JP2007154887A - Method for reducing axial compressor blade tipflow and turbine machine - Google Patents

Method for reducing axial compressor blade tipflow and turbine machine Download PDF

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Publication number
JP2007154887A
JP2007154887A JP2006323501A JP2006323501A JP2007154887A JP 2007154887 A JP2007154887 A JP 2007154887A JP 2006323501 A JP2006323501 A JP 2006323501A JP 2006323501 A JP2006323501 A JP 2006323501A JP 2007154887 A JP2007154887 A JP 2007154887A
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Prior art keywords
tip
channel
blade
air
airfoil
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JP2006323501A
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Japanese (ja)
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Zhifeng Dong
チフェン・ドン
John Joseph Rahaim
ジョン・ジョセフ・ラーイム
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General Electric Co
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General Electric Co
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    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/684—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid injection
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/307—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the tip of a rotor blade

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

<P>PROBLEM TO BE SOLVED: To provide a method and a device for reducing an axial compressor blade tipflow. <P>SOLUTION: In this turbine machine T1 of such type that has a high pressure compressor arranged in a casing, a plurality of rotary compressor blades 10 to 14 having at least one air channel 10C to 14C formed in the blades 10 to 14 to communicate air flow from a basic part to a tip part are provided, and at least one air channel 10C to 14C lets the air out of an inside region adjacent to a disc 20, pressurizes the air, and carries the removed and pressurized air into a region between the tip part of the blade and the casing C1 to shut off air flow crossing the tip part of the blade 10 to 14. <P>COPYRIGHT: (C)2007,JPO&INPIT

Description

本発明は、ブレード及びベーンのような翼形部における軸流圧縮機先端流を減少させるための方法及びタービン機械に関する。   The present invention relates to a method and turbine machine for reducing axial compressor tip flow in airfoils such as blades and vanes.

タービンエンジンの圧縮機区域におけるブレード先端流は、圧縮機効率及び失速マージンの低下を引き起こす。加えて、シール空洞内におけるベーンとブレードとの間の内部流路に沿った流れの再循環もまた、圧縮機性能を低下させる。先端流を減少させるための１つの従来技術の方法は、ブレード先端間隙を小さくすることである。それは、機械的及び／又は熱的方法を使用してケーシングとベーンとのインタフェースを制御することを含む様々な方法で行われる。これらの方法は、先端部摩擦、過剰摩耗及びエンジン効率の低下を引き起こすおそれがある。
米国特許第６，８７７，９５３号公報 米国特許第６，５７４，９６５号公報 米国特許第５，３５８，３７８号公報 Blade tip flow in the compressor area of a turbine engine causes a reduction in compressor efficiency and stall margin. In addition, flow recirculation along the internal flow path between the vanes and blades within the seal cavity also reduces compressor performance. One prior art method for reducing tip flow is to reduce the blade tip clearance. It can be done in a variety of ways, including controlling the casing-vane interface using mechanical and / or thermal methods. These methods can cause tip friction, excessive wear and reduced engine efficiency.
US Pat. No. 6,877,953 US Pat. No. 6,574,965 US Pat. No. 5,358,378

本発明の１つの態様によると、間隔を置いて近接配置したケーシング内で回転するタービン翼形部の先端部と該ケーシングとの間の空気流を減少させる方法を提供し、本方法は、翼形部の基部に隣接する入口開口と翼形部先端部上の出口開口とを有する半径方向に延びるチャネルを翼形部内に設ける段階を含む。空気は、翼形部の基部に隣接した領域から抜取られかつ加圧され、チャネル内に導入され、チャネルを通して翼形部先端部まで運ばれる。空気は、翼形部の正圧側から負圧側への軸方向空気流に抗するのに十分な圧力の下で、翼形部先端部内の出口開口を通ってチャネルから翼形部の先端部とケーシングとの間の区域内に流出する。   According to one aspect of the present invention, there is provided a method for reducing air flow between a tip of a turbine airfoil that rotates within a closely spaced casing and the casing, the method comprising: Providing in the airfoil a radially extending channel having an inlet opening adjacent to the base of the airfoil and an outlet opening on the airfoil tip. Air is drawn from the area adjacent to the base of the airfoil and pressurized, introduced into the channel, and carried through the channel to the airfoil tip. The air passes from the channel through the outlet opening in the airfoil tip and from the channel to the airfoil tip under sufficient pressure to resist axial airflow from the pressure side to the suction side of the airfoil. It flows into the area between the casing.

本発明の別の態様は、間隔を置いて近接配置したケーシング内で回転するタービン翼形部の先端部と該ケーシングとの間の空気流を減少させる方法を提供し、本方法は、翼形部の前縁側で該翼形部の基部に隣接する入口開口と翼形部先端部上の出口開口とを有する第１の半径方向に延びるチャネルを設ける段階と、翼形部の後縁側で該翼形部の基部に隣接する入口開口と翼形部先端部上の出口開口とを有する第２の半径方向に延びるチャネルを設ける段階とを含む。空気は、翼形部の基部に隣接する領域からチャネル内に抜取られかつチャネルを通して翼形部先端部まで圧送される。空気は、翼形部の正圧側から負圧側への軸方向空気流に抗するのに十分な圧力の下で、翼形部先端部内の出口開口を通ってチャネルから翼形部の先端部とケーシングとの間の区域内に流出する。   Another aspect of the present invention provides a method for reducing air flow between a tip of a turbine airfoil rotating within a closely spaced casing and the casing, the method comprising: Providing a first radially extending channel having an inlet opening adjacent to the base of the airfoil on the leading edge side of the airfoil and an outlet opening on the airfoil tip; and on the trailing edge side of the airfoil Providing a second radially extending channel having an inlet opening adjacent to the base of the airfoil and an outlet opening on the airfoil tip. Air is drawn into the channel from the region adjacent to the base of the airfoil and is pumped through the channel to the airfoil tip. The air passes from the channel through the outlet opening in the airfoil tip and from the channel to the airfoil tip under sufficient pressure to resist axial airflow from the pressure side to the suction side of the airfoil. It flows into the area between the casing.

本発明のさらに別の態様では、タービン機械用圧縮機翼形部を提供し、本圧縮機翼形部は、翼形部基部と、翼形部先端部と、翼形部基部に隣接する翼形部内の空気入口開口から翼形部先端部内の出口開口まで半径方向に延びる空気流チャネルとを含み、空気流チャネルは、翼形部の正圧側から該翼形部の負圧側への軸方向空気流に対して空気遮断を行い、それによって圧縮機翼形部先端流を減少させるようにする。   In yet another aspect of the present invention, a compressor airfoil for a turbine machine is provided, the compressor airfoil including an airfoil base, an airfoil tip, and an airfoil adjacent to the airfoil base. An airflow channel extending radially from an air inlet opening in the airfoil to an outlet opening in the airfoil tip, wherein the airflow channel is axial from the pressure side of the airfoil to the suction side of the airfoil Air blocking is applied to the air flow, thereby reducing compressor airfoil tip flow.

添付の図面に関連させて、以下に本発明を説明する。   The present invention is described below with reference to the accompanying drawings.

次に具体的に図面を参照すると、図１には、タービンエンジンＴ１の軸流圧縮機セクションの部分断面を示しており、この図により、本発明による軸流圧縮機ブレード先端流を制御するための方法及び装置を説明する。タービンエンジン「Ｔ１」は、ケーシングＣ１内に圧縮機ブレード１０〜１４と中間配置ステータベーン１５〜１９とを含む。圧縮機ブレード１０〜１４は、それぞれの前縁１０Ａ〜１４Ａを含む。   Referring now specifically to the drawings, FIG. 1 shows a partial cross section of an axial compressor section of a turbine engine T1 for controlling the axial compressor blade tip flow according to the present invention. The method and apparatus will be described. The turbine engine “T1” includes compressor blades 10 to 14 and intermediately arranged stator vanes 15 to 19 in a casing C1. The compressor blades 10-14 include respective leading edges 10A-14A.

図２に示すように、ブレード１０は、明瞭にするために拡大詳細図として示しており、またブレード１１〜１４を代表するものである。空気は、ブレード１０の前縁１０Ａ側の区域からディスク２０内の孔１０Ｂを通して抜取られかつ加圧される。孔１０Ｂは、チャネル１０Ｃと連通しており、チャネル１０Ｃは、ブレード１０を貫通して先端部まで半径方向外向きに延び、先端部においてチャネル１０Ｃは、孔１０Ｄを通って出る。チャネル１０Ｃは、ブレード１０の先端部を出る前に分岐させることができることに留意されたい。チャネル１０Ｃの寸法並びに分岐の位置及び数は、ブレードの寸法、形状及びボリュームと、エンジンの性能、定格、先端部間隙及び類似要因とに基づいて実験的に決定される。図面では、図面の縮尺に関連して先端部間隙の実物表示を示すことができないほどに先端部間隙が非常に小さいことに留意されたい。   As shown in FIG. 2, the blade 10 is shown as an enlarged detail for clarity and is representative of the blades 11-14. Air is extracted from the area on the leading edge 10A side of the blade 10 through the hole 10B in the disk 20 and pressurized. The hole 10B communicates with the channel 10C, which extends radially outwardly through the blade 10 to the tip, where the channel 10C exits through the hole 10D. Note that channel 10C may be branched before exiting the tip of blade 10. The size of channel 10C and the location and number of branches are determined empirically based on the size, shape and volume of the blades, engine performance, rating, tip clearance and similar factors. Note that in the drawing, the tip gap is so small that no real representation of the tip gap can be shown in relation to the scale of the drawing.

図３を参照すると、タービンエンジン「Ｔ２」は、ケーシングＣ２内に圧縮機ブレード３０〜３４と中間配置ステータベーン３５〜３９とを含む。圧縮機ブレード３０〜３４は、それぞれの後縁３０Ａ〜３４Ａを含む。   Referring to FIG. 3, turbine engine “T2” includes compressor blades 30-34 and intermediately disposed stator vanes 35-39 in casing C2. The compressor blades 30-34 include respective trailing edges 30A-34A.

図４では、ブレード３１は、明瞭にするために拡大詳細図として示しており、またブレード３０及び３２〜３４を代表するものである。空気は、ブレード３１の後縁３１Ａ側の区域からディスクリム４０内の孔３１Ｂを通して抜取られる。孔３１Ｂは、チャネル３１Ｃと連通しており、チャネル３１Ｃは、ブレード３１を貫通して先端部まで半径方向外向きに延び、先端部において孔３１Ｄを通って出る前に、チャネル３１Ｃは分岐させるのが好ましい。   In FIG. 4, the blade 31 is shown as an enlarged detail for clarity and is representative of blades 30 and 32-34. Air is extracted from the area on the rear edge 31A side of the blade 31 through the hole 31B in the disc rim 40. The hole 31B communicates with the channel 31C, the channel 31C extends radially outward through the blade 31 to the tip, and the channel 31C is branched before exiting through the hole 31D at the tip. Is preferred.

次に図５を参照すると、タービンエンジン「Ｔ３」は、ケーシングＣ３内に圧縮機ブレード５０〜５４と中間配置ステータベーン５５〜５９とを含む。圧縮機ブレード５０〜５４は、それぞれの前縁５０Ａ〜５４Ａとそれぞれの後縁５０Ｂ〜５４Ｂとを含む。   Referring now to FIG. 5, turbine engine “T3” includes compressor blades 50-54 and intermediately disposed stator vanes 55-59 within casing C3. The compressor blades 50-54 include respective leading edges 50A-54A and respective trailing edges 50B-54B.

図６は、明瞭にするために拡大詳細図として示しかつブレード５１及び５２〜５４を代表するものであるブレード５２を示す。空気は、ブレード５２の前縁側５２Ａ及び後縁側５２Ｂの区域の両方からディスクリム６０内の孔５２Ｃ及び５２Ｄを通して抜取られる。孔５２Ｃ及び５２Ｄは、それぞれチャネル５２Ｅ及び５２Ｆと連通しており、チャネル５２Ｅ及び５２Ｆは、ブレード５２を貫通して先端部まで半径方向外向きに延び、先端部において孔５２Ｇを通って出る前に、それらのチャネルは分岐させるのが好ましい。   FIG. 6 shows a blade 52 which is shown as an enlarged detail for clarity and is representative of blades 51 and 52-54. Air is drawn from both the leading edge side 52A and trailing edge side 52B areas of the blade 52 through holes 52C and 52D in the disc rim 60. The holes 52C and 52D communicate with channels 52E and 52F, respectively, which extend radially outward through the blade 52 to the tip and before exiting through the hole 52G at the tip. The channels are preferably branched.

上述の実施形態の各々では、ブレード先端部で吐出される空気は、ブレード先端部とケーシングとの間の先端部間隙の領域における空気流を空気力学的に遮断することによってブレード先端流を減少させるか又は阻止する。内部流路からの空気は、上述のように先端部間隙に持ち込まれて、空気遮断を形成する。抜取った空気の圧力は、圧縮機ロータのポンプ（圧送）作用によって増大して、先端部においてブレードから流出する時には、正圧側から負圧側にブレード先端部を横切る空気流に抗する。   In each of the above embodiments, the air discharged at the blade tip reduces the blade tip flow by aerodynamically blocking the air flow in the region of the tip gap between the blade tip and the casing. Or block. Air from the internal channel is brought into the tip gap as described above to form an air block. The pressure of the extracted air is increased by the pump (pumping) action of the compressor rotor, and resists the air flow across the blade tip from the positive pressure side to the negative pressure side when it flows out of the blade at the tip.

上述の方法は、低圧圧縮機（ブースタ）及び高圧圧縮機の両方に適用することができる。これらの方法を使用する場合に、負担となる流量損失は存在しない。さらに、ブレード先端部とケーシングとの間の狭い運転間隙によるのではなく空気力学的空気遮断により空気流を減少させることによって、ブレード先端部とケーシングとの間でのより大きな組立間隙を設定しかつ維持することが可能になる。従って、ベーンとブレードとの間の内部流路における再循環であるので、ブレード先端部摩擦が減少する。抜取った空気は、内部流路からブレード先端部に連続的に圧送され、従ってエンジンの運転時には常にブレード先端部に対して連続的な空気遮断を行う。   The method described above can be applied to both low pressure compressors (boosters) and high pressure compressors. There is no flow loss that is a burden when using these methods. Furthermore, by setting the larger assembly gap between the blade tip and the casing by reducing the air flow by aerodynamic air blockage rather than by a narrow operating gap between the blade tip and the casing, and It becomes possible to maintain. Accordingly, the blade tip friction is reduced due to recirculation in the internal flow path between the vane and the blade. The extracted air is continuously pumped from the internal flow path to the blade tip, and therefore, continuously shuts off the air from the blade tip during engine operation.

本出願において説明したこの方法はまた、ブリスク（ブレード一体形ディスク）、傾斜ダブテール結合形ブレード又は円周方向ダブテール結合形ブレードにも適用できる。   This method described in this application can also be applied to blisks (blade integrated discs), inclined dovetail coupled blades or circumferential dovetail coupled blades.

以上、軸流圧縮機ブレード先端流を制御するための方法及び装置を説明している。本発明の様々な細部は、本発明の技術的範囲から逸脱せずに変更することができる。さらに、本発明の好ましい実施形態及び本発明を実施するための最良の形態についての上述の説明は、例示の目的のためだけに示しているものであって、本発明を限定するためのものではなく、本発明は、特許請求の範囲によって定まる。   Thus, a method and apparatus for controlling axial compressor blade tip flow has been described. Various details of the invention may be changed without departing from the scope of the invention. Furthermore, the foregoing descriptions of the preferred embodiments of the present invention and the best mode for carrying out the present invention are presented for purposes of illustration only and are not intended to limit the present invention. Instead, the invention is defined by the claims.

本発明の１つの実施形態を示すタービンエンジンの軸流圧縮機セクションの部分断面図。1 is a partial cross-sectional view of an axial flow compressor section of a turbine engine illustrating one embodiment of the present invention. 図１に示す圧縮機の一部分の拡大部分断面図。FIG. 2 is an enlarged partial cross-sectional view of a part of the compressor shown in FIG. 1. 本発明の別の実施形態を示すタービンエンジンの圧縮機セクションの部分断面図。FIG. 4 is a partial cross-sectional view of a compressor section of a turbine engine illustrating another embodiment of the present invention. 図３に示す圧縮機の一部分の拡大部分断面図。FIG. 4 is an enlarged partial cross-sectional view of a part of the compressor shown in FIG. 3. 本発明のさらに別の実施形態を示すタービンエンジンの圧縮機セクションの部分断面図。FIG. 6 is a partial cross-sectional view of a compressor section of a turbine engine showing yet another embodiment of the present invention. 図５に示す圧縮機の一部分の拡大部分断面図。FIG. 6 is an enlarged partial cross-sectional view of a part of the compressor shown in FIG. 5.

符号の説明Explanation of symbols

Ｔ１ タービン機械
Ｔ２ タービン機械
Ｔ３ タービン機械
Ｃ１ ケーシング
Ｃ３ ケーシング
１０ ブレード
１０Ａ 前縁
１０Ｂ 入口開口
１０Ｃ チャネル
１０Ｄ 出口開口
１１ ブレード
１１Ｃ チャネル
１２ ブレード
１２Ｃ チャネル
１３ ブレード
１３Ｃ チャネル
１４ ブレード
１４Ｃ チャネル
１５ 固定ベーン
１６ 固定ベーン
１７ 固定ベーン
１８ 固定ベーン
１９ 固定ベーン
２０ ディスク
３１ ブレード
３１Ａ 後縁
３１Ｂ 入口開口
３１Ｃ チャネル
３１Ｄ 出口開口
５０ ブレード
５０Ａ 前縁
５０Ｂ 後縁
５０Ｃ 入口開口
５０Ｄ 入口開口
５０Ｆ チャネル
５２ ブレード
５２Ａ 前縁
５２Ｂ 後縁
５２Ｃ 入口開口
５２Ｄ 入口開口
５２Ｅ 第１のチャネル
５２Ｆ 第２のチャネル
５２Ｇ 出口孔 T1 turbine machine T2 turbine machine T3 turbine machine C1 casing C3 casing 10 blade 10A leading edge 10B inlet opening 10C channel 10D outlet opening 11 blade 11C channel 12 blade 12C channel 13 blade 13C channel 14 blade 14C channel 15 fixed vane 16 fixed vane 16 fixed vane 16 fixed vane 16 Vane 18 Fixed vane 19 Fixed vane 20 Disk 31 Blade 31A Rear edge 31B Inlet opening 31C Channel 31D Outlet opening 50 Blade 50A Front edge 50B Rear edge 50C Inlet opening 50D Inlet opening 50F Channel 52 Blade 52A Front edge 52B Rear edge 52C Inlet opening 52 Inlet opening 52E First channel 52F Second channel 52G Outlet hole

Claims (10)

ケーシング（Ｃ１）内に配置された高圧圧縮機と、それぞれのブレード先端部と中心ディスク（２０）に固定されたそれぞれのブレード基部とを有する複数の回転圧縮機ブレード（１０〜１４）と、前記ブレード（１０〜１４）のそれぞれのブレード間に配置された複数の固定ベーン（１５〜１９）とを有するタイプのタービン機械（Ｔ１）であって、
前記基部から前記先端部まで空気流れ連通させるように前記ブレード（１０〜１４）のそれぞれのブレード内に形成された少なくとも１つの空気チャネル（１０Ｃ、１１Ｃ、１２Ｃ、１３Ｃ、１４Ｃ）を含み、
前記少なくとも１つの空気チャネル（１０Ｃ、１１Ｃ、１２Ｃ、１３Ｃ、１４Ｃ）が、前記ディスク（２０）に隣接する内部区域から空気を抜取りかつ該空気を加圧し、前記抜取りかつ加圧した空気を該チャネル（１０Ｃ、１１Ｃ、１２Ｃ、１３Ｃ、１４Ｃ）を通して前記ブレード先端部と前記ケーシング（Ｃ１）との間の区域内に運んで、前記ブレード（１０〜１４）の先端部からの空気流を遮断するようにする、
タービン機械（Ｔ１）。 A plurality of rotary compressor blades (10-14) having a high-pressure compressor disposed in the casing (C1), a respective blade tip and a respective blade base fixed to the central disk (20); A turbine machine (T1) of the type having a plurality of stationary vanes (15-19) disposed between each of the blades (10-14),
Including at least one air channel (10C, 11C, 12C, 13C, 14C) formed in each blade of the blades (10-14) to provide air flow communication from the base to the tip;
The at least one air channel (10C, 11C, 12C, 13C, 14C) draws air from an internal area adjacent to the disk (20) and pressurizes the air, and draws the drawn and pressurized air into the channel (10C, 11C, 12C, 13C, 14C) is carried into the area between the blade tip and the casing (C1) to block air flow from the blade (10-14) tip. To
Turbine machine (T1). 前記チャネル（１０Ｃ）が、それぞれの複数の出口孔（１０Ｄ）を通して前記ブレード先端部を出る、請求項１記載のタービン機械（Ｔ１）。   The turbine machine (T1) of claim 1, wherein the channel (10C) exits the blade tip through a respective plurality of outlet holes (10D). 前記チャネル（１０Ｃ）が、前記ブレード（１０）の前縁（１０Ａ）上に空気入口開口（１０Ｂ）を含む、請求項１記載のタービン機械（Ｔ１）。   The turbine machine (T1) according to claim 1, wherein the channel (10C) includes an air inlet opening (10B) on a leading edge (10A) of the blade (10). 前記チャネル（３１Ｃ）が、前記ブレード（３１）の後縁（３１Ａ）側に空気入口開口（３１Ｂ）を含む、請求項１記載のタービン機械（Ｔ２）。   The turbine machine (T2) of claim 1, wherein the channel (31C) includes an air inlet opening (31B) on a trailing edge (31A) side of the blade (31). 前記チャネル（５０Ｆ）が、前記ブレード（５０）の前縁（５０Ａ）側の空気入口開口（５０Ｃ）と後縁（５０Ｂ）側の空気入口開口（５０Ｄ）とを含む、請求項１記載のタービン機械（Ｔ３）。   The turbine of claim 1, wherein the channel (50F) includes an air inlet opening (50C) on a leading edge (50A) side and an air inlet opening (50D) on a trailing edge (50B) side of the blade (50). Machine (T3). 前記ブレード（５２）の前縁（５２Ａ）側に空気入口開口（５２Ｃ）を有する第１のチャネル（５２Ｅ）と、
前記ブレード（５２）の後縁（５２Ｂ）側に空気入口開口（５２Ｄ）を有する第２のチャネル（５２Ｆ）と、
を含む、請求項１記載のタービン機械（Ｔ３）。 A first channel (52E) having an air inlet opening (52C) on the leading edge (52A) side of the blade (52);
A second channel (52F) having an air inlet opening (52D) on the trailing edge (52B) side of the blade (52);
The turbine machine (T3) according to claim 1, comprising: 前記チャネル（５０Ｆ）が、前記ブレード先端部内に複数の出口孔（５２Ｇ）を含む、請求項３記載のタービン機械（Ｔ１）。   The turbine machine (T1) according to claim 3, wherein the channel (50F) includes a plurality of outlet holes (52G) in the blade tip. 間隔を置いて近接配置したケーシング（Ｃ１）内で回転するタービン翼形部（１０）の先端部と該ケーシングとの間の空気流を減少させる方法であって、
（ａ）前記翼形部（１０）の基部に隣接する入口開口（１０Ｂ）と前記翼形部先端部上の出口開口（１０Ｄ）とを有する半径方向に延びるチャネル（１０Ｃ）を該翼形部（１０）内に設ける段階と、
（ｂ）前記翼形部（１０）の基部に隣接した領域から前記チャネル（１０Ｃ）内に空気を抜取りかつ該空気を加圧する段階と、
（ｃ）前記チャネル（１０Ｃ）を通して前記翼形部先端部まで空気を運ぶ段階と、
（ｄ）前記翼形部（１０）の先端部を横切る軸方向空気流に抗するのに十分な圧力の下で、前記翼形部先端部内の出口開口（１０Ｄ）を通して前記翼形部の先端部と前記ケーシング（Ｃ１）との間の区域内に空気を吐出する段階と、
含む方法。 A method of reducing air flow between a tip of a turbine airfoil (10) rotating in a closely spaced casing (C1) and the casing;
(A) a radially extending channel (10C) having an inlet opening (10B) adjacent to the base of the airfoil (10) and an outlet opening (10D) on the tip of the airfoil; (10) providing in the stage;
(B) extracting air into the channel (10C) from a region adjacent to the base of the airfoil (10) and pressurizing the air;
(C) carrying air through the channel (10C) to the airfoil tip;
(D) the tip of the airfoil through an outlet opening (10D) in the tip of the airfoil under pressure sufficient to resist axial airflow across the tip of the airfoil (10). Discharging air into the area between the part and the casing (C1);
Including methods. 前記翼形部（１０）の基部に隣接する入口開口（１０Ｂ）と前記翼形部先端部上の出口開口（１０Ｄ）とを有する半径方向に延びるチャネル（１０Ｃ）を設ける段階が、前記翼形部（１０）の前縁（１０Ａ）側に前記入口開口（１０Ｂ）を形成する段階を含む、請求項８記載の方法。   Providing a radially extending channel (10C) having an inlet opening (10B) adjacent to a base of the airfoil (10) and an outlet opening (10D) on the airfoil tip; The method according to claim 8, comprising the step of forming the inlet opening (10B) on the leading edge (10A) side of the part (10). 前記翼形部（３１）の基部に隣接する入口開口（３１Ｂ）と前記翼形部先端部上の出口開口（３１Ｄ）とを有する半径方向に延びるチャネル（３１Ｃ）を設ける段階が、前記翼形部（３１）の後縁（３１Ａ）側に前記入口開口（３１Ｂ）を形成する段階を含む、請求項８記載の方法。   Providing a radially extending channel (31C) having an inlet opening (31B) adjacent to the base of the airfoil (31) and an outlet opening (31D) on the airfoil tip; The method according to claim 8, comprising the step of forming the inlet opening (31B) on the rear edge (31A) side of the part (31).
JP2006323501A 2005-11-30 2006-11-30 Method for reducing axial compressor blade tipflow and turbine machine Pending JP2007154887A (en)

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