US3806274A - Gas turbine engine blades - Google Patents

Gas turbine engine blades Download PDF

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Publication number
US3806274A
US3806274A US00282778A US28277872A US3806274A US 3806274 A US3806274 A US 3806274A US 00282778 A US00282778 A US 00282778A US 28277872 A US28277872 A US 28277872A US 3806274 A US3806274 A US 3806274A
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United States
Prior art keywords
blade
insert
walls
flow passages
fins
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Expired - Lifetime
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US00282778A
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A Moore
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Rolls Royce PLC
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Rolls Royce 1971 Ltd
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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
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall

Definitions

  • a gas turbme blade has a hollow Interim 1 9 which [58 Field 6: Search 416/96-97, is divided form flow Passages cooling medium- 416/92
  • the flow passages are bounded by the sides of a sheetlike insert the two blade walls, and fins between the [56] References Cited insert and the blade walls; they commence at one end UNITED STATES PATENTS of the blade and extend in a spiral-like path around the opposite sides of the insert.
  • the present invention aims to provide a rotor or stator blade construction which enables improved cooling to be obtained.
  • An object of the invention is to provide an improved a gas-turbine blade with a hollow interior space which is divided to form flow passages bounded by the sides of a sheet-like insert, the two blade walls and fins between the insert and the blade walls, the flow passages commencing at one end of the blade and extending in a spiral-like path around and along the insert.
  • the invention is applied to rotor blades which are fed from the root end.
  • FIG. 1 shows a sectional elevation of a gas turbine rotor blade according to the invention, the section blade is hollow, and the space within it is effectively di-' vided into two by means of a sheet-metal insert 14 which extends substantially the full radial length of the blade being brazed to the blade at its root end but otherwise unattached to allow for differential expansion.
  • the arrangement of the insert 14 within the blade is clearly seen in FIG. 2.
  • the insert is located between a large number of pimples 15 (indicated in FIG. 1 by small crosses) and by a series of helical fins 16.
  • blade is conveniently made with the pimples and fins cast onto the interior surfaces of the blade walls l7, 18.
  • the insert 14 stops short of both the leading and trailing edges of they blade, thus leaving spaces around which air may pass in order to progress from,one side of the insert to the other.
  • air suppliedto the interior of the blade from the root end 13 can enter any one of a series of flow passages extending obliquely up one side of the insert passing round the leading or trailing edge of the insert, and then continuing obliquely up the other side of the insert.
  • An arrow traces one possible path. As it pro gresses radially outwards, a proportion of the air is allowed to escape through leading edge film cooling holes 20, and'through trailing edge slots 21. Since the volume of air flowing through the blade is thus diminished, the flow passages are designed to be of smaller cross section towards the outer shroud end 12 of the blade.
  • the velocity of the air flow may be varied by altering the angle at which the fins 16 arecast onto the interior surfacesof the blade.
  • holes 22 may be provided for the final escape of the cooling air in the angle formed by the blade proper wall 17 and the shroud 23.
  • the air thus released can be used to cool the shroudby causing it to spread over the shroud 23 along channels 24 formed in the radially inner surface thereof.
  • a blade constructed in the fashion described is relatively easy to produce, particularly since the insert is solid and therefore easy both to manufacture and to insert into the blade, while the blade itself can be conveniently cast, together with, fins and pimples, using only a single piece core.
  • the holes 20 and/or the holes 21 and/or the pimples 15 may be omit ted, the cooling medium may be liquid or a gas other than air, and the fins 16 and'pimples (if provided) may be other than integral with the blade walls.
  • At least some of the flow passages for cooling air are such that they individually serve either the leading edge of the blade or the trailing edge of the blade, but not both. This may be effected by suitable choice of the inclination of the fins 16. Such an arrangement may enable different supplies to be used for the leading and trailing edges of the blade to make allowance for the different ambient air pressures existing in those regions.
  • a gas turbine blade having opposed walls defining therebetween a hollow interior space, an insert within said space, said insert being formed of a solid sheet curved to the chamberof the blade, and fins carried by said blade and extending between said insert and said blade walls, said walls, said-insert and said fins defining in combination a plurality of flow passages for cooling fluid, each said flow passage commencing at one end of the blade and extending in a spiral-like manner around and along the insert.
  • a blade as claimed in claim 1 which includes a plurality of discrete pimples extending between the sides of the insert and the blade walls within the flow passages for providing location for the insert.
  • a blade as claimed in claim 2 wherein the fins and pimples are formed integrally with. the blade walls. 4. A blade as claimed in claim 1, wherein holes are formed in the region of the leading edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the flow passages.
  • a gas turbine rotor having a plurality of blades as claimed in claim 1, arranged for supply with cooling fluid at their root ends.
  • a gas turbine rotor according to claim 8 which includes a shroud connecting together the outer ends of 5 are terminated immediately adjacent said root end portion of said insert.
  • a gas turbine blade having opposed walls defining therebetween a hollow interior space, and an insert within said space, said insert being formed of a solid sheet curved to the chamber of the blade and connected to said walls by brazing at the root end of the blade, being otherwise unattached to allow for differential expansion, said walls being integrally formed with fins extending between said insert and said blade walls, said walls, said insert and said fins defining in combination a plurality of flow passages for cooling fluid to flow through the blade commencing at the root end of 'the blade and passing in a spiral-like manner around and along the insert.

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

Abstract

A gas turbine blade has a hollow interior space which is divided to form flow passages for cooling medium. The flow passages are bounded by the sides of a sheet-like insert the two blade walls, and fins between the insert and the blade walls; they commence at one end of the blade and extend in a spiral-like path around the opposite sides of the insert.

Description

United States Patent 1191 Moore 1 Apr. 23, 1974 [54] GAS TURBINE ENGINE BLADES 3,370,829 2/1968 Banthin et a1. 416 97 3,527,544 9/1970 Allen 416/97 UX [751 Invenm 3mm, Engla'ld 3,606,574 9/1971 Brands et al 416/97 x 1 3,628,880 12/1971 Smuland et a1 416/97 X [73] Asslgnee 53223 252; F 3,635,587 1/1972 Giesman et 111. 416/97 [22] Aug 22 1972 FOREIGN PATENTS OR APPLICATIONS [21] Appl No 282 778 853,534 10/1952 Germany 416/96 Primary Examiner Everette A. Powell, Jr. [30] Foreign Application Priority Data Attorney, Agent, or Firm-Stevens, Davis, Miller &
Aug. 25, 1971 Great Britain 39778/71 Mosher Aug, 25, 1971 Great Britain 39778/71 57 ABSTRACT [52] US. Cl. 416/97 1 51 1m. (:1. F01d 5/18 A gas turbme blade has a hollow Interim 1 9 which [58 Field 6: Search 416/96-97, is divided form flow Passages cooling medium- 416/92 The flow passages are bounded by the sides of a sheetlike insert the two blade walls, and fins between the [56] References Cited insert and the blade walls; they commence at one end UNITED STATES PATENTS of the blade and extend in a spiral-like path around the opposite sides of the insert. 2,888,243 5/1959 Pollock 416/92 1 3 3,057,597 10/ 1962 Meyer et a1... 416/96 12 Claims, 3 Drawing Figures 1 GAS TURBINE ENGINE BLADES This invention relates to the blading of gas turbine engines, and although not so limited, has particular reference to rotor blades for such engines.
As is known, it is advantageous to provide for the cooling of gas turbine rotor blades, and many different ways of doing this are known. In particular, it is known to provide a cooling air flow through the interior of the blade, and the cooling is the more effective, the greater the volume of air flow and the greater the velocity of the air as it passes through the blade.
In very small rotor blades, e.g., less than 1% inches long considerable difficulties are encountered in providing flow channels by means of which satisfactory cooling can be achieved. The present invention aims to provide a rotor or stator blade construction which enables improved cooling to be obtained.
An object of the invention is to provide an improved a gas-turbine blade with a hollow interior space which is divided to form flow passages bounded by the sides of a sheet-like insert, the two blade walls and fins between the insert and the blade walls, the flow passages commencing at one end of the blade and extending in a spiral-like path around and along the insert.
In a preferred form, the invention is applied to rotor blades which are fed from the root end.
The invention will now be described with reference to the accompanying drawing, in which:
FIG. 1 shows a sectional elevation of a gas turbine rotor blade according to the invention, the section blade is hollow, and the space within it is effectively di-' vided into two by means of a sheet-metal insert 14 which extends substantially the full radial length of the blade being brazed to the blade at its root end but otherwise unattached to allow for differential expansion. The arrangement of the insert 14 within the blade is clearly seen in FIG. 2.
It will be seen that the insert is located between a large number of pimples 15 (indicated in FIG. 1 by small crosses) and by a series of helical fins 16. The
blade is conveniently made with the pimples and fins cast onto the interior surfaces of the blade walls l7, 18.
The insert 14 stops short of both the leading and trailing edges of they blade, thus leaving spaces around which air may pass in order to progress from,one side of the insert to the other.
Referring new again to FIG. 1, it will be seen that air suppliedto the interior of the blade from the root end 13 can enter any one of a series of flow passages extending obliquely up one side of the insert passing round the leading or trailing edge of the insert, and then continuing obliquely up the other side of the insert. An arrow traces one possible path. As it pro gresses radially outwards, a proportion of the air is allowed to escape through leading edge film cooling holes 20, and'through trailing edge slots 21. Since the volume of air flowing through the blade is thus diminished, the flow passages are designed to be of smaller cross section towards the outer shroud end 12 of the blade. This can be achieved by arranging that the spacing of the fins 16 is reduced or,'mor'e conveniently, by reducing the width of the passages. This maintains the airflow velocity substantially constant. In addition, the velocity of the air flow may be varied by altering the angle at which the fins 16 arecast onto the interior surfacesof the blade.
In a preferred form of the invention, holes 22 (FIG! 3) may be provided for the final escape of the cooling air in the angle formed by the blade proper wall 17 and the shroud 23. The air thus released can be used to cool the shroudby causing it to spread over the shroud 23 along channels 24 formed in the radially inner surface thereof.
lt will be seen that a blade constructed in the fashion described is relatively easy to produce, particularly since the insert is solid and therefore easy both to manufacture and to insert into the blade, while the blade itself can be conveniently cast, together with, fins and pimples, using only a single piece core.
. If desired, in the described arrangement the holes 20 and/or the holes 21 and/or the pimples 15 may be omit ted, the cooling medium may be liquid or a gas other than air, and the fins 16 and'pimples (if provided) may be other than integral with the blade walls.
In a further variation of the described arrangement at least some of the flow passages for cooling air are such that they individually serve either the leading edge of the blade or the trailing edge of the blade, but not both. This may be effected by suitable choice of the inclination of the fins 16. Such an arrangement may enable different supplies to be used for the leading and trailing edges of the blade to make allowance for the different ambient air pressures existing in those regions.
. What is claimed is:
l. A gas turbine blade having opposed walls defining therebetween a hollow interior space, an insert within said space, said insert being formed of a solid sheet curved to the chamberof the blade, and fins carried by said blade and extending between said insert and said blade walls, said walls, said-insert and said fins defining in combination a plurality of flow passages for cooling fluid, each said flow passage commencing at one end of the blade and extending in a spiral-like manner around and along the insert.
2. A blade as claimed in claim 1, which includes a plurality of discrete pimples extending between the sides of the insert and the blade walls within the flow passages for providing location for the insert.
3. A blade as claimed in claim 2, wherein the fins and pimples are formed integrally with. the blade walls. 4. A blade as claimed in claim 1, wherein holes are formed in the region of the leading edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the flow passages.
5. A blade as claimed in claim 1, wherein holes are formed in the region of the trailing edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the flowpassages.
6. A blade as claimed in claim 1, wherein the insert has a root end portion at which it is secured within the blade by brazing, the insert being otherwise unattached to allow for differential expansion. I l
7. A blade as claimed in claim 1, wherein at least some of the flow passages are arranged for individually serving one edge only of the blade.
8. A gas turbine rotor having a plurality of blades as claimed in claim 1, arranged for supply with cooling fluid at their root ends.
9. A gas turbine rotor according to claim 8, which includes a shroud connecting together the outer ends of 5 are terminated immediately adjacent said root end portion of said insert.
12 A gas turbine blade having opposed walls defining therebetween a hollow interior space, and an insert within said space, said insert being formed of a solid sheet curved to the chamber of the blade and connected to said walls by brazing at the root end of the blade, being otherwise unattached to allow for differential expansion, said walls being integrally formed with fins extending between said insert and said blade walls, said walls, said insert and said fins defining in combination a plurality of flow passages for cooling fluid to flow through the blade commencing at the root end of 'the blade and passing in a spiral-like manner around and along the insert.

Claims (12)

1. A gas turbine blade having opposed walls defining therebetween a hollow interior space, an insert within said space, said insert being formed of a solid sheet curved to the chamber of the blade, and fins carried by said blade and extending between said insert and said blade walls, said walls, said insert and said fins defining in combination a plurality of flow passages for cooling fluid, each said flow passage commencing at one end of the blade and extending in a spiral-like manner around and along the insert.
2. A blade as claimed in claim 1, which includes a plurality of discrete pimples extending between the sides of the insert and the blade walls within the flow passages for providing location for the insert.
3. A blade as claimed in claim 2, wherein the fins and pimples are formed integrally with the blade walls.
4. A blade as claimed in claim 1, wherein holes are formed in the region of the leading edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the flow passages.
5. A blade as claimed in claim 1, wherein holes are formed in the region of the trailing edge of the blade for allowing cooling fluid to pass to the exterior of the blade from the flow passages.
6. A blade as claimed in claim 1, wherein the insert has a root end portion at which it is secured within the blade by brazing, the insert being otherwise unattached to allow for differential expansion.
7. A blade as claimed in claim 1, wherein at least some of the flow passages are arranged for individually serving one edge only of the blade.
8. A gas turbine rotor having a plurality of blades as claimed in claim 1, arranged for supply with cooling fluid at their root ends.
9. A gas turbine rotor according to claim 8, which includes a shroud connecting together the outer ends of the blades, the arrangement being such that cooling fluid may pass from said flow passages into cooling relation with the shroud.
10. A rotor according to claim 9, wherein the radially inner surface of the shroud is formed with channels, and the blade walls are formed with holes arranged for directing cooling fluid from the flow passages along the channels.
11. A blade as claimed in claim 6, wherein said fins are terminated immediately adjacent said root end portion of said insert.
12. A gas turbine blade having opposed walls defining therebetween a hollow interior space, and an insert within said space, said insert being formed of a solid sheet curved to the chamber of the blade and connected to said walls by brazing at the root end of the blade, being otherwise unattached to allow for differential expansion, said walls being integrally formed with fins extending between said insert and said blade walls, said walls, said insert and said fins defining in combination a plurality of flow passages for cooling fluid to flow through the blade commencing at the root end of the blade and passing in a spiral-like manner around and along the insert.
US00282778A 1971-08-25 1972-08-22 Gas turbine engine blades Expired - Lifetime US3806274A (en)

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GB3977871A GB1361256A (en) 1971-08-25 1971-08-25 Gas turbine engine blades
US28277972A 1972-08-22 1972-08-22

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Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3902820A (en) * 1973-07-02 1975-09-02 Westinghouse Electric Corp Fluid cooled turbine rotor blade
US4286122A (en) * 1978-03-13 1981-08-25 U.S. Philips Corporation Acoustic electrical conversion device with at least one capacitor electret element connected to an electronic circuit
US4515526A (en) * 1981-12-28 1985-05-07 United Technologies Corporation Coolable airfoil for a rotary machine
US5097660A (en) * 1988-12-28 1992-03-24 Sundstrand Corporation Coanda effect turbine nozzle vane cooling
US5919031A (en) * 1996-08-23 1999-07-06 Asea Brown Boveri Ag Coolable blade
WO1999061756A1 (en) * 1998-05-25 1999-12-02 Asea Brown Boveri Ab A component for a gas turbine
US6056508A (en) * 1997-07-14 2000-05-02 Abb Alstom Power (Switzerland) Ltd Cooling system for the trailing edge region of a hollow gas turbine blade
US6254346B1 (en) * 1997-03-25 2001-07-03 Mitsubishi Heavy Industries, Ltd. Gas turbine cooling moving blade
US20050031452A1 (en) * 2003-08-08 2005-02-10 Siemens Westinghouse Power Corporation Cooling system for an outer wall of a turbine blade
US6923247B1 (en) * 1998-11-09 2005-08-02 Alstom Cooled components with conical cooling passages
US20070172354A1 (en) * 2004-02-27 2007-07-26 Mats Annerfeldt Blade or vane for a turbomachine
US7563072B1 (en) 2006-09-25 2009-07-21 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall spiral flow cooling circuit
US20090185903A1 (en) * 2006-04-21 2009-07-23 Beeck Alexander R Turbine Blade
CN102207007A (en) * 2010-03-31 2011-10-05 通用电气公司 Interior cooling channels
US20130034429A1 (en) * 2010-04-14 2013-02-07 Dave Carter Blade or vane for a turbomachine
US20140093391A1 (en) * 2012-09-28 2014-04-03 Solar Turbines Incorporated Cooled turbine blade with trailing edge flow metering
US9228439B2 (en) 2012-09-28 2016-01-05 Solar Turbines Incorporated Cooled turbine blade with leading edge flow redirection and diffusion
US9314838B2 (en) 2012-09-28 2016-04-19 Solar Turbines Incorporated Method of manufacturing a cooled turbine blade with dense cooling fin array
US10641103B2 (en) 2017-01-19 2020-05-05 United Technologies Corporation Trailing edge configuration with cast slots and drilled filmholes
US10822963B2 (en) * 2018-12-05 2020-11-03 Raytheon Technologies Corporation Axial flow cooling scheme with castable structural rib for a gas turbine engine

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GB1400285A (en) * 1972-08-02 1975-07-16 Rolls Royce Hollow cooled vane or blade for a gas turbine engine
US4153386A (en) * 1974-12-11 1979-05-08 United Technologies Corporation Air cooled turbine vanes
US4278400A (en) * 1978-09-05 1981-07-14 United Technologies Corporation Coolable rotor blade
FR2468727A1 (en) * 1979-10-26 1981-05-08 Snecma IMPROVEMENT TO COOLED TURBINE AUBES
FR2473621A1 (en) * 1980-01-10 1981-07-17 Snecma DAWN OF TURBINE DISPENSER
US5232343A (en) * 1984-05-24 1993-08-03 General Electric Company Turbine blade
GB2189553B (en) * 1986-04-25 1990-05-23 Rolls Royce Cooled vane
US5022817A (en) * 1989-09-12 1991-06-11 Allied-Signal Inc. Thermostatic control of turbine cooling air
US5279111A (en) * 1992-08-27 1994-01-18 Inco Limited Gas turbine cooling
DE19860787B4 (en) * 1998-12-30 2007-02-22 Alstom Turbine blade with cooling channels
US7497655B1 (en) * 2006-08-21 2009-03-03 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall impingement and vortex cooling
US7722326B2 (en) * 2007-03-13 2010-05-25 Siemens Energy, Inc. Intensively cooled trailing edge of thin airfoils for turbine engines
US10443407B2 (en) * 2016-02-15 2019-10-15 General Electric Company Accelerator insert for a gas turbine engine airfoil

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DE853534C (en) * 1943-02-27 1952-10-27 Maschf Augsburg Nuernberg Ag Air-cooled gas turbine blade
US2888243A (en) * 1956-10-22 1959-05-26 Pollock Robert Stephen Cooled turbine blade
US3057597A (en) * 1959-08-20 1962-10-09 Jr Andre J Meyer Modification and improvements to cooled blades
US3370829A (en) * 1965-12-20 1968-02-27 Avco Corp Gas turbine blade construction
US3527544A (en) * 1968-12-12 1970-09-08 Gen Motors Corp Cooled blade shroud
US3606574A (en) * 1969-10-23 1971-09-20 Gen Electric Cooled shrouded turbine blade
US3628880A (en) * 1969-12-01 1971-12-21 Gen Electric Vane assembly and temperature control arrangement
US3635587A (en) * 1970-06-02 1972-01-18 Gen Motors Corp Blade cooling liner

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DE853534C (en) * 1943-02-27 1952-10-27 Maschf Augsburg Nuernberg Ag Air-cooled gas turbine blade
US2888243A (en) * 1956-10-22 1959-05-26 Pollock Robert Stephen Cooled turbine blade
US3057597A (en) * 1959-08-20 1962-10-09 Jr Andre J Meyer Modification and improvements to cooled blades
US3370829A (en) * 1965-12-20 1968-02-27 Avco Corp Gas turbine blade construction
US3527544A (en) * 1968-12-12 1970-09-08 Gen Motors Corp Cooled blade shroud
US3606574A (en) * 1969-10-23 1971-09-20 Gen Electric Cooled shrouded turbine blade
US3628880A (en) * 1969-12-01 1971-12-21 Gen Electric Vane assembly and temperature control arrangement
US3635587A (en) * 1970-06-02 1972-01-18 Gen Motors Corp Blade cooling liner

Cited By (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3902820A (en) * 1973-07-02 1975-09-02 Westinghouse Electric Corp Fluid cooled turbine rotor blade
US4286122A (en) * 1978-03-13 1981-08-25 U.S. Philips Corporation Acoustic electrical conversion device with at least one capacitor electret element connected to an electronic circuit
US4515526A (en) * 1981-12-28 1985-05-07 United Technologies Corporation Coolable airfoil for a rotary machine
US5097660A (en) * 1988-12-28 1992-03-24 Sundstrand Corporation Coanda effect turbine nozzle vane cooling
US5919031A (en) * 1996-08-23 1999-07-06 Asea Brown Boveri Ag Coolable blade
US6254346B1 (en) * 1997-03-25 2001-07-03 Mitsubishi Heavy Industries, Ltd. Gas turbine cooling moving blade
US6056508A (en) * 1997-07-14 2000-05-02 Abb Alstom Power (Switzerland) Ltd Cooling system for the trailing edge region of a hollow gas turbine blade
WO1999061756A1 (en) * 1998-05-25 1999-12-02 Asea Brown Boveri Ab A component for a gas turbine
US6382907B1 (en) 1998-05-25 2002-05-07 Abb Ab Component for a gas turbine
US6923247B1 (en) * 1998-11-09 2005-08-02 Alstom Cooled components with conical cooling passages
US20050031452A1 (en) * 2003-08-08 2005-02-10 Siemens Westinghouse Power Corporation Cooling system for an outer wall of a turbine blade
US6955525B2 (en) 2003-08-08 2005-10-18 Siemens Westinghouse Power Corporation Cooling system for an outer wall of a turbine blade
US20070172354A1 (en) * 2004-02-27 2007-07-26 Mats Annerfeldt Blade or vane for a turbomachine
US7674092B2 (en) 2004-02-27 2010-03-09 Siemens Aktiengesellschaft Blade or vane for a turbomachine
US20090185903A1 (en) * 2006-04-21 2009-07-23 Beeck Alexander R Turbine Blade
US8092175B2 (en) * 2006-04-21 2012-01-10 Siemens Aktiengesellschaft Turbine blade
US7563072B1 (en) 2006-09-25 2009-07-21 Florida Turbine Technologies, Inc. Turbine airfoil with near-wall spiral flow cooling circuit
CN102207007A (en) * 2010-03-31 2011-10-05 通用电气公司 Interior cooling channels
US20110243711A1 (en) * 2010-03-31 2011-10-06 General Electric Company Interior cooling channels
US8636463B2 (en) * 2010-03-31 2014-01-28 General Electric Company Interior cooling channels
CN102207007B (en) * 2010-03-31 2015-04-29 通用电气公司 Interior cooling channels
US9181808B2 (en) * 2010-04-14 2015-11-10 Siemens Aktiengesellschaft Blade or vane for a turbomachine
US20130034429A1 (en) * 2010-04-14 2013-02-07 Dave Carter Blade or vane for a turbomachine
US20140093391A1 (en) * 2012-09-28 2014-04-03 Solar Turbines Incorporated Cooled turbine blade with trailing edge flow metering
US9206695B2 (en) * 2012-09-28 2015-12-08 Solar Turbines Incorporated Cooled turbine blade with trailing edge flow metering
US9228439B2 (en) 2012-09-28 2016-01-05 Solar Turbines Incorporated Cooled turbine blade with leading edge flow redirection and diffusion
US9314838B2 (en) 2012-09-28 2016-04-19 Solar Turbines Incorporated Method of manufacturing a cooled turbine blade with dense cooling fin array
US10641103B2 (en) 2017-01-19 2020-05-05 United Technologies Corporation Trailing edge configuration with cast slots and drilled filmholes
US10822963B2 (en) * 2018-12-05 2020-11-03 Raytheon Technologies Corporation Axial flow cooling scheme with castable structural rib for a gas turbine engine

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US3782852A (en) 1974-01-01
GB1361256A (en) 1974-07-24
AU4587272A (en) 1974-03-07
AU467301B2 (en) 1975-11-27
DE2241192C3 (en) 1978-03-09
DE2241192B2 (en) 1977-07-21
DE2241192A1 (en) 1973-03-08
FR2150476B1 (en) 1979-04-06
FR2150476A1 (en) 1973-04-06
SE378645B (en) 1975-09-08

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