JP2013076346A - Coated turbine cooling blade - Google Patents
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- JP2013076346A JP2013076346A JP2011215912A JP2011215912A JP2013076346A JP 2013076346 A JP2013076346 A JP 2013076346A JP 2011215912 A JP2011215912 A JP 2011215912A JP 2011215912 A JP2011215912 A JP 2011215912A JP 2013076346 A JP2013076346 A JP 2013076346A
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Description
本発明は、補修して再使用する翼内部冷却を施したタービン冷却翼に関する。 The present invention relates to a turbine cooling blade subjected to blade internal cooling that is repaired and reused.
燃焼ガス温度が素材許容限界温度を超えて使用される従来のタービン冷却翼には、例えば特許文献１および特許文献２に記載のものがある。 Examples of conventional turbine cooling blades in which the combustion gas temperature exceeds the material allowable limit temperature are described in Patent Document 1 and Patent Document 2, for example.
特許文献１に記載のタービン冷却翼は、流れを乱して熱伝達を促進させるための伝熱促進リブやピンフィンを後縁冷却流路に設けることにより、後縁部分の冷却性能を向上している。 The turbine cooling blade described in Patent Document 1 improves the cooling performance of the trailing edge portion by providing heat transfer promoting ribs and pin fins in the trailing edge cooling channel for disturbing the flow and promoting heat transfer. Yes.
また、特許文献２に記載のセラミック断熱層で表面を被覆されたタービン冷却翼は、その被膜層の厚さを部位によって異なるよう施工することにより、材料温度分布を均一化してタービン翼の寿命を長くしている。 In addition, the turbine cooling blade whose surface is coated with the ceramic heat insulating layer described in Patent Document 2 is constructed so that the thickness of the coating layer varies depending on the part, thereby uniformizing the material temperature distribution and extending the life of the turbine blade. It is long.
特許文献１および２に記載されているようなタービン冷却翼は、高温運転で使用した後に溶接補修して再利用される場合が一般である。 Turbine cooling blades such as those described in Patent Documents 1 and 2 are generally used after being repaired by welding after being used in high-temperature operation.
しかし、特許文献１に記載のタービン冷却翼を再利用するために溶接補修する場合、後縁冷却流路に設けられた伝熱促進リブやピンフィンは、形状が複雑なために、高温運転前と同じ形状に復元することは困難である。このため、特許文献１に記載のタービン冷却翼の再利用時には、伝熱促進リブやピンフィンが最適な形状とならず、後縁部分の冷却性能が高温運転前に比べて低下する可能性がある。 However, when welding repair is performed in order to reuse the turbine cooling blade described in Patent Document 1, the heat transfer promotion ribs and pin fins provided in the trailing edge cooling channel are complicated in shape, It is difficult to restore the same shape. For this reason, when the turbine cooling blade described in Patent Document 1 is reused, the heat transfer promoting ribs and the pin fins do not have an optimal shape, and the cooling performance of the trailing edge portion may be lower than before the high temperature operation. .
また、特許文献２に記載のセラミック断熱層で表面を被覆されたタービン冷却翼を再利用するために溶接補修する場合、補修前に翼表面のセラミック断熱層を除去しなければならい。したがって、特許文献２に記載のタービン冷却翼の再利用には、翼表面全体へのセラミック断熱層の除去・再施工作業が必要となるため、補修コストが増加する可能性がある。 In addition, when repairing a weld by the turbine cooling blade whose surface is covered with the ceramic heat insulating layer described in Patent Document 2, the ceramic heat insulating layer on the blade surface must be removed before the repair. Therefore, the reuse of the turbine cooling blade described in Patent Document 2 requires removal and re-construction work of the ceramic heat insulation layer on the entire blade surface, which may increase the repair cost.
更に、特許文献２に記載のタービン冷却翼は、セラミック断熱層の厚さが局所的に異なるため、一度で均一厚さのセラミック断熱層を被覆する溶射で施工する場合、施工回数増加に伴いセラミック断熱層施工コストが増加する可能性がある。結果として、補修コストとしては更に増加することとなり、特許文献２に記載のタービン冷却翼は、補修して再利用するには好適でない可能性がある。 Furthermore, since the thickness of the ceramic heat insulating layer is locally different in the turbine cooling blade described in Patent Document 2, when it is applied by thermal spraying to coat the ceramic heat insulating layer having a uniform thickness at a time, the ceramic cooling blade increases with the number of times of application. Thermal insulation layer construction costs may increase. As a result, the repair cost further increases, and the turbine cooling blade described in Patent Document 2 may not be suitable for repair and reuse.
本発明は、このような実状に鑑みなされたものであり、その目的は、再利用時も高温運転前と同等の冷却性能を保持でき、かつ再利用のための溶接補修作業が比較的容易であるタービン冷却翼を提供することにある。 The present invention has been made in view of such a situation. The purpose of the present invention is to maintain the same cooling performance as before high-temperature operation even during reuse, and relatively easy welding repair work for reuse. It is to provide a turbine cooling blade.
上記目的を達成するために、本発明のタービン冷却翼は、翼高さ方向に延びる前縁及び後縁と、前記前縁から前記後縁まで延びる背側部分および腹側部分の翼壁と、前記翼壁により形成される中空部分とを備え、前記中空部分に通風した冷却空気の少なくとも一部を前記後縁から噴出するための、前記中空部分から前記後縁へ貫通した単一の断面形状を有する貫通流路を、前記翼高さ方向に複数備え、前記貫通流路を通過する冷却空気により冷却される前記翼壁の外側表面にのみ、被覆層を有する。 In order to achieve the above object, a turbine cooling blade of the present invention includes a leading edge and a trailing edge extending in a blade height direction, and a blade wall of a back side portion and a ventral side portion extending from the leading edge to the trailing edge, A single cross-sectional shape penetrating from the hollow portion to the trailing edge for ejecting at least part of the cooling air ventilated through the hollow portion from the trailing edge. A plurality of through-flow passages are provided in the blade height direction, and the coating layer is provided only on the outer surface of the blade wall cooled by the cooling air passing through the through-flow passage.
本発明によれば、再利用時も高温運転前と同等の冷却性能を保持でき、かつ再利用のための補修作業が比較的容易であるタービン冷却翼を提供することができる。 According to the present invention, it is possible to provide a turbine cooling blade that can maintain a cooling performance equivalent to that before high-temperature operation even during reuse and that is relatively easy for repair work for reuse.
以下、本発明を実施するための形態について２つの実施例を挙げ、図１および図２を参照して詳細に説明する。なお、ここで挙げた２つの実施例は軸流タービンを例に挙げて説明するが、類似構造を有する半径流タービンなどにも本発明は適用可能である。 Hereinafter, two embodiments of the present invention will be described in detail with reference to FIG. 1 and FIG. The two embodiments described here are described by taking an axial turbine as an example, but the present invention can also be applied to a radial turbine having a similar structure.
図１には、本実施例に係るタービン冷却翼１の翼形部（羽根部）１０が横断面図で示されている。このタービン冷却翼１は、定置形や航空機ガスタービンで使用され、高温ガスを作動媒体として運転される。翼形部１０は、図１の奥行き方向である翼高さ方向に延びる前縁（入口縁）１４および後縁（出口縁）１３と、前縁１４から後縁１３まで延びた背側部分１１および腹側部分１２で形成されている。 FIG. 1 is a cross-sectional view of an airfoil portion (blade portion) 10 of a turbine cooling blade 1 according to the present embodiment. The turbine cooling blade 1 is used in a stationary type or an aircraft gas turbine, and is operated using a hot gas as a working medium. The airfoil portion 10 includes a leading edge (inlet edge) 14 and a trailing edge (exit edge) 13 that extend in the blade height direction that is the depth direction of FIG. 1, and a dorsal portion 11 that extends from the leading edge 14 to the trailing edge 13. And a ventral portion 12.
また、翼形部１０の内部には、支持リブで仕切られた第一および第二冷却流路２１、２２が設けられており、第二冷却流路２２は後縁冷却流路２３と繋がっている。作動媒体である高温ガスより低温の冷却空気を、第一および第二冷却流路２１、２２へ供給することにより、背側部分１１および腹側部分１２は冷却される。第二冷却流路２２へ供給された冷却空気は、後縁冷却流路２３を通過する際、背側部分１１および腹側部分１２の後縁側を主に冷却した後、後縁１３より作動媒体中へ噴出される。 In addition, first and second cooling channels 21 and 22 partitioned by support ribs are provided inside the airfoil portion 10, and the second cooling channel 22 is connected to the trailing edge cooling channel 23. Yes. By supplying cooling air having a temperature lower than that of the hot gas, which is a working medium, to the first and second cooling flow paths 21 and 22, the back portion 11 and the ventral portion 12 are cooled. When the cooling air supplied to the second cooling channel 22 passes through the trailing edge cooling channel 23, the cooling air mainly cools the trailing edge side of the back portion 11 and the ventral portion 12, and then the working medium from the trailing edge 13. It is spouted in.
本実施例の後縁冷却流路２３は、その入口から出口までを単一の円形断面によって構成された円孔流路を翼の高さ方向に複数個並べた単純な構造により形成されている。そのため、後縁１３から放電加工などの貫通孔を加工する方法で比較的容易に製作できる。一方で、この後縁冷却流路２３は、熱伝達を促進するための伝熱促進リブやピンフィンを設けた冷却流路と比較して、伝熱を促進する手段を設けていない分、冷却性能が低くなる傾向がある。このため、タービン冷却翼１を高温運転する際、後縁冷却流路２３の冷却効果が不十分となり、タービン冷却翼１の寿命が期待より短くなる場合が生じる可能性がある。 The trailing edge cooling flow path 23 of the present embodiment is formed by a simple structure in which a plurality of circular hole flow paths having a single circular cross section are arranged in the height direction of the blade from the inlet to the outlet. . Therefore, it can be manufactured relatively easily by a method of machining a through-hole such as electric discharge machining from the rear edge 13. On the other hand, the trailing edge cooling flow path 23 has a cooling performance equivalent to that provided with no means for promoting heat transfer, compared to a cooling flow path provided with heat transfer promoting ribs or pin fins for promoting heat transfer. Tend to be low. For this reason, when the turbine cooling blade 1 is operated at a high temperature, there is a possibility that the cooling effect of the trailing edge cooling channel 23 becomes insufficient and the life of the turbine cooling blade 1 may be shorter than expected.
このような場合に備え、本実施例に係るタービン冷却翼１では、後縁冷却流路２３の冷却効果が不十分な部分として、前縁から後縁に延びる翼壁のうち後縁側の約３０％の領域に限定して、背側部分１１および腹側部分１２のそれぞれの作動媒体側の外側表面１１０、１２０に断熱性に優れた被覆層３１を被覆させている。これにより、翼外表面１１０、１２０における金属温度を低下させ寿命を延長することができる。また、被覆層３１を被覆させる領域を翼壁の後縁側の約３０％の領域に限定することで、熱負荷の低減に加え、施工時および補修時におけるコストの低減が可能となる。この被覆層３１は、たとえばプラズマ溶射法などで施工できる均一厚さのセラミック断熱層とすることが好ましい。 In preparation for such a case, in the turbine cooling blade 1 according to the present embodiment, as a portion where the cooling effect of the trailing edge cooling channel 23 is insufficient, about 30 on the trailing edge side of the blade wall extending from the leading edge to the trailing edge. The outer layer 110 and 120 on the working medium side of the dorsal side portion 11 and the ventral side portion 12 is covered with a coating layer 31 having excellent heat insulation properties. Thereby, the metal temperature in blade outer surface 110,120 can be lowered | hung and lifetime can be extended. Moreover, by limiting the area | region which coat | covers the coating layer 31 to the area | region of about 30% of the trailing edge side of a blade wall, it becomes possible to reduce the cost at the time of construction and repair in addition to reduction of a thermal load. The coating layer 31 is preferably a ceramic heat insulating layer having a uniform thickness that can be applied by, for example, a plasma spraying method.
タービン冷却翼１を一定期間の高温運転後、再利用するために溶接補修をする場合において、被覆層３１を有する場合は、まず被覆層３１を除去して翼表面１１０、１２０の金属面を露出させる必要がある。次に、高温運転で損傷した部分が在る場合、損傷箇所を切削除去して翼材料と同種材料の肉盛り溶接により、翼形部１０の形状を高温運転前と同一形状に復元する。後縁冷却流路２３において、肉盛り溶接により円孔流路が塞がってしまった場合には、貫通孔を加工する方法により、高温運転前と同一の円孔形状に復元することが可能である。同一の円孔形状に復元することで、高温運転前と同等の冷却性能を維持しつつ、次回の補修作業を比較的容易なものとすることができる。被覆層３１を有する場合は、最後に、高温運転前と同一の範囲に、高温運転前と同一な厚さの被覆層３１を再施工する。 When the turbine cooling blade 1 is welded and repaired for reuse after a certain period of high temperature operation, when the coating layer 31 is provided, the coating layer 31 is first removed to expose the metal surfaces of the blade surfaces 110 and 120. It is necessary to let Next, when there is a portion damaged by the high temperature operation, the damaged portion is cut and removed, and the shape of the airfoil portion 10 is restored to the same shape as before the high temperature operation by overlay welding of the same material as the blade material. In the trailing edge cooling channel 23, when the circular hole channel is blocked by build-up welding, it is possible to restore the same circular hole shape as before the high temperature operation by a method of processing the through hole. . By restoring the same circular hole shape, it is possible to make the next repair work relatively easy while maintaining the same cooling performance as before the high temperature operation. When the coating layer 31 is provided, finally, the coating layer 31 having the same thickness as that before the high temperature operation is reconstructed in the same range as that before the high temperature operation.
以上に説明したように、本実施例のタービン冷却翼１は、従来に比べて比較的容易な補修作業により、再利用時も高温運転前と同等の後縁冷却性能を維持することができる。 As described above, the turbine cooling blade 1 of the present embodiment can maintain the trailing edge cooling performance equivalent to that before the high-temperature operation even during reuse by a relatively easy repair work compared to the conventional one.
なお、翼外表面１１０、１２０のうち、どちらか片側のみの金属温度を低下させれば良い場合には、被覆層３１の施工範囲は、必要最小限の範囲に留めた方が、被覆層３１の施工コストを低減できるので、翼外表面１１０、１２０の必要な側のみに施工範囲を限定した方が良い。 In the case where the metal temperature on only one side of the blade outer surfaces 110 and 120 only needs to be lowered, it is better to keep the construction range of the coating layer 31 within the minimum necessary range. Therefore, it is better to limit the construction range only to the necessary side of the blade outer surfaces 110 and 120.
図２には、本実施例に係るタービン冷却翼１の翼形部（羽根部）１０が横断面図で示されている。このタービン冷却翼１は、基本的には実施例１に記載のタービン冷却翼と同じであり、実施例１と異なる部分のみ以下に詳細に説明する。 FIG. 2 is a cross-sectional view of an airfoil portion (blade portion) 10 of the turbine cooling blade 1 according to this embodiment. The turbine cooling blade 1 is basically the same as the turbine cooling blade described in the first embodiment, and only portions different from the first embodiment will be described in detail below.
本実施例のタービン冷却翼１は、被覆層３１の被覆開始部分において、被覆する被覆層３１の厚さの分だけ翼外表面１１０、１２０をあらかじめ削り、くぼみ３２を設けている。これにより、被覆層３１を被覆した部分が被覆していない部分に対して段差を生じないように翼面を形成することができる。したがって、被覆層を翼表面の一部分にしか被覆していないにも関わらず、翼表面に外側へ出張った段差が無いことにより、その段差で生じる流れの剥離を抑えることができ、空力性能の損失を低減できる。また、本実施例の後縁冷却流路２３は、単一の四角形断面によって構成されており、より高い冷却性能を得ることができる。すなわち、本実施例に係るタービン冷却翼は翼外表面１１０、１２０における金属温度を実施例１と同等以下に保ったうえで、より高い空力性能を得ることができる。 In the turbine cooling blade 1 of the present embodiment, at the coating start portion of the coating layer 31, the blade outer surfaces 110 and 120 are cut in advance by the thickness of the coating layer 31 to be coated, and the recess 32 is provided. Thereby, a blade surface can be formed so that a level difference may not arise with respect to the part which the part which coat | covered the coating layer 31 does not coat | cover. Therefore, even though the coating layer covers only a part of the blade surface, there is no step that travels outward on the blade surface, so that flow separation caused by the step can be suppressed, resulting in loss of aerodynamic performance. Can be reduced. In addition, the trailing edge cooling channel 23 of the present embodiment is configured by a single rectangular cross section, and higher cooling performance can be obtained. That is, the turbine cooling blade according to the present embodiment can obtain higher aerodynamic performance while keeping the metal temperature at the blade outer surfaces 110 and 120 at the same level or lower as that of the first embodiment.
以上で説明したように、実施例１及び実施例２に係るタービン冷却翼では、その後縁冷却流路が単一の断面形状を有する貫通流路を翼高さ方向に複数個並べた比較的単純な構造により形成されているので、高温運転後に再利用するための溶接補修する場合に、高温運転前と同じ形状に復元できる。したがって、本発明によれば、タービン冷却翼の再利用時も高温運転前と同等の後縁冷却性能を保持することができる。 As described above, in the turbine cooling blades according to the first and second embodiments, the trailing edge cooling flow path is relatively simple in which a plurality of through flow paths having a single cross-sectional shape are arranged in the blade height direction. Since it is formed by a simple structure, it can be restored to the same shape as before the high temperature operation when repairing welding for reuse after the high temperature operation. Therefore, according to the present invention, the trailing edge cooling performance equivalent to that before high-temperature operation can be maintained even when the turbine cooling blade is reused.
また、均一厚さのセラミック断熱層による翼表面を被覆する部分を、主にその後縁冷却流路により冷却される部分のみに限定することにより、再利用のための溶接補修する場合に必要となるセラミック断熱層の除去・再施工作業範囲を最小限の範囲に止めることができる。したがって、本発明によれば、タービン冷却翼の再利用のための溶接補修作業を比較的容易に行うことができる。 In addition, it is necessary when repairing welding for reuse by limiting the portion of the blade surface covered with the ceramic insulation layer of uniform thickness to only the portion cooled mainly by the trailing edge cooling channel. The removal / reconstruction work range of the ceramic thermal insulation layer can be kept to a minimum range. Therefore, according to the present invention, the welding repair work for reusing the turbine cooling blade can be performed relatively easily.
さらに、特に高い空力性能が要求される場合を考慮し、実施例２に係るタービン冷却翼では、セラミック断熱層による被覆開始部分において、被覆するセラミック断熱層厚さの分だけ翼表面をあらかじめ削っておき、セラミック断熱層被覆部分が非被覆部分に対して段差が生じないように翼面を形成することができる。したがって、特に高い空力性能が要求される場合にも、補修性を維持したまま要求を満足するタービン冷却翼を提供することができる。 Furthermore, considering the case where particularly high aerodynamic performance is required, in the turbine cooling blade according to the second embodiment, the blade surface is cut in advance by the thickness of the ceramic heat insulating layer to be coated at the coating start portion by the ceramic heat insulating layer. In addition, the blade surface can be formed so that the ceramic heat insulating layer covering portion does not have a step with respect to the non-covering portion. Therefore, even when particularly high aerodynamic performance is required, it is possible to provide a turbine cooling blade that satisfies the requirements while maintaining repairability.
１ タービン冷却翼
１０ 翼形部（羽根部）
１１ 背側部分
１２ 腹側部分
１３ 後縁（出口縁）
１４ 前縁（入口縁）
２１ 第一冷却流路
２２ 第二冷却流路
２３ 後縁冷却流路
３１ 被覆層
３２ くぼみ
１１０ 背側部分の外側表面
１２０ 腹側部分の外側表面
1 Turbine cooling blade 10 Airfoil part (blade part)
11 dorsal side 12 ventral side 13 trailing edge (exit edge)
14 Leading edge (entrance edge)
21 1st cooling flow path 22 2nd cooling flow path 23 Trailing edge cooling flow path 31 Covering layer 32 Indentation 110 Outer side surface 120 Outer side surface
Claims (4)
前記中空部分に通風した冷却空気の少なくとも一部を前記後縁から噴出するための、前記中空部分から前記後縁へ貫通した単一の断面形状を有する貫通流路を、前記翼高さ方向に複数備え、
前記貫通流路を通過する冷却空気により冷却される前記翼壁の外側表面にのみ、被覆層を有することを特徴とした被覆タービン冷却翼。 An airfoil comprising a leading edge and a trailing edge extending in a blade height direction, a blade wall of a back side portion and a ventral portion extending from the leading edge to the trailing edge, and a hollow portion formed by the blade wall. A turbine cooling blade having
A through passage having a single cross-sectional shape penetrating from the hollow portion to the trailing edge for ejecting at least part of the cooling air ventilated through the hollow portion from the trailing edge in the blade height direction. Multiple
A coated turbine cooling blade having a coating layer only on an outer surface of the blade wall cooled by cooling air passing through the through passage.
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| JP2011215912A JP2013076346A (en) | 2011-09-30 | 2011-09-30 | Coated turbine cooling blade |
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| JP2011215912A JP2013076346A (en) | 2011-09-30 | 2011-09-30 | Coated turbine cooling blade |
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| JP2013076346A true JP2013076346A (en) | 2013-04-25 |
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| JP2011215912A Pending JP2013076346A (en) | 2011-09-30 | 2011-09-30 | Coated turbine cooling blade |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112389634A (en) * | 2020-11-27 | 2021-02-23 | 北京宇航系统工程研究所 | Heat-proof front edge grid rudder under medium-high heat flow condition |
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2011
- 2011-09-30 JP JP2011215912A patent/JP2013076346A/en active Pending
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112389634A (en) * | 2020-11-27 | 2021-02-23 | 北京宇航系统工程研究所 | Heat-proof front edge grid rudder under medium-high heat flow condition |
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