TW200925388A - Cooling of a stator vane of a gas turbine engine - Google Patents

Cooling of a stator vane of a gas turbine engine Download PDF

Info

Publication number
TW200925388A
TW200925388A TW97123226A TW97123226A TW200925388A TW 200925388 A TW200925388 A TW 200925388A TW 97123226 A TW97123226 A TW 97123226A TW 97123226 A TW97123226 A TW 97123226A TW 200925388 A TW200925388 A TW 200925388A
Authority
TW
Taiwan
Prior art keywords
blade
rib
trailing edge
platform
cooling
Prior art date
Application number
TW97123226A
Other languages
Chinese (zh)
Inventor
Dmitry Valerievich Petrunin
Anton Mikhailovich Sumin
Edouard Stanislavovich Sloutski
Alexander Anatolievich Khanin
Original Assignee
Alstom Technology Ltd
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 Alstom Technology Ltd filed Critical Alstom Technology Ltd
Publication of TW200925388A publication Critical patent/TW200925388A/en

Links

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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • 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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

The invention relates to a stator vane (1) for a gas turbine engine comprising a vane portion having a first and a second side wall (4, 5) and a cooling passage (6) arranged therebetween. The cooling passage (6) leads to a trailing edge slot (10) in the trailing edge region (9) of the vane portion. At least one rib (19) is provided in the cooling passage (6) in the area of a platform (2, 3) of the stator vane (1). The at least one rib (19) extends generally in the radial direction of the stator vane (1).

Description

200925388 九、發明說明： 【發明所屬之技術領域】 本發明整體而言係關於燃氣渴輪引擎，更具體而言係關 於其中可冷卻之定葉片。 【先前技術】 在一燃氣渦輪引擎中，環狀靜定葉片排列於該旋轉動葉 片之間之該渦輪轉子之區域中。由於該定葉片將遭遇從燃 燒室流出之該熱氣體，因此有必要提供一冷卻措施，用於 冷卻該定葉片。定葉片大致藉由來自一壓縮機之壓縮空氣 冷卻，該壓縮空氣流入該定葉片中之一中空空間，以便該 葉片金屬之溫度可被維持在該材料所能經受住之溫度之 下，比如以防止產生其所不能經受之熱應力。該冷卻空氣 流過該葉片平臺中之一人口進入該葉片之前緣區域之一中 空空間。該中空空間可為一蜿蜒延伸之通道，其通向一後 緣槽’冷卻空氣可從該後緣槽中流出。 對具有傳統之冷卻空氣流通通道之定葉片^，冷卻空 氣之該壓力分佈在-定葉片之後緣區域中為不均勻的。該 後緣區域中之該冷卻空氣之該不均旬壓力分佈導致該冷卻 空氣沿該後緣穿過該後緣槽之流速不同。其可導致該定葉 片在該後緣區域中局部過冷，該情況反而會造成該=緣中 之熱應力。 因此’有必要提供一種定葉片，其可用於一燃氣渦輪引 擎’且該葉片具有-改良之冷卻結構，該結構使該定葉片 在該後緣區域中之局部過冷最小化。 132161.doc 200925388 【發明内容】 本發明將闡述該等問題。根據本發明之一第一態樣係提 供一燃氣渦輪引擎之一定葉片’該定葉片包括一葉片部 分’該葉片部分具有一第一與第二側壁，以及一冷卻通 道，該冷卻通道被設計於該第一與一第二側壁之間。該冷 卻通道通向該葉片部分之該後緣區域中之一後緣槽，且至 少一個肋條被設置於該定葉片之一平臺之區域中之該冷卻200925388 IX. INSTRUCTIONS OF THE INVENTION: TECHNICAL FIELD OF THE INVENTION The present invention relates generally to gas thirsty wheel engines, and more particularly to stator blades that are coolable therein. [Prior Art] In a gas turbine engine, an annular stationary vane is arranged in the region of the turbine rotor between the rotating vanes. Since the stator blade will encounter the hot gas flowing out of the combustion chamber, it is necessary to provide a cooling measure for cooling the stator blade. The stator blade is substantially cooled by compressed air from a compressor, the compressed air flowing into one of the hollow spaces of the stator blade such that the temperature of the blade metal can be maintained below the temperature at which the material can withstand, such as Prevents thermal stresses that it cannot withstand. The cooling air flows through one of the vane platforms into an empty space in one of the leading edge regions of the vane. The hollow space may be a meandering passage leading to a trailing edge groove from which cooling air may flow out. For a fixed vane having a conventional cooling air flow passage, the pressure distribution of the cooling air is uneven in the trailing edge region of the stationary vane. The uneven pressure distribution of the cooling air in the trailing edge region results in a different flow rate of the cooling air along the trailing edge slot along the trailing edge. It can cause the stator blade to be partially subcooled in the trailing edge region, which in turn causes thermal stress in the = edge. It is therefore necessary to provide a fixed vane that can be used in a gas turbine engine' and that has a modified cooling structure that minimizes local subcooling of the stator blade in the trailing edge region. 132161.doc 200925388 SUMMARY OF THE INVENTION The present invention will address these issues. According to a first aspect of the present invention, a certain blade of a gas turbine engine is provided. The stator blade includes a blade portion having a first and second side wall and a cooling passage, the cooling passage being designed Between the first and a second side wall. The cooling passage leads to one of the trailing edge slots of the trailing edge region of the blade portion, and the cooling is provided at least one rib in the region of one of the stator blades

通道中，該至少一肋條大致在該定葉片之徑向方向上延 伸。該肋條具有該作用，可減小該定葉片平臺之區域中之 冷卻通道中流動之該冷卻空氣之該流速。藉由減小該平臺 之區域中之冷卻空氣之流速，使該局部過冷與其次級影響 之問題最小化。 本發明之一較佳之實施例中，該冷卻通道大致具有一蜿 蜒之形式，且藉由該側壁與分隔壁界定。該分離壁大致在 該側壁之間，在該定葉片之該徑向方向上延伸。藉由該側 壁與一後分隔壁，界定該後緣區域中之該冷卻通道，該後 分隔壁從該徑向内部平臺延伸。由此該至少-肋條被佈置 在該後分隔壁與該後緣槽之間，順著該燃㈣輪引擎之該 熱燃燒氣體之該流^該肋條之位置具有減小該葉片平臺 外部與該後緣中之該冷卻空氣之該流速之作用。 在本發明之一進一步較佳之實施例中，該肋條之長度與 該後緣槽之寬度之比為5_45:1至8 15:1，較佳為6 Η。八 配合該等關，其中某些較佳之實施例之該以下描述將 使本發明之上述與其他目的、特徵與優點將變得清晰。 132161.doc 200925388 【實施方式】 圖2顯示了先前技術中之一定葉片1，其主要包括一徑向 外葉片平臺3’ 一徑向内葉片平臺2’第一與第二側壁4, 5 (如圖3 )，以及一冷卻通道，該冷卻通道被設計於該側壁 4，5之間。 該定葉片1藉由進入該冷卻通道6中之該冷卻空氣冷卻。 通過該冷卻媒介，該葉片金屬之溫度被維持在該材料所能 經受住之溫度之下’比如以防止產生其所不能經受之熱應 力。該冷卻空氣通過該外平臺3中之一冷卻空氣入口 7進入 該冷卻通道6。該冷卻通道以一蜿蜒之形式在該葉片部分 之該前緣區域8與該後緣區域9之間延伸，如圖2所示。 該冷卻通道6通向該葉片部分之該後緣區域中之一後緣 槽10 ’該冷卻空氣從該後緣槽中流出。 在該等側壁4，5之間，設置三個徑向充分延伸之分隔壁 12’ 13’ 14，其界定該蜿蜒形狀之冷卻通道。該前分隔壁 12以其該徑向内部平臺2之區域中之自由端，徑向從該外 葉片平臺3向内部延伸。該中間分隔壁13具有2個自由端， 從該外葉片平臺3之區域延伸至該徑向内平臺2之區域中之 一點。該後分隔壁14從該内平臺2延伸至該徑向外葉片平 臺3之區域中之一點。此外，導流壁15，μ，17，18被設 置於該冷卻通道6中，處於該等分隔壁12，13，14之該等 自由端之區域中’其用於引導該冷卻通道6中之該冷卻空 氣流。 如圖2所示，該等導流壁17與18沿該後緣區域9之徑向方 132161.doc 200925388 向’轉移該定葉片1之該後緣區域9中之一部分該冷卻空氣 流。 在該葉片部分之該後緣區域9中’複數個柱11被排列在 該等侧壁之間。 與該後分隔壁14與該後緣槽10之間之該徑向内葉片平臺 2之區域中之冷卻空氣之流速相比，在該第一分隔壁12與 該後緣槽10之間之該外葉片平臺3之區域中之該冷卻空氣 之流速相對較高》此將導致該前分隔壁12與該後緣槽1〇之 間之該外葉片平臺3中之該定葉片1之局部過冷，該情況接 著會造成該定葉片中之不良熱應力。 圖1顯示根據本發明之一較佳實施之一定葉片1。類似元 件將以類似參考標號標註。 該定葉片1藉由經引入該冷卻通道6中之該冷卻媒介而冷 卻。透過該冷卻媒介’該葉片金屬之溫度被維持在該材料 所能經受住之溫度之下，比如以防止產生其所不能經受之 熱應力。較佳地，該渦輪之一壓縮機產生之壓縮空氣被作 為該冷卻媒介而使用。該冷卻媒介通過該外平臺3中之一 冷卻空氣入口 7流入該冷卻通道6 ^該冷卻通道以一蜿蜒之 形式在該葉片部分之該前緣區域8與該後緣區域9之間延 伸，如圖1所示。該冷卻通道6通向該葉片部分之該後緣區 域中之一後緣槽1 〇，該冷卻空氣從該後緣槽中流出。在該 疋葉片1之該等側壁4，5之間，設置三個徑向充分延伸之 刀隔壁12 ’ 13 ’ 14 ’其界定該婉蜒形狀之冷卻通道6。 根據本發明，至少一肋條19被設置於該定葉片丨之一葉 132161.doc -9^ 200925388 片平臺2，3之區域中之該冷卻通道6中。該術語”肋條&quot;意 味著一側壁4或5中之突出，而非一分隔壁，其從一侧壁* 延伸至另-側壁5。至少有-肋條19之該目的為減小流經 該肋條19與該反向側壁4，5之間之該冷卻媒介流之速度。 該肋條19大致在該定葉片丨之該徑向方向上延伸。較佳 地，該肋條19在該徑向方向上充分延伸，如圖2所示無 論如何其可在另一實施例中以一接近該徑向方向之角度延 伸。一彎曲形之肋條19也為可能。較佳地，該定葉片由一 鑄造製程完成，因此該至少一肋條19作為—突出形成於該 葉片之側壁4，5中之一個上。該肋條19之高度，比如突出 垂直於該側壁4，5之該肋條之高度，與該肋條輪廓之該形 狀由測試決定。該肋條19之輪廓被設計使得該肋條19與該 反向側壁之間之該冷卻媒介流速減小。 ❹ 較佳地，按流經該燃氣渦輪引擎之該渦輪之該熱燃燒氣 體之方向20 ’該肋條19被排列在該後分隔壁14與該後緣槽 10之間。根據該蜿蜒而成之冷卻通道6之幾何學，該肋條 19可從該徑向外葉片平臺3向内徑向延伸，及/或從該徑向 内葉片平臺2向外徑向延伸。如果該後分隔壁14從該内葉 片平臺2延伸’則該肋條19從該外葉片平臺3延伸。如果該 後分隔壁14從該外葉片平臺3延伸，則該肋條19從該内葉 片平臺2延伸。如果該後分隔壁丨4具有2個自由端，則可提 供2個肋條19分別從該内葉片平臺2或外葉片平臺3延伸。 如圖1所示’該後分隔壁14之該自由端之區域令，至少 提供一個導流壁17，18用於導引該後緣區域9中之該冷卻 132161.doc • 10- 200925388 媒介。較佳地’提供兩個導流壁17，18。該等導流壁I?， 18之目的為順著該定葉片1之徑向方向’轉移部分流入該 後緣區域之冷卻媒介。如圖1所示之實施例中，在該後分 隔壁14與該後緣槽1〇之間之該後分隔壁14之該徑向外區域 中，該等導流壁17，1 8被排列用於轉移一部分該後緣區域 9徑向向内之冷卻空氣流。 較佳地，該至少一肋條19以該弦向方向被排列在後分隔 壁14與複數個柱11之間，該等柱在該後緣區域9中之該等 侧壁4，5之間。該肋條19在該定葉片之該徑向方向上之長 度與該該後緣槽10之寬度21(如圖3所示）之比在5.45:1至 8.15:1之間，較佳為6.8:1。較佳地，該肋條19從該徑向外 平臺3向内徑向延伸至該至少一導流葉片14之該徑向内端 22之徑向向外之一點，如圖！所示。 如上所述，該後緣區域9中之該肋條丨9之位置與該肋條 19之該相對尺寸有利地減小了該葉片平臺2，3之該區域中 之該冷卻媒介之流速。透過減小該冷卻媒介之流速，使得 該定葉片1在該區域中之局部冷卻最小化。 本發明之較多修正與其他實施例將進入熟悉此項本發明 所適宜之技術的技術者之思維，本發明對上述描述與附圖 所闡述之教義有益。因此，應瞭解，本發明並非被限制於 該所揭不之具體之實施例，且該等修正與其他實施例將被 包含於該附屬請求項之範圍内。雖然具體項目被採用，但 疋其僅被用於大致與描述之意義，而非為限制之目的。 【圖式簡單說明】 132161.doc 200925388 藉由參考附圖所示意之實施例描述本發明，且參考該等 附圖，以下更多細節將被描述。 該等附圖示意性顯示於： 圖1係根據本發明之一實施例之一定葉片之一橫截面。 圖2係先前技術中之一定葉片之一橫截面。 圖3係沿A-A線截取圖1中之該定葉片之一橫截面。 【主要元件符號說明】In the passage, the at least one rib extends substantially in a radial direction of the stator blade. The rib has the function of reducing the flow rate of the cooling air flowing in the cooling passage in the region of the fixed blade platform. The problem of localized subcooling and its secondary effects is minimized by reducing the flow rate of cooling air in the region of the platform. In a preferred embodiment of the invention, the cooling passage has a substantially 蜿-shaped form and is defined by the side wall and the dividing wall. The separating wall extends substantially between the side walls in the radial direction of the stator blade. The cooling passage in the trailing edge region is defined by the side wall and a rear dividing wall, the rear dividing wall extending from the radially inner platform. Thereby the at least-rib is disposed between the rear partition wall and the trailing edge groove, the flow of the hot gas along the combustion (four) wheel engine has a position to reduce the outer portion of the blade platform The effect of this flow rate of the cooling air in the trailing edge. In a further preferred embodiment of the invention, the ratio of the length of the rib to the width of the trailing edge groove is from 5 to 45:1 to 8 15:1, preferably 6 Η. The above and other objects, features and advantages of the present invention will become apparent from the <RTIgt; 132161.doc 200925388 [Embodiment] FIG. 2 shows a certain blade 1 of the prior art, which mainly comprises a radially outer blade platform 3', a radially inner blade platform 2' first and second side walls 4, 5 (eg Figure 3), and a cooling channel, which is designed between the side walls 4, 5. The stator blade 1 is cooled by the cooling air entering the cooling passage 6. Through the cooling medium, the temperature of the blade metal is maintained below the temperature at which the material can withstand, e.g., to prevent the creation of thermal stresses that it cannot withstand. The cooling air enters the cooling passage 6 through a cooling air inlet 7 in the outer platform 3. The cooling passage extends in a meandering manner between the leading edge region 8 of the blade portion and the trailing edge region 9, as shown in FIG. The cooling passage 6 leads to a trailing edge groove 10' of the trailing edge region of the blade portion from which the cooling air flows. Between the side walls 4, 5, three radially extending partition walls 12' 13' 14 are provided which define the crucible-shaped cooling passages. The front partition wall 12 extends radially inwardly from the outer blade platform 3 with its free end in the region of the radially inner platform 2. The intermediate dividing wall 13 has two free ends extending from a region of the outer blade land 3 to a point in the region of the radially inner platform 2. The rear dividing wall 14 extends from the inner platform 2 to a point in the region of the radially outer blade platform 3. Furthermore, the flow guiding walls 15, μ, 17, 18 are arranged in the cooling channel 6, in the region of the free ends of the dividing walls 12, 13, 14 'which are used to guide the cooling channels 6 This cooling air flow. As shown in Fig. 2, the flow guiding walls 17 and 18 transfer the cooling air flow to a portion of the trailing edge region 9 of the fixed vane 1 along the radial direction 132161.doc 200925388 of the trailing edge region 9. In the trailing edge region 9 of the blade portion, a plurality of posts 11 are arranged between the side walls. The distance between the first partition wall 12 and the trailing edge groove 10 is compared with the flow rate of the cooling air in the region of the radially inner blade land 2 between the rear partition wall 14 and the trailing edge groove 10. The flow rate of the cooling air in the region of the outer blade platform 3 is relatively high. This will result in partial subcooling of the stator blade 1 in the outer blade platform 3 between the front partition wall 12 and the trailing edge groove 1 This condition then causes undesirable thermal stresses in the stator blade. Figure 1 shows a certain blade 1 in accordance with a preferred embodiment of the present invention. Similar elements will be labeled with similar reference numerals. The stator blade 1 is cooled by the cooling medium introduced into the cooling passage 6. The temperature of the blade metal is maintained below the temperature at which the material can withstand through the cooling medium, e.g., to prevent thermal stresses that it cannot withstand. Preferably, compressed air produced by a compressor of the turbine is used as the cooling medium. The cooling medium flows into the cooling passage 6 through a cooling air inlet 7 in the outer platform 3. The cooling passage extends between the leading edge region 8 and the trailing edge region 9 of the blade portion in a meandering manner. As shown in Figure 1. The cooling passage 6 leads to a trailing edge groove 1 in the trailing edge region of the blade portion, and the cooling air flows out from the trailing edge groove. Between the side walls 4, 5 of the weir blade 1, three radially extending blade partitions 12' 13 ' 14 ' are provided which define the weir-shaped cooling passages 6. In accordance with the present invention, at least one rib 19 is disposed in the cooling passage 6 in the region of one of the blades 13161.doc -9^200925388. The term "rib" means a protrusion in a side wall 4 or 5, rather than a dividing wall extending from a side wall* to another side wall 5. The purpose of at least the rib 19 is to reduce flow through the The velocity of the cooling medium flow between the rib 19 and the opposite side walls 4, 5. The rib 19 extends substantially in the radial direction of the stator blade. Preferably, the rib 19 is in the radial direction. Fully extending, as shown in Fig. 2, in any other embodiment, may extend at an angle close to the radial direction in another embodiment. A curved rib 19 is also possible. Preferably, the stator is made by a casting process Finishing, so that the at least one rib 19 is formed as a protrusion on one of the side walls 4, 5. The height of the rib 19, such as the height of the rib protruding perpendicular to the side walls 4, 5, and the rib profile The shape is determined by testing. The profile of the rib 19 is designed such that the flow rate of the cooling medium between the rib 19 and the reverse side wall is reduced. 较佳 Preferably, the turbine flows through the gas turbine engine. The direction of the hot combustion gas 20 'the rib 19 is arranged Between the rear partition wall 14 and the trailing edge groove 10. Depending on the geometry of the cooling passage 6 formed by the weir, the rib 19 can extend radially inward from the radially outer blade platform 3, and/or Extending radially outward from the radially inner blade platform 2. If the rear dividing wall 14 extends from the inner blade platform 2, the rib 19 extends from the outer blade platform 3. If the rear dividing wall 14 is from the outer blade The platform 3 extends and the rib 19 extends from the inner blade platform 2. If the rear dividing wall 4 has two free ends, two ribs 19 can be provided extending from the inner blade platform 2 or the outer blade platform 3, respectively. The region of the free end of the rear partition wall 14 as shown in Fig. 1 provides at least one flow guiding wall 17, 18 for guiding the cooling 132161.doc • 10-200925388 medium in the trailing edge region 9. Preferably, two flow guiding walls 17, 18 are provided. The purpose of the flow guiding walls I?, 18 is to transfer the cooling medium flowing into the trailing edge region along the radial direction of the fixed blade 1. In the embodiment shown in FIG. 1, the rear partition wall 14 between the rear partition wall 14 and the trailing edge groove 1〇 In the radially outer region, the flow guiding walls 17, 18 are arranged for transferring a portion of the trailing edge region 9 radially inward of the cooling air flow. Preferably, the at least one rib 19 is oriented in the chord direction Arranged between the rear partition wall 14 and a plurality of columns 11 between the side walls 4, 5 in the trailing edge region 9. The length of the ribs 19 in the radial direction of the stator blade The ratio of the width 21 (shown in Figure 3) of the trailing edge groove 10 is between 5.45:1 and 8.15:1, preferably 6.8:1. Preferably, the rib 19 is from the radially outer platform. 3 extends radially inwardly to a point radially outward of the radially inner end 22 of the at least one guide vane 14, as shown in the figure! Shown. As described above, the position of the ribs 9 in the trailing edge region 9 and the relative dimension of the ribs 19 advantageously reduce the flow rate of the cooling medium in the region of the blade platforms 2, 3. By reducing the flow rate of the cooling medium, local cooling of the stationary blade 1 in this region is minimized. The many modifications and other embodiments of the invention will be apparent to those skilled in the <RTIgt; Therefore, it is to be understood that the invention is not intended to be limited Although specific items are used, they are used only for the purpose of description and not for the purpose of limitation. BRIEF DESCRIPTION OF THE DRAWINGS The invention will be described by reference to the embodiments illustrated in the accompanying drawings, and reference to the accompanying drawings. The drawings are schematically shown in: Figure 1 is a cross-section of a blade of an embodiment in accordance with one embodiment of the present invention. Figure 2 is a cross section of one of the prior art blades. Figure 3 is a cross-sectional view of one of the stator blades of Figure 1 taken along line A-A. [Main component symbol description]

1 定葉片 2 内平臺 3 外平臺 4 第一側壁 5 第二側壁 6 冷卻通道 7 冷卻入口 8 前緣區域 9 後緣部分 10 後緣槽 11 柱 12 分隔壁 13 分隔壁 14 分隔壁 15 導流壁 16 導流壁 17 導流壁 132161.doc -12- 200925388 18 19 20 21 22 導流壁 肋條 流向 寬度 内端 Ο ❹ 132161.doc -131 fixed blade 2 inner platform 3 outer platform 4 first side wall 5 second side wall 6 cooling channel 7 cooling inlet 8 leading edge region 9 trailing edge portion 10 trailing edge slot 11 column 12 partition wall 13 partition wall 14 partition wall 15 16 Diversion wall 17 Diversion wall 132161.doc -12- 200925388 18 19 20 21 22 Diversion wall rib flow direction width inner end Ο 161 132161.doc -13

Claims (1)

200925388 十、申請專利範圍： 1. 一種燃氣渦輪引擎定葉片（1)，其包括一葉片部分，該禁 片部分具有一第一與一第二側壁（4，5)以及一冷卻通道 (6) ’該冷卻通道被設置於該第一與該第二側壁之間；該 冷卻通道（6)通向該葉片部分之該後緣區域（9)中之一後 - 緣槽（10) ’且至少一個肋條（19)被設置於該定葉片之一 • 葉片平臺U ’ 3)之區域中之該冷卻通道（6)中，該至少一 肋條（19)大致延伸在該定葉片（1)之徑向方向。 〇 2.如請求項1之定葉片，其中該定葉片（1)設置有一徑向外 平臺（3)與一徑向内平臺（2)，該葉片部分在該徑向外平 臺（3)與該徑向内平臺之間延伸，因此該至少一肋條 (19)從該徑向外平臺（3)向内徑向延伸或從該徑向内平臺 (2)向外徑向延伸。 3.如請求項丨或2之定葉片，其中該冷卻通道（6)大致具有一 蜿蜒之形式，且藉由該等側壁（4，5)與分隔壁，13， @ 14)界疋，該等分離壁在該等側壁之間，且大致在該定葉 片（1)之該徑向方向上延伸’因此該後緣區域中之該冷卻 通道藉由該等側壁（1，5)與一後分隔壁（14)界定，該後 刀隔壁從該控向内平臺（2)延伸，因此該至少一肋條（19) ‘ 係、纟該燃氣渦輪引擎之該熱燃燒氣體之該流向(20)上被 佈置在該後分隔壁（14)與該後緣槽（1〇)之間。 1 ·如#求項3之定葉片，其中至少一導流壁（17，18)被設置 於該後刀隔壁(14)之該徑向外區域中用於引導該冷卻 媒介之一部分向内徑向流動於該後緣區域（9)中。 132161.doc 200925388 5·如凊求項丨或2之定葉片，其中該肋條（19)之長度與該後 緣槽（10)之寬度之比為5.45:1至8.15:1，較佳為6.8:1。 6·如》月求項1或2之定葉片，其中該至少一肋條（丨9)在該徑 向方向或接近該徑向方向上充分延伸。 7. 如凊求項1或2之定葉片，其中該至少一肋條（19)從該徑 向外平臺（3)向内徑向延伸至該至少一導流壁（17，18)之 該徑向内端之徑向外部一點。 8. 如請求項1或2之定葉片，其中在該等側壁（4，5)之間延 伸之複數個柱（11)被設置於該至少一肋條（19)與該後緣槽 (1〇)之間。 9·如請求項1或2之定葉片，其係由一鑄造製程完成。 10.如請求項1或2之定葉片，其中該至少一肋條（19)具有— 輪廓，以便減小該肋條（19)之該區域中之冷卻媒介之該 流速。200925388 X. Patent Application Range: 1. A gas turbine engine stator blade (1) comprising a blade portion having a first and a second side wall (4, 5) and a cooling passage (6) 'The cooling passage is disposed between the first and the second side wall; the cooling passage (6) leads to one of the trailing edge regions (9) of the blade portion (9) and at least a rib (19) is disposed in the cooling passage (6) in the region of one of the fixed blades, the blade platform U'3), the at least one rib (19) extending substantially in the diameter of the fixed blade (1) Direction. 〇 2. The blade of claim 1, wherein the fixed blade (1) is provided with a radially outer platform (3) and a radially inner platform (2), the blade portion being on the radially outer platform (3) The radially inner platforms extend between the at least one rib (19) extending radially inwardly from the radially outer platform (3) or radially outwardly from the radially inner platform (2). 3. The blade of claim 2 or 2, wherein the cooling passage (6) has a substantially 蜿蜒 shape and is bounded by the side walls (4, 5) and the dividing wall, 13, @ 14), The separating walls are between the side walls and extend substantially in the radial direction of the stator blade (1) so that the cooling channels in the trailing edge region are separated by the side walls (1, 5) A rear partition wall (14) defines that the rear blade partition extends from the inner control platform (2) such that the at least one rib (19) 'sends the flow of the hot combustion gas of the gas turbine engine (20) The upper portion is disposed between the rear partition wall (14) and the trailing edge groove (1〇). 1) The blade of claim 3, wherein at least one flow guiding wall (17, 18) is disposed in the radially outer region of the rear blade partition (14) for guiding a portion of the cooling medium to an inner diameter The flow is in the trailing edge region (9). 132161.doc 200925388 5. If the blade of the item or the blade is 2, the ratio of the length of the rib (19) to the width of the trailing edge groove (10) is 5.45:1 to 8.15:1, preferably 6.8. :1. 6. The blade of claim 1 or 2, wherein the at least one rib (丨9) extends sufficiently in or near the radial direction. 7. The blade of claim 1 or 2, wherein the at least one rib (19) extends radially inwardly from the radially outer platform (3) to the diameter of the at least one flow guiding wall (17, 18) A little radially outward of the inner end. 8. The blade of claim 1 or 2, wherein a plurality of columns (11) extending between the side walls (4, 5) are disposed in the at least one rib (19) and the trailing edge groove (1〇 )between. 9. The blade of claim 1 or 2 is completed by a casting process. 10. The blade of claim 1 or 2, wherein the at least one rib (19) has a profile to reduce the flow rate of the cooling medium in the region of the rib (19). 132161.doc132161.doc
TW97123226A 2007-06-20 2008-06-20 Cooling of a stator vane of a gas turbine engine TW200925388A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
CH9872007 2007-06-20

Publications (1)

Publication Number Publication Date
TW200925388A true TW200925388A (en) 2009-06-16

Family

ID=38475900

Family Applications (1)

Application Number Title Priority Date Filing Date
TW97123226A TW200925388A (en) 2007-06-20 2008-06-20 Cooling of a stator vane of a gas turbine engine

Country Status (2)

Country Link
TW (1) TW200925388A (en)
WO (1) WO2008155248A1 (en)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113550794A (en) * 2021-09-10 2021-10-26 中国航发湖南动力机械研究所 Multi-cavity efficient cooling structure and cooling method for turbine rotor blade

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8985940B2 (en) 2012-03-30 2015-03-24 Solar Turbines Incorporated Turbine cooling apparatus
KR101691095B1 (en) * 2015-04-20 2016-12-29 연세대학교 산학협력단 Structure of discrete guide vane in the internal cooling channel to control local cooling performance on internal surface
DE102018119572A1 (en) 2018-08-13 2020-02-13 Man Energy Solutions Se Cooling system for active cooling of a turbine blade
US11015455B2 (en) * 2019-04-10 2021-05-25 Pratt & Whitney Canada Corp. Internally cooled turbine blade with creep reducing divider wall

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1350424A (en) * 1971-07-02 1974-04-18 Rolls Royce Cooled blade for a gas turbine engine
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
US5052889A (en) * 1990-05-17 1991-10-01 Pratt & Whintey Canada Offset ribs for heat transfer surface
US5488825A (en) * 1994-10-31 1996-02-06 Westinghouse Electric Corporation Gas turbine vane with enhanced cooling
US6607356B2 (en) * 2002-01-11 2003-08-19 General Electric Company Crossover cooled airfoil trailing edge
AU2005284134B2 (en) * 2004-09-16 2008-10-09 General Electric Technology Gmbh Turbine engine vane with fluid cooled shroud
US7575414B2 (en) * 2005-04-01 2009-08-18 General Electric Company Turbine nozzle with trailing edge convection and film cooling

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113550794A (en) * 2021-09-10 2021-10-26 中国航发湖南动力机械研究所 Multi-cavity efficient cooling structure and cooling method for turbine rotor blade
CN113550794B (en) * 2021-09-10 2022-12-06 中国航发湖南动力机械研究所 Multi-cavity efficient cooling structure and cooling method for turbine rotor blade

Also Published As

Publication number Publication date
WO2008155248A1 (en) 2008-12-24

Similar Documents

Publication Publication Date Title
US8858159B2 (en) Gas turbine engine component having wavy cooling channels with pedestals
US10641107B2 (en) Turbine blade with tip overhang along suction side
JP6334123B2 (en) Rotor blade and rotor blade cooling method
US9017012B2 (en) Ring segment with cooling fluid supply trench
US9206695B2 (en) Cooled turbine blade with trailing edge flow metering
US20130330180A1 (en) Passage channel for a turbomachine and turbomachine
US8662825B2 (en) Cooled aerofoil blade or vane
US8870535B2 (en) Airfoil
JP2006002757A (en) Internally cooled turbo-machine member and method of redesigning its structure
US10267163B2 (en) Airfoil turn caps in gas turbine engines
US11421549B2 (en) Cooled airfoil, guide vane, and method for manufacturing the airfoil and guide vane
JP2005337259A (en) Rotor blade
TW200925388A (en) Cooling of a stator vane of a gas turbine engine
US10563519B2 (en) Engine component with cooling hole
US8757961B1 (en) Industrial turbine stator vane
US7137784B2 (en) Thermally loaded component
US20080181784A1 (en) Convectively cooled gas turbine blade
JP2011038515A (en) Turbine end wall cooling structure
JP5615121B2 (en) Radially fed high temperature axial steam turbine
JP2016130514A (en) Turbine airfoil
US10041352B2 (en) Stator of a turbine of a gas turbine with improved cooling air routing
CN106715833A (en) Cooled turbine blade having internal connecting ribs between the cooling spaces, the ribs having rated break points for reducing thermal stress
US20130022444A1 (en) Low pressure turbine exhaust diffuser with turbulators
RU2479726C1 (en) Turbo machine cooled blade element
US20160186577A1 (en) Cooling configurations for turbine blades