JPH04219205A - High temperature member made of ceramic - Google Patents
High temperature member made of ceramicInfo
- Publication number
- JPH04219205A JPH04219205A JP8758191A JP8758191A JPH04219205A JP H04219205 A JPH04219205 A JP H04219205A JP 8758191 A JP8758191 A JP 8758191A JP 8758191 A JP8758191 A JP 8758191A JP H04219205 A JPH04219205 A JP H04219205A
- Authority
- JP
- Japan
- Prior art keywords
- gas turbine
- temperature
- ceramic
- high temperature
- pores
- 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- Withdrawn
Links
- 239000000919 ceramic Substances 0.000 title claims abstract description 27
- 238000001816 cooling Methods 0.000 claims abstract description 26
- 239000011148 porous material Substances 0.000 claims abstract description 13
- 239000002826 coolant Substances 0.000 claims description 10
- 238000000034 method Methods 0.000 abstract description 9
- 238000000465 moulding Methods 0.000 abstract description 5
- 238000005245 sintering Methods 0.000 abstract description 5
- 239000000126 substance Substances 0.000 abstract description 4
- 239000007789 gas Substances 0.000 description 47
- 239000000567 combustion gas Substances 0.000 description 7
- 239000002657 fibrous material Substances 0.000 description 6
- 229910010293 ceramic material Inorganic materials 0.000 description 5
- 238000010304 firing Methods 0.000 description 5
- 239000002994 raw material Substances 0.000 description 5
- 229910052581 Si3N4 Inorganic materials 0.000 description 4
- 230000007797 corrosion Effects 0.000 description 4
- 238000005260 corrosion Methods 0.000 description 4
- 239000000835 fiber Substances 0.000 description 4
- 238000002347 injection Methods 0.000 description 4
- 239000007924 injection Substances 0.000 description 4
- 230000035939 shock Effects 0.000 description 4
- HBMJWWWQQXIZIP-UHFFFAOYSA-N silicon carbide Chemical compound [Si+]#[C-] HBMJWWWQQXIZIP-UHFFFAOYSA-N 0.000 description 4
- 229910010271 silicon carbide Inorganic materials 0.000 description 4
- HQVNEWCFYHHQES-UHFFFAOYSA-N silicon nitride Chemical compound N12[Si]34N5[Si]62N3[Si]51N64 HQVNEWCFYHHQES-UHFFFAOYSA-N 0.000 description 4
- 238000012360 testing method Methods 0.000 description 4
- 239000011230 binding agent Substances 0.000 description 3
- 230000000052 comparative effect Effects 0.000 description 3
- 230000003628 erosive effect Effects 0.000 description 3
- 239000002184 metal Substances 0.000 description 3
- 229910002077 partially stabilized zirconia Inorganic materials 0.000 description 3
- 238000012545 processing Methods 0.000 description 3
- 238000009826 distribution Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 238000011156 evaluation Methods 0.000 description 2
- 239000000446 fuel Substances 0.000 description 2
- 238000000227 grinding Methods 0.000 description 2
- 238000010438 heat treatment Methods 0.000 description 2
- 238000002156 mixing Methods 0.000 description 2
- IJGRMHOSHXDMSA-UHFFFAOYSA-N Atomic nitrogen Chemical compound N#N IJGRMHOSHXDMSA-UHFFFAOYSA-N 0.000 description 1
- 229910045601 alloy Inorganic materials 0.000 description 1
- 239000000956 alloy Substances 0.000 description 1
- PNEYBMLMFCGWSK-UHFFFAOYSA-N aluminium oxide Inorganic materials [O-2].[O-2].[O-2].[Al+3].[Al+3] PNEYBMLMFCGWSK-UHFFFAOYSA-N 0.000 description 1
- 238000001354 calcination Methods 0.000 description 1
- 238000005238 degreasing Methods 0.000 description 1
- 229910001873 dinitrogen Inorganic materials 0.000 description 1
- 210000003746 feather Anatomy 0.000 description 1
- 238000001746 injection moulding Methods 0.000 description 1
- 238000004898 kneading Methods 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 229910003465 moissanite Inorganic materials 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 229910002076 stabilized zirconia Inorganic materials 0.000 description 1
- XLYOFNOQVPJJNP-UHFFFAOYSA-N water Substances O XLYOFNOQVPJJNP-UHFFFAOYSA-N 0.000 description 1
- FIXNOXLJNSSSLJ-UHFFFAOYSA-N ytterbium(III) oxide Inorganic materials O=[Yb]O[Yb]=O FIXNOXLJNSSSLJ-UHFFFAOYSA-N 0.000 description 1
Landscapes
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
【０００１】0001
【産業上の利用分野】本発明は、セラミック製高温部材
に係り、更に詳しくはガスタービン部材等の高温に晒さ
れるセラミック製部材に関する。本セラミック製部材は
ガスタービン用動翼、静翼などに好適に使用することが
できる。BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to high-temperature ceramic members, and more particularly to ceramic members exposed to high temperatures such as gas turbine members. This ceramic member can be suitably used for gas turbine rotor blades, stationary blades, and the like.
【０００２】0002
【従来の技術】窒化珪素、炭化珪素、部分安定化ジルコ
ニアなどのセラミック材は高耐熱性、高耐摩耗性、高硬
度、高耐食性等の優れた特性を有しているため、機械部
品の一部として使用されている。これらのセラミックは
相次ぐ改良、および設計の適性化などによってその利用
分野は広がりつつある。[Prior Art] Ceramic materials such as silicon nitride, silicon carbide, and partially stabilized zirconia have excellent properties such as high heat resistance, high wear resistance, high hardness, and high corrosion resistance. It is used as a part. The fields of use of these ceramics are expanding due to successive improvements and optimization of designs.
【０００３】そして近年になり、次世代エンジンとして
ガスタービンエンジンが注目を集めている。ガスタービ
ンエンジンは高温の燃焼ガスを直接タービンロータに当
て動力を得る回転型エンジンであり、構造上燃焼器以外
の圧縮機、タービンロータおよび回転式熱交換器などの
各要素が回転機械であることに特長を有するものである
。このため、ガスタービンでは、排ガスの低公害性、使
用燃料の多様化、低振動・低騒音性、エンジンの軽量化
等のメリットが期待できるものである。[0003] In recent years, gas turbine engines have been attracting attention as next-generation engines. A gas turbine engine is a rotary engine that generates power by directing high-temperature combustion gas to the turbine rotor, and structurally, each element other than the combustor, such as the compressor, turbine rotor, and rotary heat exchanger, is a rotating machine. It has the following features. Therefore, gas turbines can be expected to have advantages such as low exhaust gas pollution, diversification of fuels used, low vibration and noise, and lightweight engines.
【０００４】しかし、上記のようなメリットがあるにも
拘らず、現在までに実用化に至っていない大きな理由と
しては燃費の面で従来エンジンを超えることができない
ためである。従って、ガスタービンエンジンの実用化に
当っては機関熱効率を向上させることが重要な課題であ
り、このためにはタービン入口のガス温度（ＴＩＴ）を
上昇させることが必須条件といえる。ガスタービンが即
、セラミックガスタービンといわれているのは、まさに
この点からであり、耐熱合金よりも耐熱性に優れたセラ
ミック材の実用化・開発が急務となっているのである。However, despite the advantages mentioned above, the main reason why it has not been put into practical use to date is that it cannot exceed conventional engines in terms of fuel efficiency. Therefore, when putting gas turbine engines into practical use, it is an important issue to improve the engine thermal efficiency, and for this purpose, it is an essential condition to increase the gas temperature (TIT) at the turbine inlet. It is precisely for this reason that gas turbines are called ceramic gas turbines, and there is an urgent need to commercialize and develop ceramic materials that have better heat resistance than heat-resistant alloys.
【０００５】[0005]
【発明が解決しようとする課題】ところで、セラミック
材を、例えばＴＩＴが１５００℃を超えるような高温用
ガスタービン部材として適用する場合、部分的には１６
００℃を超える温度領域が存在することとなり、セラミ
ック材といえども強度低下が発生し、またエロージョン
（浸食）あるいはコロージョン（腐食）の影響により、
ガスタービン部材としての信頼性の低下、および寿命の
短縮化等の問題が生じてくる。従って、本発明は、上記
したような高温に晒される領域・部材として使用しても
強度低下を生じず、信頼性の高いセラミック製の高温部
材を提供することを目的とするものである。[Problems to be Solved by the Invention] By the way, when ceramic materials are used as high-temperature gas turbine members where the TIT exceeds 1,500°C, some
There will be a temperature range exceeding 00℃, and even ceramic materials will experience a decrease in strength, and due to the effects of erosion or corrosion,
Problems such as a decrease in reliability as a gas turbine member and a shortened life span arise. Therefore, an object of the present invention is to provide a highly reliable ceramic high-temperature member that does not cause a decrease in strength even when used as a region or member exposed to high temperatures as described above.
【０００６】[0006]
【課題を解決するための手段】そしてその目的は、本発
明によれば、セラミック体の所定位置に、冷却媒体を流
通させるための冷却機構用細孔を備えたセラミック製高
温部材により達成することができる。本発明に係るセラ
ミック製高温部材は、１０００℃以上という高温条件下
で使用されるのに適しており、特にガスタービン用動翼
、静翼等のガスタービン部材として、好ましく用いるこ
とができる。また、冷却媒体を流通させるための冷却機
構用細孔はその径を０．２ｍｍφ以上とすることが、目
詰まりを防止する上で望ましい。[Means for Solving the Problems] According to the present invention, this object is achieved by a ceramic high-temperature member provided with cooling mechanism pores for circulating a cooling medium at predetermined positions of a ceramic body. I can do it. The ceramic high-temperature member according to the present invention is suitable for use under high-temperature conditions of 1000° C. or higher, and can be particularly preferably used as gas turbine members such as gas turbine rotor blades and stationary blades. Further, in order to prevent clogging, it is desirable that the diameter of the cooling mechanism pores for circulating the cooling medium be 0.2 mm or more.
【０００７】[0007]
【作用】本発明のセラミック製高温部材は、その所定位
置に冷却媒体を流通させるための冷却機構用細孔を備え
ているため、例えばガスタービン部材に適用した場合、
タービン入口ガス温度（ＴＩＴ）を１５００℃以上にす
ることができ、熱効率が大幅に向上する。また、セラミ
ック製高温部材で問題となるシャットダウン時の熱衝撃
を回避することができ、信頼性、寿命が大幅に向上する
。[Function] The ceramic high-temperature member of the present invention is equipped with cooling mechanism pores for circulating a cooling medium at predetermined positions, so when applied to a gas turbine member, for example,
The turbine inlet gas temperature (TIT) can be increased to 1500° C. or higher, greatly improving thermal efficiency. Additionally, thermal shock during shutdown, which is a problem with ceramic high-temperature components, can be avoided, significantly improving reliability and service life.
【０００８】本発明のセラミック製高温部材では、その
表面を冷却媒体によりクーリングできるため、ガスター
ビン部材の場合、燃焼ガスとの直接的な接触を回避でき
、エロージョン（浸食）あるいはコロージョン（腐食）
を効果的に防止する。また金属製ガスタービンの冷却と
比較して、冷却媒体量は少量でよく、冷却による熱効率
の低下は少ない。[0008] In the ceramic high-temperature member of the present invention, its surface can be cooled with a cooling medium, so in the case of a gas turbine member, direct contact with combustion gas can be avoided, and erosion or corrosion can be avoided.
effectively prevent. Also, compared to cooling a metal gas turbine, the amount of cooling medium required is small, and the decrease in thermal efficiency due to cooling is small.
【０００９】次に、本発明のセラミック製高温部材のう
ちガスタービン部材の製造方法の例を説明する。まず、
化学繊維等の繊維状物を準備する。繊維状物は、目的と
する径と焼成時の収縮を見込んだ割掛率から計算した太
さとする。次いで化学繊維の場合には熱処理を施して繊
維状物を真直とする。なお、金属繊維を用いる場合には
熱処理は必要ではない。Next, an example of a method for manufacturing a gas turbine member among the ceramic high-temperature members of the present invention will be explained. first,
Prepare fibrous materials such as chemical fibers. The thickness of the fibrous material is calculated from the desired diameter and the cut rate that takes into account shrinkage during firing. Next, in the case of chemical fibers, heat treatment is applied to straighten the fibrous material. Note that heat treatment is not necessary when using metal fibers.
【００１０】次に、セラミック原料に焼結助剤を添加混
合後、バインダーを添加し混練して得られる成形原料を
、前記繊維状物が所定位置に配置された金型内に射出し
てガスタービン部材の成形体を得る。次いで、この成形
体を脱脂後、セラミックの種類に応じた焼成温度、焼成
雰囲気下で焼成することにより繊維状物を燃焼消失させ
て、目的の冷却機構用細孔を備えたガスタービン部材を
得ることができる。Next, after adding and mixing a sintering aid to the ceramic raw material, a binder is added and kneaded, and the resulting molding raw material is injected into a mold in which the fibrous material is placed at a predetermined position, and gas A molded body of a turbine member is obtained. Next, this molded body is degreased and then fired at a firing temperature and under a firing atmosphere depending on the type of ceramic to burn off the fibrous materials and obtain a gas turbine member having the desired cooling mechanism pores. be able to.
【００１１】図１は上記方法により得られたガスタービ
ン用動翼（回転翼）３を示すもので、羽根部１と翼根部
２とを有し、かつガスタービン用動翼３の所定位置に冷
却機構用細孔４を形成したものである。なお、ガスター
ビン部材に細孔を形成する手段としては、繊維状物を燃
焼消失させる方法のほか、繊維の端部を把持して引き抜
く方法、成形時加工もしくはバインダー仮焼後の乾式加
工、あるいは焼結後の研削加工、超音波加工などの方法
を採用することができる。FIG. 1 shows a moving blade (rotor blade) 3 for a gas turbine obtained by the above method, which has a blade part 1 and a blade root part 2, and has a rotor blade 3 at a predetermined position of the moving blade 3 for a gas turbine. A cooling mechanism pore 4 is formed therein. In addition to the method of burning and extinguishing the fibrous material, methods for forming pores in gas turbine components include a method of grasping and pulling out the ends of the fibers, processing during molding or dry processing after calcination of the binder, or Methods such as grinding after sintering, ultrasonic processing, etc. can be employed.
【００１２】セラミック製高温部材の所定位置に配置さ
れる細孔４は冷却機構用であり、冷却媒体が流通する。
冷却媒体としては種類は問わないが、空気、水などを挙
げることができる。本発明の対象とするセラミック材と
しては、一般的に知られているすべてのセミックが適用
でき、例えばアルミナ、窒化珪素、炭化珪素、部分安定
化ジルコニア、安定化ジルコニアなどに適用できるが、
いわゆる難削材と呼ばれる窒化珪素、炭化珪素、部分安
定化ジルコニアを対象とすることが特に有効である。The pores 4 arranged at predetermined positions in the ceramic high-temperature member are for a cooling mechanism, through which a cooling medium flows. The cooling medium may be of any type, but may include air, water, and the like. As the ceramic material to which the present invention is applied, all generally known ceramics can be applied, such as alumina, silicon nitride, silicon carbide, partially stabilized zirconia, stabilized zirconia, etc.
It is particularly effective to target silicon nitride, silicon carbide, and partially stabilized zirconia, which are so-called difficult-to-cut materials.
【００１３】[0013]
【実施例】以下、本発明を実施例に基づき更に詳しく説
明するが、本発明はこれらの実施例に限られるものでは
ない。EXAMPLES The present invention will be explained in more detail below based on Examples, but the present invention is not limited to these Examples.
【００１４】（実施例１）窒化珪素原料に、Ｙ２ Ｏ３
、Ｙｂ２ Ｏ３ 、ＳｉＣを焼結助剤として添加し、
混合した後、さらにバインダーを添加し混練して得られ
た成形原料を、ピアノ線材製ワイヤーを所定位置に配置
した射出成形用金型内に射出して、図１に示すような羽
根部１と翼根部２とを有し、かつ多数の冷却孔４を形成
したガスタービン用動翼３を成形した。次に、この成形
体からワイヤーを引き抜いた後、約３００℃で脱脂し、
次いで窒素ガス雰囲気中１７００℃の温度で焼成するこ
とにより、冷却孔４を備えたガスタービン用動翼３を製
造した。(Example 1) Y2 O3 was added to the silicon nitride raw material.
, Yb2O3, and SiC are added as sintering aids,
After mixing, the molding raw material obtained by adding a binder and kneading is injected into an injection mold in which a piano wire wire is placed at a predetermined position to form a blade part 1 as shown in FIG. A gas turbine rotor blade 3 having a blade root portion 2 and a large number of cooling holes 4 was molded. Next, after pulling out the wire from this molded body, it is degreased at about 300°C,
Next, the rotor blade 3 for a gas turbine provided with the cooling holes 4 was manufactured by firing at a temperature of 1700° C. in a nitrogen gas atmosphere.
【００１５】（比較例）実施例１と、ピアノ線材製ワイ
ヤーを射出成形用金型内の所定位置に配置しないこと以
外は、すべて同一の条件で、図３に示すガスタービン用
動翼５を製造した。(Comparative Example) The gas turbine rotor blade 5 shown in FIG. 3 was manufactured under the same conditions as in Example 1, except that the piano wire wire was not placed at a predetermined position in the injection mold. Manufactured.
【００１６】（実施例２）実施例１と同一の成形原料を
用い、これを射出成形用金型内に射出して、図２に示す
形状のガスタービン用静翼１０を成形した。次いで、実
施例１と同一条件で脱脂、焼成を行ない、焼結体のガス
タービン用静翼１０を得た。ここで、ガスタービン用静
翼１０は冷却孔６，７，８，９を備えたものであるが、
冷却孔６は射出成形時、射出成形用金型内の所定位置に
芯金を配置することにより形成した。一方、冷却孔７〜
９は、それぞれ焼結後の研削加工により形成した。(Example 2) Using the same molding raw material as in Example 1, it was injected into an injection mold to mold a stator vane 10 for a gas turbine having the shape shown in FIG. 2. Next, degreasing and firing were performed under the same conditions as in Example 1 to obtain a sintered stationary blade 10 for a gas turbine. Here, the gas turbine stationary blade 10 is equipped with cooling holes 6, 7, 8, and 9.
The cooling holes 6 were formed by placing a core metal at a predetermined position within the injection mold during injection molding. On the other hand, cooling hole 7~
No. 9 was formed by grinding after sintering.
【００１７】［評価］以上の実施例１，２および比較例
で得られたガスタービン部材を用い、下記のようにして
評価試験を行なった。■実施例１で得られた図１に示す
ガスタービン用動翼３を、冷却孔４に冷却空気を流しな
がら、１５００℃の燃焼ガス流中に１００時間保持した
。その結果、ガスタービン用動翼３の部材表面平均温度
を１１００℃とすることができた。また、燃焼ガスを止
め、試験後のガスタービン用動翼３を観察したところ、
何等変化は見られなかった。[Evaluation] Using the gas turbine members obtained in Examples 1 and 2 and Comparative Example above, an evaluation test was conducted as follows. (2) The gas turbine rotor blade 3 shown in FIG. 1 obtained in Example 1 was held in a combustion gas flow at 1500° C. for 100 hours while cooling air was allowed to flow through the cooling holes 4. As a result, the average temperature of the member surface of the gas turbine rotor blade 3 could be set to 1100°C. In addition, when the combustion gas was stopped and the gas turbine rotor blade 3 was observed after the test,
No changes were observed.
【００１８】■また、実施例２で得られた図２に示すガ
スタービン用静翼１０を、冷却孔６，７，８，９に冷却
空気を流しながら、１５００℃の燃焼ガス流中に１００
時間保持した。その結果、ガスタービン用静翼１０の部
材表面平均温度を１０００℃とするとともに、部材の温
度分布についても温度差（Ｔｍａｘ−Ｔｍｉｎ ）を１
００℃以内にすることができ、シャットダウン時の熱衝
撃を低減することができた。また、燃焼ガスを止め、試
験後のガスタービン用静翼１０を観察したところ、何等
変化は見られなかった。■Also, the gas turbine stationary blade 10 shown in FIG.
Holds time. As a result, the average surface temperature of the member surface of the gas turbine stationary blade 10 was set to 1000°C, and the temperature difference (Tmax-Tmin) of the member temperature distribution was set to 1.
It was possible to reduce the thermal shock at the time of shutdown by lowering the temperature to within 00°C. Further, when the combustion gas was stopped and the gas turbine stationary blade 10 was observed after the test, no change was observed.
【００１９】■比較例で得られた図３のガスタービン用
動翼５を、１５００℃の燃焼ガス流中に１００時間保持
したところ、動翼５の部材表面平均温度が１４００℃と
なった。また、図２に示すガスタービン用静翼１０につ
いて、冷却空気を流さずに１５００℃の燃焼ガス流中に
１００時間保持したところ、静翼１０の部材表面平均温
度が１４００℃となった。この場合、部材の温度分布は
２５０℃となり、シャットダウン時の熱衝撃が大きかっ
た。さらに、試験後、両者を観察したところ、最高温度
（１６５０〜１７００℃）となる領域近傍に、顕著なエ
ロージョンおよびコロージョンが生じていた。(2) When the gas turbine rotor blade 5 shown in FIG. 3 obtained in the comparative example was held in a combustion gas flow at 1500°C for 100 hours, the average temperature of the member surface of the rotor blade 5 was 1400°C. Further, when the stator blade 10 for a gas turbine shown in FIG. 2 was held in a combustion gas flow at 1500°C for 100 hours without flowing cooling air, the average temperature of the member surface of the stator blade 10 was 1400°C. In this case, the temperature distribution of the member was 250° C., and the thermal shock during shutdown was large. Furthermore, when both were observed after the test, significant erosion and corrosion had occurred near the region where the maximum temperature (1650 to 1700°C) was reached.
【００２０】[0020]
【発明の効果】以上説明した通り、本発明のセラミック
製高温部材によれば、冷却媒体を流通させるための冷却
機構用細孔を備えているため、例えばガスタービン部材
に適用した場合、タービン入口ガス温度（ＴＩＴ）を１
５００℃以上にすることができ、熱効率を大幅に向上さ
せることができる。さらに、セラミック製高温部材で問
題となるシャットダウン時の熱衝撃を回避することがで
き、信頼性、寿命を大幅に向上させることができるとい
う利点を有する。Effects of the Invention As explained above, the ceramic high-temperature member of the present invention is provided with cooling mechanism pores for circulating a cooling medium, so when applied to a gas turbine member, for example, Gas temperature (TIT) is 1
The temperature can be increased to 500° C. or higher, and thermal efficiency can be significantly improved. Furthermore, it has the advantage that thermal shock during shutdown, which is a problem with ceramic high-temperature members, can be avoided, and reliability and lifespan can be significantly improved.
【図１】本発明の一実施例であるガスタービン用動翼を
示す斜視図である。FIG. 1 is a perspective view showing a rotor blade for a gas turbine, which is an embodiment of the present invention.
【図２】本発明の他の実施例であるガスタービン用静翼
を示す斜視図である。FIG. 2 is a perspective view showing a stator blade for a gas turbine according to another embodiment of the present invention.
【図３】細孔を有さないガスタービン用動翼を示す斜視
図である。FIG. 3 is a perspective view showing a gas turbine rotor blade without pores.
１ 羽根部 ２ 翼根部 ３ ガスタービン用動翼 ４ 冷却孔 ５ ガスタービン用動翼 ６ 冷却孔 ７ 冷却孔 ８ 冷却孔 ９ 冷却孔 １０ ガスタービン用静翼 1. Feather part 2 Wing root 3. Moving blades for gas turbines 4 Cooling holes 5. Moving blades for gas turbines 6 Cooling holes 7 Cooling holes 8 Cooling holes 9 Cooling holes 10 Stator blade for gas turbine
Claims (5)
を流通させるための冷却機構用細孔を備えたことを特徴
とするセラミック製高温部材。1. A ceramic high-temperature member characterized in that a ceramic body is provided with pores for a cooling mechanism at predetermined positions for circulating a cooling medium.
れる請求項１記載のセラミック製高温部材。2. The ceramic high-temperature member according to claim 1, which is used under a temperature condition of 1000° C. or higher.
求項１記載のセラミック製高温部材。3. The ceramic high-temperature member according to claim 1, which is used as a gas turbine member.
セラミック製高温部材。4. The ceramic high temperature member according to claim 1, wherein the cooling medium is air.
上である請求項１記載のセラミック製高温部材。5. The ceramic high-temperature member according to claim 1, wherein the cooling mechanism pores have a diameter of 0.2 mm or more.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP8758191A JPH04219205A (en) | 1990-03-31 | 1991-03-27 | High temperature member made of ceramic |
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| JP2-86000 | 1990-03-31 | ||
| JP02086000 | 1990-03-31 | ||
| JP8758191A JPH04219205A (en) | 1990-03-31 | 1991-03-27 | High temperature member made of ceramic |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| JPH04219205A true JPH04219205A (en) | 1992-08-10 |
Family
ID=26427024
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| JP8758191A Withdrawn JPH04219205A (en) | 1990-03-31 | 1991-03-27 | High temperature member made of ceramic |
Country Status (1)
| Country | Link |
|---|---|
| JP (1) | JPH04219205A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH0624858A (en) * | 1992-03-25 | 1994-02-01 | Ngk Insulators Ltd | Ceramic made structural member |
| US5342166A (en) * | 1992-04-06 | 1994-08-30 | Ngk Insulators, Ltd. | Ceramic gas-turbine nozzle with cooling fine holes and method for preparing the same |
| JP2001132405A (en) * | 1999-09-17 | 2001-05-15 | General Electric Co <Ge> | Device for reducing thermal stress in turbine aerofoil |
-
1991
- 1991-03-27 JP JP8758191A patent/JPH04219205A/en not_active Withdrawn
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH0624858A (en) * | 1992-03-25 | 1994-02-01 | Ngk Insulators Ltd | Ceramic made structural member |
| US5342166A (en) * | 1992-04-06 | 1994-08-30 | Ngk Insulators, Ltd. | Ceramic gas-turbine nozzle with cooling fine holes and method for preparing the same |
| US5425174A (en) * | 1992-04-06 | 1995-06-20 | Ngk Insulators, Ltd. | Method for preparing a ceramic gas-turbine nozzle with cooling fine holes |
| JP2001132405A (en) * | 1999-09-17 | 2001-05-15 | General Electric Co <Ge> | Device for reducing thermal stress in turbine aerofoil |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| KR100527685B1 (en) | Use of high temperature insulation for ceramic matrix composites in gas turbines | |
| JP4166977B2 (en) | High temperature corrosion resistant alloy material, thermal barrier coating material, turbine member, and gas turbine | |
| CA2461699C (en) | Hybrid ceramic material composed of insulating and structural ceramic layers | |
| US9598973B2 (en) | Seal systems for use in turbomachines and methods of fabricating the same | |
| US7655326B2 (en) | Thermal barrier coating material and method for production thereof, gas turbine member using the thermal barrier coating material, and gas turbine | |
| US7094021B2 (en) | Gas turbine flowpath structure | |
| JP4031631B2 (en) | Thermal barrier coating material, gas turbine member and gas turbine | |
| US20120141289A1 (en) | Airfoil with wrapped leading edge cooling passage | |
| US20080219854A1 (en) | Turbine component with axially spaced radially flowing microcircuit cooling channels | |
| US7918647B1 (en) | Turbine airfoil with flow blocking insert | |
| JP2013212972A (en) | Fiber reinforced barrier coating, method for applying barrier coating to component, and turbomachinery component | |
| JP2003160852A (en) | Thermal insulating coating material, manufacturing method therefor, turbine member and gas turbine | |
| WO1991015438A1 (en) | Ceramic high-temperature member | |
| JP2004003486A (en) | Method and device for prolonging effective service life of aerofoil of gas turbine engine | |
| JP4166978B2 (en) | High temperature corrosion resistant alloy material, thermal barrier coating material, turbine member, and gas turbine | |
| JP4303480B2 (en) | Turbine blades and casting equipment | |
| JPH04219205A (en) | High temperature member made of ceramic | |
| JP6067869B2 (en) | Turbine aero foil abradable coating system and corresponding turbine blades | |
| WO2019112662A1 (en) | Wall structure with three dimensional interface between metal and ceramic matrix composite portions | |
| US11261741B2 (en) | Ceramic airfoil trailing end configuration | |
| JP6224161B2 (en) | Rotor blade for gas turbine | |
| Nakakado et al. | Strength design and reliability evaluation of a hybrid ceramic stator vane for industrial gas turbines | |
| Brown | Hybrid laser process for shaped turbine-airfoil cooling holes | |
| RU122122U1 (en) | COMPONENT TURBINE SHOULDER BLADE | |
| JPH0624858A (en) | Ceramic made structural member |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| A300 | Application deemed to be withdrawn because no request for examination was validly filed |
Free format text: JAPANESE INTERMEDIATE CODE: A300 Effective date: 19980514 |