CN86104500A - The improvement structure of combustion gas turbine moving vane coolant channel - Google Patents

The improvement structure of combustion gas turbine moving vane coolant channel Download PDF

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Publication number
CN86104500A
CN86104500A CN 86104500 CN86104500A CN86104500A CN 86104500 A CN86104500 A CN 86104500A CN 86104500 CN86104500 CN 86104500 CN 86104500 A CN86104500 A CN 86104500A CN 86104500 A CN86104500 A CN 86104500A
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China
Prior art keywords
section
flow area
duct
blade
mentioned
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Pending
Application number
CN 86104500
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Chinese (zh)
Inventor
保罗·克拉伦斯·霍尔登
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CBS Corp
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Westinghouse Electric Corp
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Publication of CN86104500A publication Critical patent/CN86104500A/en
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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
    • 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
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/231—Three-dimensional prismatic cylindrical
    • 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
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/232—Three-dimensional prismatic conical

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

Abstract

A moving vane of combustion gas turbine has a blade part, manyly passes blade from the coolant channel that extends radially outwardly.The freezing mixture duct is taper, at the blade intermediate section less flow area is arranged.Like this, in whole duct, can produce cooling action relatively uniformly.

Description

The improvement structure of combustion gas turbine moving vane coolant channel
The present invention relates to the moving vane of combustion gas turbine, relate in particular to its cooling system.
The turbine blade that rotates is used air cooling usually, and air flows radially outward through many apertures that leaf pushes up through the blade root of associating.In the technology formerly, the aperture of common whole blade and root equates that perhaps blade partly is the first equal diameter, and second diameter of root for equating.The duct of root is bigger usually, do not need to prevent from the pressure loss of emphasis cooled region.The diameter in blade portion cooling agent duct must be very little, makes this place's coolant flow speed height and obtain needed heat-transfer coefficient.
No matter consider still to consider that from cooling the key Design district of blade is the stage casing district of blade, adjusts the number in aperture, hole and the flow of freezing mixture according to design idea usually from stress.When freezing mixture was outwards mobile along blade, it had been accepted the heat of hot blade passage gas and has significantly been heated, and therefore, blade stage casing district coolant temperature is higher than near the coolant temperature the blade rim significantly.
Near the blade rim coolant temperature is low will certainly to be cooled to be lower than the desired temperature of stress design to blade.The heat that near the blade rim cold excessively means the freezing mixture actual absorption than it should absorb from the wheel rim district many, stage casing district coolant temperature has just increased as a result, make coolant flow with (or) metal temperature is than height when cold not occurring.
Because the reduction of centrifugal stress more is far more than the heating of vane tip freezing mixture, so the leaf top also can occur cold.It is that pressure loss level is higher than the level that may reach when producing the coolant flow that reduced unit's flow area in cold-zone with cooling off the heat transfer level that requirement is complementary that wheel rim and leaf pushed up a cold important consequence.The reduction of crossing the cold-zone pressure loss makes that coolant flow may increase on the stage casing district unit flow area, can increase cooling thereby infeed under the pressure certain blade root.Like this, in design, just can select higher turbine inlet temperature for use, perhaps, under certain turbine inlet temperature, can select lower cooling flow for use.
And, in the technology of casting moving vane, usually cast out radially freezing mixture duct in the blade with core.The chance that the structure weakness of core throws away core fracture and blade is more than what expect, thereby has increased the unit manufacture cost of blade.
Therefore, basic purpose of the present invention provides a kind of new blade structure, and it has improved coolant channel, makes cooling more effective, and operating turbine is better, and it is higher to make efficient.
Given this purpose, the invention belongs to a moving vane of combustion gas turbine, the blade part that it has a root and extends from root, above-mentioned blade has many freezing mixture ducts on leaf top that extend to from blade root along the leaf height for the ANALYSIS OF COOLANT FLOW that enters from blade root, above-mentioned each freezing mixture duct has an inboard rim section to reach from the outward extending second portion of inboard wheel rim, it is characterized in that: the second portion in above-mentioned rim section and duct has corresponding flow area, the coolant flow that makes unit flow area in the rim section duct low than in the stage casing district duct second portion, thus the cold excessively of blade rim district prevented.
Also most preferred embodiment of the present invention is described by way of example below with reference to accompanying drawing, wherein:
Fig. 1 illustrates the perspective view of a combustion gas turbine blade, the coolant channel of blade principle according to the invention;
Fig. 2 illustrates the part of blade section, and the freezing mixture duct structure on the blade of presentation graphs 1 is the taper duct;
Fig. 3 and Fig. 4 represent other several embodiments of the taper freezing mixture duct structure that obtains according to the present invention.
The combustion gas turbine blade 10 that illustrates on Fig. 1 has blade root 12 and blade 14, and it arranges principle according to the invention.Freezing mixture duct or passage 16 extend radially outwardly along blade height.Air stream flows into from the freezing mixture compensate opening of blade root, outwards flows air stream cooling blade 14 and be exposed to surface in the blade path gas of heat along freezing mixture duct 16.
The stage casing district 18 of blade is the key Design district, and duct number, aperture and coolant flow at first will satisfy the cooling needs in this district.And the structure of total blade cooling system all will meet the requirement of this key area in design.
Coolant temperature low than stage casing district 18 in the duct 16 in blade rim district 20, this is because freezing mixture will be heated when blade outwards flows.In the design formerly, the freezing mixture of lower temperature can be crossed cold blade rim 20, the coolant temperature in stage casing district 18 is higher than cross the temperature when not cold.Coolant flow and (or) metal temperature all is higher than with eliminating the cold degree that can reach when reducing coolant temperature.
Can reduce or eliminate the cold excessively of blade rim district significantly with freezing mixture provided by the present invention duct structure, allow coolant flow and metal temperature all lower, improve the efficient of cooling effectiveness and turbine.
Referring to Fig. 2, freezing mixture duct 16 comprises rim section 22 and external lateral portion 24, and external lateral portion 24 passes blade stage casing district from rim section 22 and extends to leaf top 26.
To tapered with external lateral portion 24 joints, interior big outer little, the aperture of external lateral portion 24 is then constant substantially along its length from entry end for the rim section 22 in freezing mixture duct 16.
Because the duct is taper, so flow area big than stage casing district 18 in duct in the wheel rim district 20.The coolant flow that reduces unit flow area in the wheel rim district 20 has just reduced this district's heat-transfer coefficient and amount of cooling water.The design parameter of turbine and blade is determining the quantity and the length of taper, to reduce and basic the cold excessively of wheel rim district 20 of eliminating.
The freezing mixture duct of taper also reduces the pressure loss in freezing mixture duct.And, can obtain higher flows on the stage casing district unit flow area, regularly metal temperature is also low to infeed pressure one at the blade root place.Therefore, use the span formula duct cooling technology can be with higher turbine inlet temperature.Under certain turbine inlet temperature,, can utilize this point to reduce the cooling flow because this design segments has higher unit flow area flow.
Embodiment illustrated in fig. 3ly be used for reducing or eliminate the cold excessively of blade rim 20 and leaf top 21.Freezing mixture duct 16a comprises the wheel rim district 22a of taper, isometrical stage casing district 24a and along the outside tapered leaf top part 25a of ANALYSIS OF COOLANT FLOW direction.
When effectively flow area increased among the duct 25a of place, leaf top, coolant flow speed just reduced, and conducts heat also to reduce.Therefore, suitably select the taper quantity of duct 25a and length just can significantly reduce or eliminate cold, thisly coldly excessively reduce and cause owing to leaf pushes up 21 place's centrifugal stresses.Its primary income is the pressure loss that has further reduced freezing mixture, and the permission coolant flow is lower or the cooling in stage casing district 18 is increased.
Can duct 16a be made of two taper 22a that put upside down and 25a head rest head ground from the isometrical stage casing district 24a of duct 16a cancellation.At this moment, can and conduct heat according to stress and determine the tie point position of duct 22a and 25a.
For the pressure loss being reduced to minimum, the taper duct is preferably continuous taper, no matter is to adopt the nonlinear form that adopts in linear (as shown in the figure) or the design.In addition, the freezing mixture duct is preferably circular cross-section, but the present invention also can be used for noncircular cross section.
An alternative embodiment of the invention is shown on Fig. 4.Wherein, freezing mixture duct 16b is step-like, provides different cross sections along its length, cold excessively with remarkable reduction vane region.
In this example, freezing mixture duct 16b comprises wheel rim district 22b, stage casing district 24b and Ye Ding district 25b.Wheel rim district 22b is by the medial segment 22b in first aperture 1And the segmentum laterale 22b of smaller aperture due 2Form.Stage casing district 24b is near wheel rim segmentum laterale 22b 2Form, 22b is compared in its aperture 2Littler.Ye Ding district 25b is by 25b 1And 25b 2Form, its aperture is strengthened successively.It is fixed that each segment length of freezing mixture duct 16b and aperture are come according to the requirement of stress and heat transfer, and cold excessively with remarkable reduction vane region improves the cooling effectiveness of blade and the efficient of turbine.
Another advantage of using taper freezing mixture duct is to allow to make blade with casting technique, uses core to cast out the duct of freezing mixture.Therefore, can cast out freezing mixture duct 16 and 16a with the taper core, compare with the core that in blade, forms isometrical freezing mixture duct of common usefulness, in the wheel rim district and the feature of the bigger taper core of Ye Ding district diameter be the intensity height, because core intensity height, therefore just reduced that core breaks and blade gets rid of and takes off, reduced manufacture cost.
In addition, the taper core that intensity is high can allow to make less aperture, stage casing district, freezing mixture duct.In view of the above, the improvement of blade cooling and the reduction of blade coolant flow also can realize.

Claims (6)

1, moving vane of combustion gas turbine has a blade root to reach from the extended blade part of blade root, above-mentioned blade has many freezing mixture ducts that extend to the leaf top from blade root along the leaf height, for the ANALYSIS OF COOLANT FLOW that enters from blade root, there are an inboard rim section and outward extending thus second portion in above-mentioned each freezing mixture duct, it is characterized in that: the second portion in above-mentioned rim section and duct has corresponding flow area, make the flow of freezing mixture on rim section unit's flow area be lower than the flow of intermediate section duct second portion unit area, thereby prevent the cold excessively of blade rim district.
2, blade as claimed in claim 1 is characterized in that: the duct of above-mentioned rim section is that the entry end of first flow area becomes continuous taper to the outlet end that has dwindled flow area (being positioned at rim section and above-mentioned second portion joint) from the cross section.
3, blade as claimed in claim 2 is characterized in that: the flow area of whole above-mentioned duct second portion is equal substantially.
4, blade as claimed in claim 3 is characterized in that: the flow area of above-mentioned duct second portion equals the flow area of above-mentioned rim section duct outlet end.
5, blade as claimed in claim 2, it is characterized in that: above-mentioned duct second portion is formed by two sections, the flow area of first section total length is equal substantially, and equal the flow area of above-mentioned rim section duct outlet end, second section stretches out along the leaf height by first section, its entry end flow area equals first section flow area, becomes continuous taper from its entry end to the outlet end that has enlarged flow area (at Ye Dingchu).
6, blade as claimed in claim 1, it is characterized in that: above-mentioned duct second portion comprises intermediate section and rim section, whole intermediate section flow area is equal substantially, and rim section comprises entrance and outlet section, first flow area of the full section of entrance is equal substantially, outlet section links to each other with intermediate section, and the flow area of its full section is also equal substantially, but area less than above-mentioned first flow area greater than the flow area of above-mentioned intermediate section; Above-mentioned duct second portion also comprises duct, one first top section, this section links to each other with above-mentioned intermediate section and duct, second top section respectively, the latter has an outlet at Ye Dingchu, the flow area of above-mentioned first top section is greater than the flow area of intermediate section, and the flow area of above-mentioned second top section that outlet arranged is greater than the flow area of above-mentioned first top section.
CN 86104500 1985-07-03 1986-07-01 The improvement structure of combustion gas turbine moving vane coolant channel Pending CN86104500A (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US75165785A 1985-07-03 1985-07-03
US751,657 1985-07-03

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CN86104500A true CN86104500A (en) 1987-02-04

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CN 86104500 Pending CN86104500A (en) 1985-07-03 1986-07-01 The improvement structure of combustion gas turbine moving vane coolant channel

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EP (1) EP0207799A3 (en)
JP (1) JPS6210402A (en)
CN (1) CN86104500A (en)
CA (1) CA1262868A (en)
IE (1) IE861475L (en)

Cited By (11)

* Cited by examiner, † Cited by third party
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CN1083051C (en) * 1994-08-26 2002-04-17 亚瑞亚勃朗勃威力有限公司 Wall cooled by reflecting flow
CN1318735C (en) * 2005-12-26 2007-05-30 北京航空航天大学 Pulsing impact cooling blade for gas turbine engine
CN100346059C (en) * 2002-11-20 2007-10-31 三菱重工业株式会社 Turbine blade and gas turbine
CN101749053A (en) * 2008-12-08 2010-06-23 通用电气公司 Hollow passages
CN102213109A (en) * 2010-04-12 2011-10-12 通用电气公司 Turbine bucket having a radial cooling hole
CN102562309A (en) * 2010-12-21 2012-07-11 株式会社东芝 Transition piece and gas turbine
CN102741506A (en) * 2010-03-03 2012-10-17 三菱重工业株式会社 Rotor blade for gas turbine, method for manufacturing same, and gas turbine using rotor blade
CN102953766A (en) * 2011-08-24 2013-03-06 通用电气公司 Axially cooled airfoil
CN103038453A (en) * 2010-06-11 2013-04-10 西门子能量股份有限公司 Component wall having diffusion sections for cooling in a turbine engine
CN108868897A (en) * 2017-05-11 2018-11-23 通用电气公司 The insertion piece of turbine engine airfoil part
CN110159357A (en) * 2019-06-04 2019-08-23 北京航空航天大学 A kind of aero engine turbine blades reducing and expansion type air supply channel promoting passive security

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US4893987A (en) * 1987-12-08 1990-01-16 General Electric Company Diffusion-cooled blade tip cap
US5413463A (en) * 1991-12-30 1995-05-09 General Electric Company Turbulated cooling passages in gas turbine buckets
US6539627B2 (en) 2000-01-19 2003-04-01 General Electric Company Method of making turbulated cooling holes
GB0229908D0 (en) * 2002-12-21 2003-01-29 Macdonald John Turbine blade
US6997679B2 (en) * 2003-12-12 2006-02-14 General Electric Company Airfoil cooling holes
US7413406B2 (en) 2006-02-15 2008-08-19 United Technologies Corporation Turbine blade with radial cooling channels
US8511992B2 (en) * 2008-01-22 2013-08-20 United Technologies Corporation Minimization of fouling and fluid losses in turbine airfoils
US8157527B2 (en) * 2008-07-03 2012-04-17 United Technologies Corporation Airfoil with tapered radial cooling passage
JP2010053749A (en) * 2008-08-27 2010-03-11 Mitsubishi Heavy Ind Ltd Blade for turbine
US8511990B2 (en) * 2009-06-24 2013-08-20 General Electric Company Cooling hole exits for a turbine bucket tip shroud
EP2354453B1 (en) * 2010-02-02 2018-03-28 Siemens Aktiengesellschaft Turbine engine component for adaptive cooling
US8506251B2 (en) 2010-03-03 2013-08-13 Mitsubishi Heavy Industries, Ltd. Gas turbine blade, manufacturing method therefor, and gas turbine using turbine blade
CA3123208C (en) 2012-08-14 2023-10-03 Loop Energy Inc. Fuel cell flow channels and flow fields
WO2014026287A1 (en) 2012-08-14 2014-02-20 Powerdisc Development Corporation Ltd. Fuel cell components, stacks and modular fuel cell systems
US9644277B2 (en) 2012-08-14 2017-05-09 Loop Energy Inc. Reactant flow channels for electrolyzer applications
US20140161625A1 (en) * 2012-12-11 2014-06-12 General Electric Company Turbine component having cooling passages with varying diameter
US10408079B2 (en) 2015-02-18 2019-09-10 Siemens Aktiengesellschaft Forming cooling passages in thermal barrier coated, combustion turbine superalloy components
WO2017121689A1 (en) 2016-01-15 2017-07-20 Siemens Aktiengesellschaft Gas turbine aerofoil
CN109075358B (en) 2016-03-22 2021-10-19 环能源公司 Fuel cell flow field design for thermal management
JP7234006B2 (en) * 2019-03-29 2023-03-07 三菱重工業株式会社 High temperature parts and method for manufacturing high temperature parts
KR102630916B1 (en) * 2019-06-05 2024-01-29 미츠비시 파워 가부시키가이샤 Turbine blades and turbine blade manufacturing method and gas turbine
JP6637630B1 (en) * 2019-06-05 2020-01-29 三菱日立パワーシステムズ株式会社 Turbine blade, method of manufacturing turbine blade, and gas turbine

Family Cites Families (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1018747A (en) * 1964-11-13 1966-02-02 Rolls Royce Aerofoil shaped blade for fluid flow machines
JPS4825103B1 (en) * 1967-06-05 1973-07-26
JPS5520041A (en) * 1978-07-29 1980-02-13 Noto Denshi Kogyo Kk Piezoelectric device
JPS5669423A (en) * 1979-11-09 1981-06-10 Hitachi Ltd Air-cooled blade of gas turbine

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN1083051C (en) * 1994-08-26 2002-04-17 亚瑞亚勃朗勃威力有限公司 Wall cooled by reflecting flow
CN100346059C (en) * 2002-11-20 2007-10-31 三菱重工业株式会社 Turbine blade and gas turbine
CN1318735C (en) * 2005-12-26 2007-05-30 北京航空航天大学 Pulsing impact cooling blade for gas turbine engine
CN101749053A (en) * 2008-12-08 2010-06-23 通用电气公司 Hollow passages
CN101749053B (en) * 2008-12-08 2015-09-02 通用电气公司 hollow channel
CN102741506A (en) * 2010-03-03 2012-10-17 三菱重工业株式会社 Rotor blade for gas turbine, method for manufacturing same, and gas turbine using rotor blade
CN102741506B (en) * 2010-03-03 2015-07-01 三菱日立电力系统株式会社 Rotor blade for gas turbine, method for manufacturing same, and gas turbine using rotor blade
CN102213109A (en) * 2010-04-12 2011-10-12 通用电气公司 Turbine bucket having a radial cooling hole
CN103038453A (en) * 2010-06-11 2013-04-10 西门子能量股份有限公司 Component wall having diffusion sections for cooling in a turbine engine
CN102562309A (en) * 2010-12-21 2012-07-11 株式会社东芝 Transition piece and gas turbine
US9200526B2 (en) 2010-12-21 2015-12-01 Kabushiki Kaisha Toshiba Transition piece between combustor liner and gas turbine
CN102953766A (en) * 2011-08-24 2013-03-06 通用电气公司 Axially cooled airfoil
CN108868897A (en) * 2017-05-11 2018-11-23 通用电气公司 The insertion piece of turbine engine airfoil part
CN110159357A (en) * 2019-06-04 2019-08-23 北京航空航天大学 A kind of aero engine turbine blades reducing and expansion type air supply channel promoting passive security

Also Published As

Publication number Publication date
IE861475L (en) 1987-01-03
EP0207799A2 (en) 1987-01-07
CA1262868A (en) 1989-11-14
JPS6210402A (en) 1987-01-19
EP0207799A3 (en) 1988-09-14

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