CN85108282A - The multi-chamber airfoil cooling insert that gas turbine blades is used - Google Patents

The multi-chamber airfoil cooling insert that gas turbine blades is used Download PDF

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Publication number
CN85108282A
CN85108282A CN85108282.3A CN85108282A CN85108282A CN 85108282 A CN85108282 A CN 85108282A CN 85108282 A CN85108282 A CN 85108282A CN 85108282 A CN85108282 A CN 85108282A
Authority
CN
China
Prior art keywords
mentioned
chamber
inserting member
turbine
cup
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Withdrawn
Application number
CN85108282.3A
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Chinese (zh)
Other versions
CN1004291B (en
Inventor
索马斯·M·塞苏克
威廉姆·爱得华·诺思
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CBS Corp
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Westinghouse Electric Corp
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Filing date
Publication date
Application filed by Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Publication of CN85108282A publication Critical patent/CN85108282A/en
Publication of CN1004291B publication Critical patent/CN1004291B/en
Expired legal-status Critical Current

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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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • 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
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid

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

Abstract

A kind of gas turbine blades of airfoil shape, it has the inserting member 22 of a single integral body, the back chamber 32,34,36 that the floor 38,40,42 that inserting member is radially extended by polylith is divided into a cup 30 and continues, throttling arrangement 48 is equipped with at back chamber inlet place, the air stream 46 that flows into cup is then unrestricted, make chamber height after the pressure ratio of cup, thereby be higher than the speed of after the low pressure chamber impact air-flows by impacting pore 58 and 60 ejections by its speed of impact air-flow of impacting pore 56 ejections.

Description

The multi-chamber airfoil cooling insert that gas turbine blades is used
The present invention relates to a kind of gas turbine, particularly relate to a kind of hollow gas turbine blades of airfoil shape with forward position wall.This kind gas turbine airfoil shape blade is equipped with an inserting member, and integral body is arranged to the air cooling that blade is provided.
The different sections of known stator blade need different cooling degree.Blade structure is a feature, is in requirement section low or medium cooling degree, and this cooling degree can utilize the impact air-flow of aiming at the blade inwall to reach.Even do not need highly the blade of cooling for those, the required cooling degree of different parts also may be different on the blade, and the forward position district of blade has higher thermal load usually, and in the back near the Hou Yanchu of blade, thermal load may be quite low.
The structure that the purpose of this invention is to provide a kind of blade and inserting member, blade has single inner chamber in this structure, the hollow inserting member of a single integral body wherein has been installed, inserting member has compartment and gas shock pore, and they all are in order to make the impact cooling of blade inwall and external heat load to be interknited and special.
According to the present invention, a kind of gas turbine comprises a kind of hollow turbine blade of airfoil shape and the air cooled hollow inserting member of a single integral body; Turbine blade has a forward position wall, have an air outlet slit line of rabbet joint back along part with the two sides is withstand voltage and the sidewall of anti-suction, and both sidewalls constitutes a single inner chamber that is connected with the above-mentioned air outlet slit line of rabbet joint; The cross section of hollow inserting member substantially with the airfoil shape complementation, inserting member places in the above-mentioned cavity, along the tangential degree that roughly is full of whole above-mentioned cavity that extends to, the dividing plate that has polylith radially to extend in this inserting member, inside is divided into one is positioned at the cup of above-mentioned blade forward position part and the back chamber of separately continuing that at least two communicate with each other, on the inserting member sidewall of all above-mentioned cups and chamber, back, many impact pores are arranged, a longitudinal end of above-mentioned cup and chamber, back is connected with a cooling air source, inserting member has the device that makes the air throttling that flows into chamber, above-mentioned back, make pressure in the above-mentioned cup relatively be higher than pressure in the chamber, above-mentioned back, therefore, its speed of impact air-flow that sprays to above-mentioned forward position partial blade inwall of the pore by above-mentioned cup will be higher than significantly from the speed of the impact air-flow of the pore ejection of chamber, above-mentioned back.
Inserting member is equipped with the dividing plate that polylith radially extends easily, inside is divided into a cup that is positioned at above-mentioned blade forward position part and at least two each other to the back chamber of separately continuing that small part communicates, on the wall of all above-mentioned cups and chamber, back, many impact pores are arranged, a longitudinal end of chamber all is connected with a cooling air source before and after all, and have a device that makes the air throttling that flows into the chamber, back, make pressure in the cup relatively be higher than the pressure in the chamber, above-mentioned back, therefore, its speed of impact air-flow that sprays to above-mentioned forward position partial blade inwall of the pore by above-mentioned cup will be higher than significantly from the speed of the impact air-flow of the pore ejection of chamber, back.
The present invention is described with reference to the drawings below as an example, in the accompanying drawing:
Fig. 1 is the chordwise section figure by blade and inserting member, and it takes from the cross section of the I-I line intercepting along Fig. 2;
Fig. 2 is the part front view and the partial cross section figure of blade and inserting member, corresponding to the view along II among Fig. 1-II line intercepting.
Fig. 1 represents a hollow blade with single inner chamber, and inner chamber is formed by forward position portion, concave side walls 14, convex sidewall 16, and the downstream portion of latter two opposing sidewalls has constituted one and had the back along part 18 of the line of rabbet joint 20.Hot gas passes through the general direction of blade shown in dotted arrow among Fig. 1.
The cross section of the air cooled hollow inserting member 22 of single integral body has airfoil shape, and with the airfoil shape complementation of blade, it is along the tangential degree that roughly is full of the whole blade cavity that extends to.Though inserting member has whole shapes of aerofoil, from Fig. 1, can see, inserting member is part 24 bulging slightly ahead of the curve, is occurring bulging in the back similarly along part 26.The wall of the intermediate portion 28 between the bulging of front and back, and the distance between the blade wall is even substantially.
The compartment 32,34,36 that the inside of whole inserting member 22 is divided into cup 30 and is arranged backward successively by the divider 38,40 and 42 that radially extends, divider 38,40 and 42 also plays structure binding effect.
The radial inner end of all compartments all is closed, and the radial outer end of compartment is communicated with the cooling air source.Can be expressly understood that from Fig. 2 radial outer end 44 is to open wide fully, makes air flow directly into cup 30, shown in arrow among Fig. 2 46.Though back chamber 32,34,36 also is communicated with the cooling air source, but the air-flow that enters the chamber, back utilizes radially extension 48 throttlings of inserting member, radially the extension is made of relative wall 50 and cover plate 52, and cover plate 52 prevents that cooling air from flowing directly into the chamber, back in the mode that cup flows into air.Cooling air enters the chamber, back by 54 throttlings of the aperture on the wall 50.Throttling makes that the air pressure of chamber, back is lower than cup 30.
With reference to two width of cloth accompanying drawings, all compartments all are provided with the impact pore on sidewall.The pore that is provided with on the cup sidewall is represented with numeral 56, as clear seeing among Fig. 1.Impact pore behind the inserting member on the chamber convex sidewall is represented with numeral 58, and the pore on the concave side walls is represented with 60.All impact pore and align, and radially extend substantially.The pore of cup capable 56 spacing each other, the spacing than pore on the convex sidewall of back chamber between capable is wideer, and also the spacing between capable is wideer than most of pores on the concave side walls except first low pressure chamber 32.The chamber is spacious mutually logical by the many pores on two dividing plates or floor 40 and 42 after three.Because dividing plate 40 and 42 will lack a spacing 64 at compartment radial outer end place, therefore also open communication each other between the chamber after three.
Inserting member part ahead of the curve has the pit 66 of outside relief, along part similar pit 68 is arranged in the back, is used for making between the wall of inserting member and the blade wall keeping suitable spacing.
Had as shown in drawings with the arrangement described in the literary composition after, cup 30 is kept the pressure higher than back chamber 32,34,36, what therefore its velocity ratio of cooling blast that penetrates from cup penetrated from the pore of chamber, back wants high, make the higher air-flow of speed penetrate on the front area of higher forward position of the thermal load of blade and convex surface, and the air-flow that the chamber is penetrated after the low pressure penetrates with lower speed, on the relatively low zone of cooling aerofoil blade heat loads on board.The gap ratio that the pore in middle string district is capable is less, obtains than the more uniform cooling effect of wide spacing high velocity air.
The example of the typical pressure level that compartment can be kept, cup are 160 pounds/square inch (1102 * 10 3Crust), back chamber is 155 pounds/square inch (1068 * 10 3Crust), the pressure in the space between inserting member and the relative blade wall is 150 pounds/square inch (1033 * 10 3Crust).

Claims (7)

1, a kind of gas turbine, it comprises a kind of hollow turbine blade of airfoil shape and the air cooled hollow inserting member of a single integral body; It is characterized in that: turbine blade has a forward position wall, have an air outlet slit line of rabbet joint back along part with the two sides is withstand voltage and the sidewall of anti-suction, and both sidewalls constitutes a single inner chamber that communicates with the above-mentioned air outlet slit line of rabbet joint; The cross section of hollow inserting member substantially with the airfoil shape complementation, inserting member places in the above-mentioned cavity, along the tangential degree that roughly is full of whole above-mentioned cavity that extends to, polylith is arranged to the dividing plate that extends in this inserting member, inside is divided into one is positioned at the cup of above-mentioned blade forward position part and the back chamber of separately continuing that at least two communicate with each other, on the inserting member sidewall of all above-mentioned cups and chamber, back, many impact pores are arranged, a longitudinal end of above-mentioned cup and chamber, back is connected with a cooling air source, inserting member has the device that makes the air throttling that flows into chamber, above-mentioned back, make pressure in the above-mentioned cup relatively be higher than pressure in the chamber, above-mentioned back, therefore, its speed of impact air-flow that sprays to above-mentioned forward position partial blade inwall of the gas by above-mentioned cup will be higher than significantly from the speed of the impact air-flow of the pore ejection of chamber, above-mentioned back.
2, a kind of as desired gas turbine in the claim 1, the line-spacing of the impact pore in this turbine in the above-mentioned cup is wideer than the line-spacing that the great majority in the chamber, above-mentioned back impact pore.
3, a kind of as desired gas turbine in claim 1 or 2, above-mentioned throttling arrangement comprises the radially epitaxial part of inserting member at radial outer end place, chamber, above-mentioned back in this turbine, and above-mentioned inserting member many throttle orifices on the epitaxial part radially.
4, a kind of as any one desired gas turbine in the claim 1 to 3, chamber, above-mentioned back comprises three compartments at least in this turbine.
5, a kind of as any one desired gas turbine in the claim 1 to 4, the above-mentioned part that separates is made of hard extension floor in this turbine, do not bore a hole on first floor that separates mutually in the back chamber that above-mentioned cup and first are continued, and two of continuing later have out above the floor
6, a kind of as desired gas turbine in the claim 5, the length that two floors in above-mentioned back extend radially outwardly in this turbine is less than above-mentioned first floor, makes chamber, above-mentioned back in its radially outer open communication each other.
7, a kind of gas turbine that comprises the hollow turbine blade of airfoil shape, its structure and purposes are as described herein substantially and illustrated with reference to this paper accompanying drawing.
CN85108282.3A 1984-11-15 1985-11-14 Multi-chamber airfoil cooling insert for turbine vane Expired CN1004291B (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US67184684A 1984-11-15 1984-11-15
US671,846 1984-11-15

Publications (2)

Publication Number Publication Date
CN85108282A true CN85108282A (en) 1986-08-27
CN1004291B CN1004291B (en) 1989-05-24

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
CN85108282.3A Expired CN1004291B (en) 1984-11-15 1985-11-14 Multi-chamber airfoil cooling insert for turbine vane

Country Status (9)

Country Link
EP (1) EP0182588B1 (en)
JP (1) JPS61126302A (en)
KR (1) KR860004224A (en)
CN (1) CN1004291B (en)
CA (1) CA1221915A (en)
DE (1) DE3565298D1 (en)
IN (1) IN163070B (en)
IT (1) IT1186049B (en)
MX (1) MX161444A (en)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101825115A (en) * 2010-03-31 2010-09-08 北京航空航天大学 Blade with built-in bed frame-type pneumatic damping device
CN102251814A (en) * 2003-02-27 2011-11-23 通用电气公司 Gas turbine engine turbine nozzle bifurcated impingement baffle
CN103089326A (en) * 2011-10-31 2013-05-08 通用电气公司 Method and apparatus for cooling gas turbine rotor blades
CN104088673A (en) * 2008-11-07 2014-10-08 三菱重工业株式会社 Vane for turbine
CN104254670A (en) * 2012-04-27 2014-12-31 通用电气公司 Durable turbine vane
CN104254669A (en) * 2011-12-06 2014-12-31 西门子公司 Turbine blade incorporating trailing edge cooling design
CN106661945A (en) * 2014-09-04 2017-05-10 西门子公司 Internal Cooling System With Insert Forming Nearwall Cooling Channels In An Aft Cooling Cavity Of A Gas Turbine Airfoil
CN108868899A (en) * 2017-05-09 2018-11-23 通用电气公司 Impingement insert
CN109983203A (en) * 2016-11-23 2019-07-05 通用电气公司 Cooling structure for turbine part

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GB2189553B (en) * 1986-04-25 1990-05-23 Rolls Royce Cooled vane
US5212940A (en) * 1991-04-16 1993-05-25 General Electric Company Tip clearance control apparatus and method
US5207556A (en) * 1992-04-27 1993-05-04 General Electric Company Airfoil having multi-passage baffle
US6183192B1 (en) * 1999-03-22 2001-02-06 General Electric Company Durable turbine nozzle
DE10004128B4 (en) 2000-01-31 2007-06-28 Alstom Technology Ltd. Air-cooled turbine blade
US6609880B2 (en) * 2001-11-15 2003-08-26 General Electric Company Methods and apparatus for cooling gas turbine nozzles
US6652220B2 (en) * 2001-11-15 2003-11-25 General Electric Company Methods and apparatus for cooling gas turbine nozzles
US7871246B2 (en) * 2007-02-15 2011-01-18 Siemens Energy, Inc. Airfoil for a gas turbine
EP2706195A1 (en) * 2012-09-05 2014-03-12 Siemens Aktiengesellschaft Impingement tube for gas turbine vane with a partition wall
GB201417476D0 (en) 2014-10-03 2014-11-19 Rolls Royce Plc Internal cooling of engine components
US10329932B2 (en) 2015-03-02 2019-06-25 United Technologies Corporation Baffle inserts
US9849510B2 (en) 2015-04-16 2017-12-26 General Electric Company Article and method of forming an article
US9976441B2 (en) 2015-05-29 2018-05-22 General Electric Company Article, component, and method of forming an article
US9850763B2 (en) 2015-07-29 2017-12-26 General Electric Company Article, airfoil component and method for forming article
US10253986B2 (en) 2015-09-08 2019-04-09 General Electric Company Article and method of forming an article
US10087776B2 (en) 2015-09-08 2018-10-02 General Electric Company Article and method of forming an article
US10739087B2 (en) * 2015-09-08 2020-08-11 General Electric Company Article, component, and method of forming an article
US10655477B2 (en) 2016-07-26 2020-05-19 General Electric Company Turbine components and method for forming turbine components
US10260363B2 (en) 2016-12-08 2019-04-16 General Electric Company Additive manufactured seal for insert compartmentalization
US10480347B2 (en) 2018-01-18 2019-11-19 United Technologies Corporation Divided baffle for components of gas turbine engines
US10697309B2 (en) 2018-04-25 2020-06-30 Raytheon Technologies Corporation Platform cover plates for gas turbine engine components

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US2873944A (en) * 1952-09-10 1959-02-17 Gen Motors Corp Turbine blade cooling
FR1177035A (en) * 1957-05-28 1959-04-20 Snecma Method and device for cooling machine parts
US3540810A (en) * 1966-03-17 1970-11-17 Gen Electric Slanted partition for hollow airfoil vane insert
US3891348A (en) * 1972-04-24 1975-06-24 Gen Electric Turbine blade with increased film cooling
GB1587401A (en) * 1973-11-15 1981-04-01 Rolls Royce Hollow cooled vane for a gas turbine engine
CH584346A5 (en) * 1974-11-08 1977-01-31 Bbc Sulzer Turbomaschinen
US4297077A (en) * 1979-07-09 1981-10-27 Westinghouse Electric Corp. Cooled turbine vane

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN102251814B (en) * 2003-02-27 2015-03-11 通用电气公司 Gas turbine engine turbine nozzle bifurcated impingement baffle
CN102251814A (en) * 2003-02-27 2011-11-23 通用电气公司 Gas turbine engine turbine nozzle bifurcated impingement baffle
CN104088673A (en) * 2008-11-07 2014-10-08 三菱重工业株式会社 Vane for turbine
CN104088673B (en) * 2008-11-07 2016-03-09 三菱日立电力***株式会社 turbine blade
CN101825115A (en) * 2010-03-31 2010-09-08 北京航空航天大学 Blade with built-in bed frame-type pneumatic damping device
CN103089326A (en) * 2011-10-31 2013-05-08 通用电气公司 Method and apparatus for cooling gas turbine rotor blades
CN104254669A (en) * 2011-12-06 2014-12-31 西门子公司 Turbine blade incorporating trailing edge cooling design
CN104254670A (en) * 2012-04-27 2014-12-31 通用电气公司 Durable turbine vane
US9506351B2 (en) 2012-04-27 2016-11-29 General Electric Company Durable turbine vane
CN106661945A (en) * 2014-09-04 2017-05-10 西门子公司 Internal Cooling System With Insert Forming Nearwall Cooling Channels In An Aft Cooling Cavity Of A Gas Turbine Airfoil
US9863256B2 (en) 2014-09-04 2018-01-09 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of an airfoil usable in a gas turbine engine
CN109983203A (en) * 2016-11-23 2019-07-05 通用电气公司 Cooling structure for turbine part
CN109983203B (en) * 2016-11-23 2021-12-21 通用电气公司 Cooling structure for turbine component
CN108868899A (en) * 2017-05-09 2018-11-23 通用电气公司 Impingement insert

Also Published As

Publication number Publication date
IT8522785A0 (en) 1985-11-11
MX161444A (en) 1990-09-27
JPH0379522B2 (en) 1991-12-19
CN1004291B (en) 1989-05-24
JPS61126302A (en) 1986-06-13
IN163070B (en) 1988-08-06
IT1186049B (en) 1987-11-18
DE3565298D1 (en) 1988-11-03
CA1221915A (en) 1987-05-19
EP0182588B1 (en) 1988-09-28
KR860004224A (en) 1986-06-18
EP0182588A1 (en) 1986-05-28

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