳造に
よつて製作したもので、第３図に示すものは、内
側シユラウド２および外側シユラウド３を突設し
てある翼部１の内部に中空部４を配設してその冷
却空気孔５から空気を流入させ、翼後縁部に配置
してある冷却空気出口孔６から該空気を排出させ
て冷却を行なうもので、高度な対流冷却方式を採
用しており、また第４図に示すものは、翼部１の
内部に多数の中子冷却孔７を穿設させてある中子
８を嵌設させ、該翼部翼面に沿設してある翼面冷
却孔９および翼後縁部に配置した冷却空気出口孔
６から空気を排出させて冷却するもので、対流冷
却方式、翼面表層冷却方式および衝突効果をもつ
た強制冷却方式の三方式を併用させた高度な冷却
方式例を採用しており、ともにその冷却空気は空
気圧縮機によつて圧縮され、ガスタービンの燃焼
器およびタービンを通過する空気を抽気して使用
するために、その量の大小がガスタービンの熱効
率および出力に大きく影響を与える。BACKGROUND TECHNOLOGY The gas temperature at the turbine inlet of a gas turbine tends to increase as thermal efficiency improves and output increases, and as a result, the thermal load on the turbine rotor and stationary blades also increases. The strength of the metal material decreases significantly at high temperatures, so it is necessary to maintain the temperature of the metal material low even if the gas temperature is increased. As countermeasures, two measures have been taken to date: improving the quality of the material and cooling with air; however, the former is said to increase the cost, and the latter is said to be accompanied by a decrease in performance. now,
To explain conventional turbine blades that are made of heat-resistant alloys and cooled by air, we will explain them using a stator vane as an example. The one shown in the figure has a hollow part 4 disposed inside a wing part 1 on which an inner shroud 2 and an outer shroud 3 are protruded, and air flows in through the cooling air hole 5, and the hollow part 4 is arranged at the trailing edge of the wing. This system performs cooling by discharging the air from a cooling air outlet hole 6, which employs an advanced convection cooling system.The system shown in FIG. A core 8 with child cooling holes 7 drilled therein is fitted, and air is drawn through the blade cooling holes 9 along the blade surface of the wing and the cooling air outlet holes 6 arranged at the trailing edge of the blade. It uses an advanced cooling method that combines three methods: convection cooling, wing surface cooling, and forced cooling with collision effect. Since the air compressed by the compressor and passed through the combustor and turbine of the gas turbine is extracted and used, the amount of air extracted greatly affects the thermal efficiency and output of the gas turbine.
前述した従来対策を施してあるタービン翼にお
いては、その材質および冷却技術についての進歩
が極めて著しく、また原価低減および性能面から
もほぼ限界に達したとも考えられるが、近来脚光
を浴びているセラミツク材が実用されつつある現
状に鑑み、該セラミツク材をガスタービン高温
部、特にタービン翼部に従来の超耐熱合金に代え
て使用しようとする機運にある。 Regarding turbine blades, which have the conventional countermeasures mentioned above, advances in materials and cooling technology have been extremely remarkable, and it is thought that they have almost reached their limits in terms of cost reduction and performance. In view of the current situation where ceramic materials are being put into practical use, there is an opportunity to use ceramic materials in the high-temperature parts of gas turbines, particularly in turbine blades, in place of conventional super heat-resistant alloys.
しかるに、セラミツク材の熱膨脹係数が金属材
料の三ないし五分の一であるので、両材料の結合
に際して適切な対策を構じておかないと、熱膨脹
差のために高温域で大きな熱応力が生じ、特にセ
ラミツク材には延性がほとんどないから、該材で
製作された部品が破損し、さらにその結合部分が
緩徐するなどの欠点が避けられない。 However, since the coefficient of thermal expansion of ceramic materials is three to one-fifth that of metal materials, if appropriate measures are not taken when joining the two materials, the difference in thermal expansion will cause large thermal stress at high temperatures. In particular, since ceramic materials have almost no ductility, it is inevitable that parts made of this material will break and that the joints between them will become loose.
発明が解決しようとする問題点
この発明は、ガスタービンのタービン翼にセラ
ミツク材を適用してその金属材料との熱膨脹差に
よる破損を防止するとともに、その結合部分の緩
徐現象を除去することにある。Problems to be Solved by the Invention The present invention is to apply a ceramic material to the turbine blades of a gas turbine to prevent damage due to the difference in thermal expansion between the ceramic material and the metal material, and to eliminate the slowing phenomenon in the joint portion. .
問題点を解決するための手段
この発明は、セラミツク製内側シユラウドおよ
び同材製外側シユラウド間に同材製タービン翼部
を同材製充填材を介設して挿嵌させて熱膨脹係数
の小さい合金製のステイボルトで締結させるとと
もに、該ステイボルトに冷却空気を通気させてな
るものである。Means for Solving the Problems The present invention provides a method for inserting a turbine blade made of the same material between an inner shroud made of ceramic and an outer shroud made of the same material, with a filler made of the same material interposed therebetween, so that an alloy having a small coefficient of thermal expansion can be used. They are fastened together using stay bolts manufactured by the company, and cooling air is vented through the stay bolts.
作 用
したがつて、この発明の構成によれば、内、外
側シユラウドおよび翼部相互間に介設させてある
充填材が緩衝作用をする上に、締結ボルトとの熱
膨脹差がほとんどなくなるから、過大熱応力を発
生することなく、また該ボルト結合力が弱化して
緩徐することがなく、さらに冷却された締結ボル
トの強度が保持される。Therefore, according to the structure of the present invention, the filler interposed between the inner and outer shrouds and the wing parts not only acts as a buffer, but also almost eliminates the difference in thermal expansion with the fastening bolt. The strength of the cooled fastening bolt is maintained without generating excessive thermal stress, without weakening or slowing down the bolt connection force.
実施例
つぎに、この発明の実施例を図面によつて説明
すると、第１および第２図において、タービン翼
部として静翼を挙例して説述すれば、セラミツク
製翼部１０をセラミツク製内側シユラウド１１お
よび同材製外側シユラウド１２間に同材充填材１
３を介設させて挿嵌させ、前記内、外側シユラウ
ドおよび翼部を貫通して該両側シユラウドの夫々
の外側に内側シユラウド当て金１５および外側シ
ユラウド当て金１４を配設させてステイボルト１
６を前記内側シユラウド側から挿入させ、該ボル
トの前記外側シユラウド側端を回り止め座金１８
を介在させて締結ナツト１７によつて緊締させる
とともに、前記ステイボルトのボルト軸心に冷却
空気導入孔２０を穿設して該導入孔のステイボル
ト１６の頭部近傍に軸心に直交して穿孔してある
冷却空気導入透孔２１に接続させ、該導入透孔を
前記内側シユラウド当て金の内側シユラウド１１
当接側に配設してある冷却空気通路１９に連通さ
せるとともに、前記導入孔のステイボルト１７端
部分から冷却空気２２を送流自在にしているもの
で、前記翼部、内、外側シユラウドおよび充填材
に使用するセラミツク材としては、耐熱性、高温
強度、耐熱衝撃性および高靭性に優れた窒化けい
素、あるいは炭化けい素の常圧焼結品が推奨され
るが、前記材料以外でもその熱膨脹係数が前者材
の3.3×10-6／℃（ニツケル基超合金の五分の
一）、後者材の4.3×10-6／℃（同じく四分の一）
に相当するものであればよく、また前記ステイボ
ルト、内、外シユラウド当て金、締結ナツトの
夫々を熱膨脹係数の小さい合金材料で製作し、さ
らに前記冷却空気の空気源として空気圧縮機の吐
出空気を使用するものとし、したがつて常温から
の温度上昇は、例えば400℃程度であり、一方ガ
ス温度を1250℃とすれば、温度上昇は1230℃（空
温を20℃）となり、セラミツク材が無冷却でほぼ
ガス温度になるとすれば、その温度上昇は冷却さ
れたステイボルト１６の約3.0倍になるので、熱
膨脹量は両者ほとんど差がなくなる。Embodiment Next, an embodiment of the present invention will be described with reference to the drawings. In FIGS. 1 and 2, a stator blade will be described as an example of a turbine blade. A filling material 1 made of the same material is placed between the inner shroud 11 and the outer shroud 12 made of the same material.
The stay bolt 1 is inserted through the inner and outer shrouds and the wing portions, and an inner shroud stopper 15 and an outer shroud stopper 14 are disposed on the outside of each of the both shrouds.
6 is inserted from the inner shroud side, and the outer shroud side end of the bolt is fitted with a locking washer 18.
is interposed and tightened with a fastening nut 17, and a cooling air introduction hole 20 is bored in the bolt axis of the stay bolt, and the cooling air introduction hole 20 is perpendicular to the axis near the head of the stay bolt 16. The cooling air introduction hole 21 is connected to the perforated cooling air introduction hole 21, and the introduction hole is connected to the inner shroud 11 of the inner shroud pad.
It communicates with the cooling air passage 19 provided on the contact side, and allows cooling air 22 to flow freely from the end portion of the stay bolt 17 of the introduction hole, and connects the wing portion, inner and outer shrouds, and As the ceramic material used for the filler, pressureless sintered products of silicon nitride or silicon carbide, which have excellent heat resistance, high-temperature strength, thermal shock resistance, and high toughness, are recommended. The coefficient of thermal expansion is 3.3×10 -6 /℃ of the former material (one-fifth of the nickel-based superalloy), and 4.3×10 -6 /℃ (also one-fourth) of the latter material.
In addition, each of the stay bolts, inner and outer shroud pads, and fastening nuts may be made of an alloy material with a small coefficient of thermal expansion, and the discharge air of an air compressor may be used as the air source of the cooling air. Therefore, the temperature rise from room temperature is, for example, about 400°C. On the other hand, if the gas temperature is 1250°C, the temperature rise is 1230°C (air temperature 20°C), and the temperature rise from room temperature is about 400°C. If the stay bolt 16 is brought to almost the gas temperature without cooling, the temperature increase will be about 3.0 times that of the cooled stay bolt 16, so there will be almost no difference in the amount of thermal expansion between the two.
発明の効果
上述したように、この発明は、内、外側シユラ
ウドおよび翼部をセラミツク化するとともに、そ
の間に同材からなる充填材を介設してセラミツク
材相互間の緩衝材として作用させている上に、該
セラミツク材と熱膨脹差のほとんどないステイボ
ルトで締結させているので、該セラミツク製諸部
品の破損を防止できるとともに、結合部分が緩む
現象が起こることがなく、さらに締結しているボ
ルトを冷却させているから、その強度の低下のお
それが全くないなど、従来タービン翼にセラミツ
ク材を適用した場合の欠点を除去しているので、
そのセラミツク化の実用を実現でき、したがつ
て、ガスタービン入口ガス温度を上昇させられる
ので、タービン性能の格段の向上となり、さらに
セラミツク化の実用とともに量産にともなう焼結
用金型の製作償却費の低下となつて原価低減が大
幅に行なわれるなど、この発明の産業上の利用価
値は極めて広大である。Effects of the Invention As described above, in this invention, the inner and outer shrouds and the wing portions are made of ceramic, and a filler made of the same material is interposed between them to act as a buffer between the ceramic materials. Since the ceramic material is fastened with a stay bolt that has almost no difference in thermal expansion, it is possible to prevent damage to the ceramic parts and prevent the joint from loosening, and furthermore, the bolts that are fastened are Since the ceramic material is cooled, there is no risk of its strength decreasing, which eliminates the drawbacks of conventional ceramic materials used in turbine blades.
It is possible to put ceramic into practical use, and as a result, the gas temperature at the gas turbine inlet can be raised, resulting in a significant improvement in turbine performance.Furthermore, along with putting ceramic into practical use, manufacturing and depreciation costs for sintering molds associated with mass production can be reduced. The industrial utility value of this invention is extremely vast, such as the reduction in cost resulting in a significant reduction in cost.
第１図は、この発明の実施例を示す要部を切断
面であらわした側面図、第２図は、前図のＡ−Ａ
矢視図、第３図は、対流冷却方式を適用した従来
タービン翼の要部を切開してあらわした斜視図、
第４図は、対流冷却方式、翼面表層冷却方式およ
び強制冷却方式を併用した従来タービン翼の要部
を切開してあらわした斜視図である。
１０……セラミツク製翼部、１１……同材製内
側シユラウド、１２……同材製外側シユラウド、
１３……同材製充填材、１６……ステイボルト、
２０……冷却空気導入孔、２１……冷却空気導入
透孔、２２……冷却空気。
FIG. 1 is a cross-sectional side view showing the main parts of an embodiment of the present invention, and FIG.
Fig. 3 is a perspective view showing the main parts of a conventional turbine blade to which a convection cooling method is applied;
FIG. 4 is a cutaway perspective view showing the main parts of a conventional turbine blade that uses a convection cooling method, a blade surface cooling method, and a forced cooling method. 10...Ceramic wing part, 11...Inner shroud made of the same material, 12...Outer shroud made of the same material,
13... Filler made of the same material, 16... Stay bolt,
20...Cooling air introduction hole, 21...Cooling air introduction hole, 22...Cooling air.
Claims (1)
側シユラウド間に同材製タービン翼部を同材製充
填材を介設して挿嵌させて熱膨脹係数の小さい合
金製のステイボルトで締結させるとともに、該ス
テイボルトに冷却空気を通気させることを特徴と
するガスタービンのタービン翼。1 A turbine blade made of the same material is inserted between an inner shroud made of ceramic and an outer shroud made of the same material with a filler made of the same material interposed, and fastened with a stay bolt made of an alloy with a small coefficient of thermal expansion, and the stay is A gas turbine turbine blade characterized by ventilating cooling air through bolts.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP18817284A JPS6166802A (en) | 1984-09-10 | 1984-09-10 | Turbine blade of gas turbine |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP18817284A JPS6166802A (en) | 1984-09-10 | 1984-09-10 | Turbine blade of gas turbine |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| JPS6166802A JPS6166802A (en) | 1986-04-05 |
| JPH0425404B2 true JPH0425404B2 (en) | 1992-04-30 |
Family
ID=16219016
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP18817284A Granted JPS6166802A (en) | 1984-09-10 | 1984-09-10 | Turbine blade of gas turbine |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPS6166802A (en) |
Families Citing this family (37)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH0739805B2 (en) * | 1986-04-22 | 1995-05-01 | 株式会社東芝 | Turbine seal clearance adjustment device |
| JPS63191204U (en) * | 1987-05-29 | 1988-12-09 | ||
| US4987736A (en) * | 1988-12-14 | 1991-01-29 | General Electric Company | Lightweight gas turbine engine frame with free-floating heat shield |
| US6000906A (en) * | 1997-09-12 | 1999-12-14 | Alliedsignal Inc. | Ceramic airfoil |
| US6164903A (en) * | 1998-12-22 | 2000-12-26 | United Technologies Corporation | Turbine vane mounting arrangement |
| US9133724B2 (en) | 2012-01-09 | 2015-09-15 | General Electric Company | Turbomachine component including a cover plate |
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| US10428663B2 (en) | 2016-11-17 | 2019-10-01 | United Technologies Corporation | Airfoil with tie member and spring |
| US10808554B2 (en) | 2016-11-17 | 2020-10-20 | Raytheon Technologies Corporation | Method for making ceramic turbine engine article |
| US10415407B2 (en) | 2016-11-17 | 2019-09-17 | United Technologies Corporation | Airfoil pieces secured with endwall section |
| US10309238B2 (en) | 2016-11-17 | 2019-06-04 | United Technologies Corporation | Turbine engine component with geometrically segmented coating section and cooling passage |
| US10731495B2 (en) | 2016-11-17 | 2020-08-04 | Raytheon Technologies Corporation | Airfoil with panel having perimeter seal |
| US10746038B2 (en) | 2016-11-17 | 2020-08-18 | Raytheon Technologies Corporation | Airfoil with airfoil piece having radial seal |
| US10436062B2 (en) | 2016-11-17 | 2019-10-08 | United Technologies Corporation | Article having ceramic wall with flow turbulators |
| US10309226B2 (en) | 2016-11-17 | 2019-06-04 | United Technologies Corporation | Airfoil having panels |
| US10480334B2 (en) | 2016-11-17 | 2019-11-19 | United Technologies Corporation | Airfoil with geometrically segmented coating section |
| US10436049B2 (en) | 2016-11-17 | 2019-10-08 | United Technologies Corporation | Airfoil with dual profile leading end |
| US10767487B2 (en) | 2016-11-17 | 2020-09-08 | Raytheon Technologies Corporation | Airfoil with panel having flow guide |
| US10408082B2 (en) | 2016-11-17 | 2019-09-10 | United Technologies Corporation | Airfoil with retention pocket holding airfoil piece |
| US10598029B2 (en) | 2016-11-17 | 2020-03-24 | United Technologies Corporation | Airfoil with panel and side edge cooling |
| US10711616B2 (en) | 2016-11-17 | 2020-07-14 | Raytheon Technologies Corporation | Airfoil having endwall panels |
| US10830071B2 (en) | 2017-01-23 | 2020-11-10 | General Electric Company | System and method for the hybrid construction of multi-piece parts |
-
1984
- 1984-09-10 JP JP18817284A patent/JPS6166802A/en active Granted
Also Published As
| Publication number | Publication date |
|---|---|
| JPS6166802A (en) | 1986-04-05 |
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