EP0182588B1 - Multi-chamber airfoil cooling insert for turbine vane - Google Patents

Multi-chamber airfoil cooling insert for turbine vane Download PDF

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Publication number
EP0182588B1
EP0182588B1 EP85308230A EP85308230A EP0182588B1 EP 0182588 B1 EP0182588 B1 EP 0182588B1 EP 85308230 A EP85308230 A EP 85308230A EP 85308230 A EP85308230 A EP 85308230A EP 0182588 B1 EP0182588 B1 EP 0182588B1
Authority
EP
European Patent Office
Prior art keywords
chambers
rearward
turbine
insert
ports
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.)
Expired
Application number
EP85308230A
Other languages
German (de)
French (fr)
Other versions
EP0182588A1 (en
Inventor
Thomas M. Szewczuk
William Edward North
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
CBS Corp
Original Assignee
Westinghouse Electric Corp
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 Westinghouse Electric Corp filed Critical Westinghouse Electric Corp
Publication of EP0182588A1 publication Critical patent/EP0182588A1/en
Application granted granted Critical
Publication of EP0182588B1 publication Critical patent/EP0182588B1/en
Expired legal-status Critical Current

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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
    • 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

Definitions

  • This invention relates to a combustion turbine and in particular to an airfoil-shaped, hollow turbine vane having a leading edge wall.
  • Such turbine airfoil vanes provided with an insert, with the arrangement as a whole providing for air cooling of the vanes.
  • the vane structure is of a character and in a stage calling for a low or a moderate level of cooling, which level of cooling can be carried out by the use of impingement jets directed against the interior walls of the vane. Even with those vanes which do not require a high level of cooling, the degree of cooling required at different locations on the vane may differ, with the leading edge region of the vane typically having a relatively higher heat load while downstream and toward the trailing edge of the vane the heat load may be significantly lower.
  • French Patent No. 2,457,965 discloses turbine blades with inserts defining a front chamber to which a cooling gas is supplied and escapes partially through passages in the leading edge of the blade and partially into a space between the insert in the front chamber and the blade wall. The cooling gas then enters the interior of an insert in the rear chamber from where the cooling gas is discharged through openings therein into a space formed between the insert and the blade walls before it is discharged through openings in the trailing edge of the blade.
  • U.S. Patent 2,873,944 also discloses turbine blades or vanes with cooling air guide inserts defining different chambers provided with openings for even distribution of the cooling air to the inner surfaces of the blade or vane walls.
  • the present invention resides in a combustion turbine with airfoil-shaped, hollow, turbine vanes having leading edge walls, trailing edge portions with exit air slots therein, and pressure and suction sidewalls defining internal cavities in communication with said exit air slots, with hollow inserts of substantially complementary airfoil shape in cross section located in said cavities and extending in a chordwise direction for substantially the entire extent of said cavities in spaced relationship from the walls thereof, characterized in that said inserts are single unitary structures with a plurality of radially extending partitions in each insert dividing the interior thereof into a forward chamber in the leading edge portion of said vane, and at least two separate, successively rearward chambers in communication with each other, a plurality of impingement ports in the insert walls of all of the said chambers, one radial end portion of said chambers being in communication with a source of cooling air, and that for throttling the flow into said rearward chambers are provided such that said forward chamber is at a relatively higher pressure than said rearward chambers and the imp
  • the single, unitary, air-cooling, hollow insert 22 has an airfoil shape in cross section which is complementary to the vane airfoil shape, and extends in a chordwise direction for substantially the entire extent of the vane cavity. While the insert does have the overall shape of an airfoil, it may be seen in Figure 1 that at the leading edge portion 24 the insert is bulged somewhat, a similar bulged arrangement being provided at the trailing edge portion 26.
  • the intermediate portion 28 has walls which are basically uniformly spaced from the vane wails throughout the intermediate extent between the front and rear bulges.
  • the unitary insert 22 has its interior divided into a forward chamber 30 and successively rearward chambers 32, 34, and 36, by the radially-extending partition means 38, 40, and 42, which also perform a structural tying function.
  • the radially inner ends of all of the chambers are closed while the radially outer ends of the chambers are in communication with a source of cooling air.
  • the radially outer end 44 is completely open so that cooling air flows directly into the forward chamber 30 as indicated by the arrow 46 in Figure 2.
  • the flow into these chambers is throttled by means of a radial extension 48 of the insert comprising opposite walls 50 capped by plate 52 which prevents the direct admission of the cooling air into the rearward chambers in the fashion in which the forward chambers receives its air, the cooling air being throttled into the rearward chambers by the provision of the holes 54 in the walls 50.
  • the throttling results in the rearward chambers being at a lower pressure than the forward chamber 30.
  • all of the chambers are provided with impingement ports in their sidewalls.
  • Those ports provided in the forward chamber sidewalls are identified by the numeral 56 as best seen in Figre 1.
  • the impingement ports in the convex sidewall of the insert for the rearward chambers are designated 58 while those in the concave wall are designated 60.
  • All of the impingement ports are in rows which extend substantially radially.
  • the rows of ports 56 of the forward chamber are more widely spaced from each other than the rows of ports from-the rearward chambers on the convex side, and most of the concave side with the exception of the spacing of the rows of ports of the concave side of the first low-pressure chamber 32.
  • the three rearward chambers are open to each other through the provision of a series of ports 62 in both of the partitions or ribs 40 and 42.
  • the rearward chambers are also in open communication with each other at the radially outer portion of the chambers by virtue of the partitions 40 and 42 stopping short of the space 64 at the radially outer ends of the chambers.
  • the insert has dimples 66 embossed outwardly in its leading edge portion and similar dimples 68 in its trailing edge portion to properly space the insert walls from the vane walls.
  • the forward chamber 30 is maintained at a higher pressure than the rearward chambers 32, 34, and 36, so that the cooling jets issuing from the forward chamber are projected at a higher velocity than those exiting through the ports of the rearward chambers so that the higher velocity jets are projected at the higher heat load leading edge and forward convex surface areas of the vane, while the jets issuing from the lower pressure rearward chambers are projected at a lower velocity for cooling the relativeiy lower heat load regions of the airfoil vane.
  • the relatively closer spaced rows of ports throughout the midchord region is to obtain more uniform cooling than would be obtained with widely-spaced high velocity jets.
  • Typical pressures at which the chambers can be maintained may be in order of, for example, 160 psig (1102 E+03Pa) for the forward chamber, 155 psig (1068 E+03Pa) for the rearward chambers, with the pressures in the spaces between the insert and the opposing vane walls being 150 psig (1033 E+03Pa).

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

Description

  • This invention relates to a combustion turbine and in particular to an airfoil-shaped, hollow turbine vane having a leading edge wall. Such turbine airfoil vanes provided with an insert, with the arrangement as a whole providing for air cooling of the vanes.
  • It is known that different stages of the stator vanes require different levels of cooling. The vane structure is of a character and in a stage calling for a low or a moderate level of cooling, which level of cooling can be carried out by the use of impingement jets directed against the interior walls of the vane. Even with those vanes which do not require a high level of cooling, the degree of cooling required at different locations on the vane may differ, with the leading edge region of the vane typically having a relatively higher heat load while downstream and toward the trailing edge of the vane the heat load may be significantly lower.
  • In order to achieve such different cooling rates the gas turbine blades as shown for example in French Patent 2,290,573 have inserts with air cooling jet openings adjacent the leading end of the blades where cooling is mostly needed.
  • French Patent No. 2,457,965 discloses turbine blades with inserts defining a front chamber to which a cooling gas is supplied and escapes partially through passages in the leading edge of the blade and partially into a space between the insert in the front chamber and the blade wall. The cooling gas then enters the interior of an insert in the rear chamber from where the cooling gas is discharged through openings therein into a space formed between the insert and the blade walls before it is discharged through openings in the trailing edge of the blade. U.S. Patent 2,873,944 also discloses turbine blades or vanes with cooling air guide inserts defining different chambers provided with openings for even distribution of the cooling air to the inner surfaces of the blade or vane walls.
  • It is the principal object of the present invention to provide a vane and insert structure in which a vane having a single internal cavity is provided with a single, unitary, hollow insert provided with a chamber arrangement and jet impingement ports all tailored to relate the impingement cooling of the walls to the external heat load.
  • With this object in view the present invention resides in a combustion turbine with airfoil-shaped, hollow, turbine vanes having leading edge walls, trailing edge portions with exit air slots therein, and pressure and suction sidewalls defining internal cavities in communication with said exit air slots, with hollow inserts of substantially complementary airfoil shape in cross section located in said cavities and extending in a chordwise direction for substantially the entire extent of said cavities in spaced relationship from the walls thereof, characterized in that said inserts are single unitary structures with a plurality of radially extending partitions in each insert dividing the interior thereof into a forward chamber in the leading edge portion of said vane, and at least two separate, successively rearward chambers in communication with each other, a plurality of impingement ports in the insert walls of all of the said chambers, one radial end portion of said chambers being in communication with a source of cooling air, and that for throttling the flow into said rearward chambers are provided such that said forward chamber is at a relatively higher pressure than said rearward chambers and the impingement jets through said ports of said forward chamber against said interior vane walls of said leading edge portion are at a significantly higher velocity than the impingement jets exiting the ports of said rearward chambers.
  • The invention will now be described, by way of example, with reference to the accompanying drawings in which:
    • Figure 1 is a chordwise sectional view through the vane and insert as would appear from a section taken along the line I-I of Figure 2; and
    • Figure 2 is a view partly in elevation and partly in section of the vane and insert, and corresponding to a view taken along the line II-II of Figure I.
    • Figure 1 shows a hollow vane having a single internal cavity defined by the leading edge section, a concave sidewall 14, a convex sidewall 16, the downstream portions of these opposite sidewalls defining a trailing edge portion is provided with a slot 20 therein. The general direction of the hot gas past the vane is as indicated by the dash line arrow in Figure 1.
  • The single, unitary, air-cooling, hollow insert 22 has an airfoil shape in cross section which is complementary to the vane airfoil shape, and extends in a chordwise direction for substantially the entire extent of the vane cavity. While the insert does have the overall shape of an airfoil, it may be seen in Figure 1 that at the leading edge portion 24 the insert is bulged somewhat, a similar bulged arrangement being provided at the trailing edge portion 26. The intermediate portion 28 has walls which are basically uniformly spaced from the vane wails throughout the intermediate extent between the front and rear bulges.
  • The unitary insert 22 has its interior divided into a forward chamber 30 and successively rearward chambers 32, 34, and 36, by the radially-extending partition means 38, 40, and 42, which also perform a structural tying function.
  • The radially inner ends of all of the chambers are closed while the radially outer ends of the chambers are in communication with a source of cooling air. As may be best understood from Figure 2, the radially outer end 44 is completely open so that cooling air flows directly into the forward chamber 30 as indicated by the arrow 46 in Figure 2. While the rearward chambers 32, 34, and 36 are also in communication with the source of cooling air, the flow into these chambers is throttled by means of a radial extension 48 of the insert comprising opposite walls 50 capped by plate 52 which prevents the direct admission of the cooling air into the rearward chambers in the fashion in which the forward chambers receives its air, the cooling air being throttled into the rearward chambers by the provision of the holes 54 in the walls 50. The throttling results in the rearward chambers being at a lower pressure than the forward chamber 30.
  • Referring to both figures, all of the chambers are provided with impingement ports in their sidewalls. Those ports provided in the forward chamber sidewalls are identified by the numeral 56 as best seen in Figre 1. The impingement ports in the convex sidewall of the insert for the rearward chambers are designated 58 while those in the concave wall are designated 60. All of the impingement ports are in rows which extend substantially radially. The rows of ports 56 of the forward chamber are more widely spaced from each other than the rows of ports from-the rearward chambers on the convex side, and most of the concave side with the exception of the spacing of the rows of ports of the concave side of the first low-pressure chamber 32. The three rearward chambers are open to each other through the provision of a series of ports 62 in both of the partitions or ribs 40 and 42. The rearward chambers are also in open communication with each other at the radially outer portion of the chambers by virtue of the partitions 40 and 42 stopping short of the space 64 at the radially outer ends of the chambers.
  • The insert has dimples 66 embossed outwardly in its leading edge portion and similar dimples 68 in its trailing edge portion to properly space the insert walls from the vane walls.
  • With the arrangement as shown and described, the forward chamber 30 is maintained at a higher pressure than the rearward chambers 32, 34, and 36, so that the cooling jets issuing from the forward chamber are projected at a higher velocity than those exiting through the ports of the rearward chambers so that the higher velocity jets are projected at the higher heat load leading edge and forward convex surface areas of the vane, while the jets issuing from the lower pressure rearward chambers are projected at a lower velocity for cooling the relativeiy lower heat load regions of the airfoil vane. The relatively closer spaced rows of ports throughout the midchord region is to obtain more uniform cooling than would be obtained with widely-spaced high velocity jets.
  • Typical pressures at which the chambers can be maintained may be in order of, for example, 160 psig (1102 E+03Pa) for the forward chamber, 155 psig (1068 E+03Pa) for the rearward chambers, with the pressures in the spaces between the insert and the opposing vane walls being 150 psig (1033 E+03Pa).

Claims (6)

1. A combustion turbine with airfoil-shaped, hollow, turbine vanes (12) having leading edge walls, trailing edge portions (18) with exit air slots (20) therein, and pressure and suction sidewalls (14,16) defining internal cavities in communication with said exit air slots, with hollow inserts (22) of substantially complementary airfoil shape in cross section located in said cavities and extending in a chordwise direction for substantially the entire extent of said cavities in spaced relationship from the walls thereof, characterized in that said inserts (22) are single unitary structures with a plurality of radially extending partitions (38,40,42) in each insert (22) dividing the interior thereof into a forward chamber (30) in the leading edge portion of said vane (12), and at least two separate, successively rearward chambers (32,34,36) in communication with each other, a plurality of impingement ports (56,58,60) in the insert walls of all of said chambers (30,36), one radial end portion of said chambers (30,36) being in communication with a source of cooling air, and that means (52,54) for throttling the flow into said rearward chambers (32-36) are provided such that said forward chamber (30) is at a relatively higher pressure than said rearward chambers (32,34,36) and the impingement jets through said ports (56) of said forward chamber (30) against said interior vane walls (14,16) of said leading edge portion are at a significantly higher velocity than the impingement jets exiting the ports (58,60) of said rearward chambers (32,34,36).
2. A turbine as claimed in claim 1, characterized in that said impingement ports (56) in said forward chamber (30) are arranged at a greater distance from one another than the impingement ports (58,60) in said rearward chambers (32,34,36).
3. A turbine as claimed in claim 1 or 2, characterized in that said throttling means includes a radially outwardly extending portion (48) of the insert (22) at the radially outer ends of said rearward chambers (32,34,36), and a plurality of throttling holes (54) in said portion (48) of said insert (22).
4. A turbine as claimed in any one of claims 1 to 3, characterized in that said rearward chambers comprise at least three chambers (32,34,36).
5. A turbine as claimed in any one of claims 1 to 4, characterized in that said partitions (38,40,42) comprise rigidly extending ribs, the first rib (38) separating said forward chamber (30) from the first successively rear chamber (32) being imperforate, and successive rearward second ribs (40,42) having openings (62) therein.
6. A combustion turbine as claimed in claim 5, characterized in that said second ribs (40,42) extend radially outwardly less than said first rib (38) so that said rearward chambers (32,34,36) are in open communication with each other in their radially outer portions.
EP85308230A 1984-11-15 1985-11-13 Multi-chamber airfoil cooling insert for turbine vane Expired EP0182588B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US67184684A 1984-11-15 1984-11-15
US671846 1984-11-15

Publications (2)

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EP0182588A1 EP0182588A1 (en) 1986-05-28
EP0182588B1 true EP0182588B1 (en) 1988-09-28

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Family Applications (1)

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EP85308230A Expired EP0182588B1 (en) 1984-11-15 1985-11-13 Multi-chamber airfoil cooling insert for turbine vane

Country Status (9)

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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)

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US9850763B2 (en) 2015-07-29 2017-12-26 General Electric Company Article, airfoil component and method for forming article
US10260363B2 (en) 2016-12-08 2019-04-16 General Electric Company Additive manufactured seal for insert compartmentalization
US10655477B2 (en) 2016-07-26 2020-05-19 General Electric Company Turbine components and method for forming turbine components
US10697309B2 (en) 2018-04-25 2020-06-30 Raytheon Technologies Corporation Platform cover plates for gas turbine engine components

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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
US7008185B2 (en) * 2003-02-27 2006-03-07 General Electric Company Gas turbine engine turbine nozzle bifurcated impingement baffle
US7871246B2 (en) * 2007-02-15 2011-01-18 Siemens Energy, Inc. Airfoil for a gas turbine
CN104088673B (en) * 2008-11-07 2016-03-09 三菱日立电力系统株式会社 turbine blade
CN101825115B (en) * 2010-03-31 2011-09-28 北京航空航天大学 Blade with built-in bed frame-type pneumatic damping device
US20130104567A1 (en) * 2011-10-31 2013-05-02 Douglas Gerard Konitzer Method and apparatus for cooling gas turbine rotor blades
US9004866B2 (en) * 2011-12-06 2015-04-14 Siemens Aktiengesellschaft Turbine blade incorporating trailing edge cooling design
US9506351B2 (en) 2012-04-27 2016-11-29 General Electric Company Durable turbine vane
EP2706195A1 (en) * 2012-09-05 2014-03-12 Siemens Aktiengesellschaft Impingement tube for gas turbine vane with a partition wall
WO2016036366A1 (en) 2014-09-04 2016-03-10 Siemens Aktiengesellschaft Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil
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
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
US10400608B2 (en) * 2016-11-23 2019-09-03 General Electric Company Cooling structure for a turbine component
US10494948B2 (en) * 2017-05-09 2019-12-03 General Electric Company Impingement insert
US10480347B2 (en) 2018-01-18 2019-11-19 United Technologies Corporation Divided baffle for components of gas turbine engines

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Cited By (4)

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Publication number Priority date Publication date Assignee Title
US9850763B2 (en) 2015-07-29 2017-12-26 General Electric Company Article, airfoil component and method for forming 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
US10697309B2 (en) 2018-04-25 2020-06-30 Raytheon Technologies Corporation Platform cover plates for gas turbine engine components

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
CN85108282A (en) 1986-08-27
KR860004224A (en) 1986-06-18
EP0182588A1 (en) 1986-05-28

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