US3600890A - Turbine cooling construction - Google Patents

Turbine cooling construction Download PDF

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US3600890A
US3600890A US780944A US3600890DA US3600890A US 3600890 A US3600890 A US 3600890A US 780944 A US780944 A US 780944A US 3600890D A US3600890D A US 3600890DA US 3600890 A US3600890 A US 3600890A
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turbine
double walled
disk
casing
double
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US780944A
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Loren H White
David K Dorer
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Raytheon Technologies Corp
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United Aircraft Corp
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/22Fuel supply systems
    • F02C7/224Heating fuel before feeding to the burner
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/16Cooling of plants characterised by cooling medium

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  • This invention relates to gas turbine engines, and more specifically to a cooling construction for the turbine components of a gas turbine engine.
  • gas turbine performance is directly related to the temperature of the gas entering the turbine.
  • the benefits to be expected from cooling of gas turbine components, both in terms of higher output and higher efficiency, are so well known as to require no further elaboration
  • Many proposals have been directed to this end, particularly to the cooling of the turbine blades, which are the most critical point, in the turbine wheel. Notwithstanding these schemes, most gas turbines are kept within endurable temperature limits simply by diluting the motive gases with excess air, which'results in low efficiency. 1
  • the present invention avoids the penalties to efficiency by employing a cooling construction which employs a cryogenic fuel as the cooling fluid.
  • a cooling construction which employs a cryogenic fuel as the cooling fluid.
  • the use of this type system represents a significant advancement in turbine cooling technology since its effectiveness is such that turbine inlet temperatures, corresponding to stoichiometric fuel-air ratios, may be achieved. More specifically, this means that turbine inlet temperatures maybe almost doubled in relation to current engines and that substantial gains in gas turbine performance can be realized.
  • the primary object of the present invention is to provide a turbine cooling construction, the effectiveness of which is such that turbine inlet temperatures corresponding to stoichiometric fuel-air ratios may be achieved, hence providing substantial gains in gas turbine performance.
  • the present invention accomplishes the foregoing objective by using a cooling fluid which is a cryogenic fuel.
  • the coolant is supplied from an external source to a first flow path means which is in communication with passageways within both the turbine disk and turbine blades; therefore, the coolant flows from the external source through the turbine wheel and blades, hence maintaining the foregoing components at a desired temperature.
  • the actual metal temperatures of the turbine disk and blade are controlled by the coolant velocity, which is a function of the flow coefficient through the passageways of the first flow path means, the turbine disk, and the turbine blades.
  • a third flow path means is provided to the turbine nozzle upstream of the turbine disk and blades.
  • the third flow path is in communication with a second flow path hereinafter described, and transmits the coolant fluid from the turbine disk area to the turbine nozzle vanes, which contain passageways for the flow of the coolant therethrough.
  • the second flow path means hereinbefore mentioned transmits the fluid from the turbine disk exit area to the combustion chamber which is normally located upstream of the turbine nozzle vane.
  • the combustion chamber contains injector means for injecting the coolant, which is a cryogenic fuel, into the combustion chamber and hence provides an additional source of motive gases to the turbine.
  • the present invention additionally handles a problem which is generally associated with a cooling construction which utilizes a fuel as the cooling medium, the problem generally being one of an increased fire hazard inthe' engine.
  • the problem generally being one of an increased fire hazard inthe' engine.
  • the present inventiom'theturbinedisk is ,util. ized'to provide an air compartment forward or upstream of the disk, while providing a ,fuel compartment aft or downstream of the disk. This siibstantiallyminimizes the tire ":hazard within the present construction. m
  • the turbine cooling construction of the invention is shown as applied to a turbojet engine of conventional design, the Savin US. Pat. No. 2,747,367 being a typical example.
  • the engine includes a compressor, not shown herein, which discharges air into combustion chamber 2.
  • Fuel is suitably supplied to combustion chamber 2 and the products of combustion are utilized to drive turbine 4.
  • the turbine 4 includes a turbine rotor assembly including turbine wheel or disk 8, on which is mounted a plurality of turbine blades 10.
  • the center of the turbine disk 8 is closed whereby the disk 8 forms a solid wallfacing the forward compartment 60.
  • a plurality of passageways 12 Positioned within turbine disk 8 is a plurality of passageways 12, the passageways 12 having appropriate inlets l4 and outlets 16.
  • the inlets l4 and outlets 16 of the disk open to the rearward side of the solid wall formed by the disk.
  • a plurality of passageways 18, the passageways l2 and I8 being in communication with one another.
  • first flow path means 20 Extending from supply means 80 in turbine casing 6 is first flow path means 20.
  • first flow path means 20 comprises a double-wall construction, one wall being indicated by the reference character 22 and the second wall by the reference character 24.
  • first flow path means are shown as passing through a passageway 26 in exit guide vane 28 positioned downstream of the turbine 4, it therefore being clear that at least exit guide vanes 28 may be cooled by any coolant flowing through the exit guide vanes.
  • the coolant it has been found desirable from the standpoint of engine-thrust weight performance to employ a coolant with both a'coolant and a fuel. Genetically, it has been found that a cryogenic fuel satisfies this criteria.
  • the coolant is supplied to the supply means 80 from where it flows through the first flow path means 20, through exit guide vanes 28 and hence connectively cools the exit guide vanes 28.
  • the coolant continues to flow through first flow path means 20 to turbine disk inlet 14 and then internally of turbine disk 8 through passageways l2 and turbine blades 10 through passageways 18.
  • passageways l2 and 18 are in communication and it should be noted that for the sake of simplicity only one passageway in the dish, one turbine blade, one passageway in the turbine blade, one exit guide vane and one passageway in the exit guide vane have been illustrated, and that obviously in the usual engine construction, this is not the conventional construction.
  • first flow path means 20 and other flow path means hereinafter described is of significance in that at the flow coefficient through these flow path means that actually controls or determines the level of the metal temperature of the turbine disk 8 and turbine blades 10. More specifically, metal temperatures throughout the entire turbine construction are controlled by the coolant velocity which is controlled as a function of the flow coefficient of the flow paths.
  • second flow path means 30 which like first flow path 20 comprises a double-wall construction as indicated at 32 and 34.
  • second flow path means extend from outlet 16, through turbine casing 6 and passes forwardly to combustion chamber 2.
  • Combustion chamber 2 includes injector means 36 through which air from the engine compressor (not shown) and the coolant are injected into combustion chamber 2.
  • the cooling construction of the present invention includes one other flow path means, i.e., a third flow path means 40 which extends from the second flow path means 30 through turbine inlet nozzle 42 and into injector means 36. As hereinbefore explained, only one nozzle vane 42 and coolant passageway 46 are illustrated for purposes of simplicity.
  • the construction of the third flow path means is similar to the first and second flow path means and comprises a double-wall passageway as indicated by reference characters 48 and 50.
  • a problem generally associated with a cooling system using fuel as the coolant is the fire hazard involved.
  • the present embodiment of the invention minimized the fire hazard by compartmenting the engine with turbine disk 8. More specifically, all engine air remains forward of disk 8 as in compartment 60 and any free coolant or cryogenic fuel is contained aft or downstream of disk 8 as in compartment 62.
  • a gas turbine engine including a combustion section, a turbine section and an exit section, a double walled outer casing extending around said combustion section, turbine section and exit section, said turbine section including a turbine rotor mounted for rotation therein, said rotor including a rotor disk, a plurality of blades mounted on'said turbine disk, the center of said rotor disk being closed, said rotor disk thereby forming a solid wall, fixed means extending rearwardly from said disk with its outer wall forming an annular exit passageway with the casing, inlet means at the rear end of thete xi t section of the double walled outer casing for admitting "a cryogenic fluid, exit guide vanes extending between the inner wall of said doublewall casing and the outer wall of the fixed means within said annular exit passageway, said fixed means being double walled with a central passage means connecting the rear double wall to a location adjacent the rear of the disk and a concentric passage means connecting the front double wall to a second location adjacent the rear
  • double walled casing and forwardly of the turbine rotor forming an annular combustion chamber with said double walled outer casing, turbine inlet nozzles extending between the double walled outer casing and the rear end of the double walled inner casing just upstream of the blades, said nozzles having a passageway thercthrough connecting .the interior of the two double walls, a plurality of injection means connecting the double walled outer casing to the forward end of the double walled inner casing, said injection means having openings for injecting a fluid into the combination chamber.

Abstract

A turbine cooling construction for use in a gas turbine engine which permits the use of stoichiometric temperatures within the turbine. More specifically, an inlet turbine temperature in excess of 3,000* F. may be utilized.

Description

lnventors Loren H. White East Hartford, Conn.; David K. Dorer, Rochester, Mich. Appl. No. 780,944 Filed Nov. 29, 1968 Patented Aug. 24, I971 Assignee United Aircraft Corporation East Hartford, Conn.
TURBINE COOLING CONSTRUCTION References Cited UNITED STATES PATENTS 3,377,803 4/1968 Prachar .l 60/3966 FOREIGN PATENTS 1 711,985 7/1954 Great Britain 60/3966 Primary Examiner-Samuel Feinberg Anomey-Jack N. McCarthy peratures within the turbine. More specifically, an inlet turbine temperature in excess of 3,000 F. may be utilized.
lClaim,-1Dnwlngl'lg.
us. ca. 60/39., 60/267,4l5/l77,4l6/95 in. ca. F02c7/12 FieldofSeareh... 416/95; 415/177, 178, l75;60/39.66
M A I I i a W i I 1 2 36 I II me i Z 47 i'i H This invention relates to gas turbine engines, and more specifically to a cooling construction for the turbine components of a gas turbine engine.
Fundamentally, gas turbine performance is directly related to the temperature of the gas entering the turbine. The benefits to be expected from cooling of gas turbine components, both in terms of higher output and higher efficiency, are so well known as to require no further elaboration Many proposals have been directed to this end, particularly to the cooling of the turbine blades, which are the most critical point, in the turbine wheel. Notwithstanding these schemes, most gas turbines are kept within endurable temperature limits simply by diluting the motive gases with excess air, which'results in low efficiency. 1
In most of the prior art constructions, turbine disks and buckets have been cooled by circulation of air over or through the wheel and through the buckets. Since very considerable amounts of air are required for a small degree of cooling and the air must be forced through against the pressure of the turbine motive fluid, this too is inefficient.
The present invention avoids the penalties to efficiency by employing a cooling construction which employs a cryogenic fuel as the cooling fluid. The use of this type system represents a significant advancement in turbine cooling technology since its effectiveness is such that turbine inlet temperatures, corresponding to stoichiometric fuel-air ratios, may be achieved. More specifically, this means that turbine inlet temperatures maybe almost doubled in relation to current engines and that substantial gains in gas turbine performance can be realized.
SUMMARY OF THE INVENTION The primary object of the present invention is to provide a turbine cooling construction, the effectiveness of which is such that turbine inlet temperatures corresponding to stoichiometric fuel-air ratios may be achieved, hence providing substantial gains in gas turbine performance.
The present invention accomplishes the foregoing objective by using a cooling fluid which is a cryogenic fuel. The coolant is supplied from an external source to a first flow path means which is in communication with passageways within both the turbine disk and turbine blades; therefore, the coolant flows from the external source through the turbine wheel and blades, hence maintaining the foregoing components at a desired temperature. The actual metal temperatures of the turbine disk and blade are controlled by the coolant velocity, which is a function of the flow coefficient through the passageways of the first flow path means, the turbine disk, and the turbine blades.
To provide an overall turbine with a substantially increased efficiency, a third flow path means is provided to the turbine nozzle upstream of the turbine disk and blades. The third flow path is in communication with a second flow path hereinafter described, and transmits the coolant fluid from the turbine disk area to the turbine nozzle vanes, which contain passageways for the flow of the coolant therethrough. It should be clear, therefore, that the coolant cools not only the turbine disk and blades but also the turbine nozzle vanes, the metal temperature of each being a function of the flow coefficient of each flow path means. The second flow path means hereinbefore mentioned transmits the fluid from the turbine disk exit area to the combustion chamber which is normally located upstream of the turbine nozzle vane. The combustion chamber contains injector means for injecting the coolant, which is a cryogenic fuel, into the combustion chamber and hence provides an additional source of motive gases to the turbine.
The present invention additionally handles a problem which is generally associated with a cooling construction which utilizes a fuel as the cooling medium, the problem generally being one of an increased fire hazard inthe' engine. In the preferred embodiment of the present inventiom'theturbinedisk is ,util. ized'to provide an air compartment forward or upstream of the disk, while providing a ,fuel compartment aft or downstream of the disk. This siibstantiallyminimizes the tire ":hazard within the present construction. m
BRIEF DESCRIPTION O'FTHE pitiAwtNo DESCRIPTION OF THE PREFERRED EMBODIMENT Referring now to the single FIGURE, the turbine cooling construction of the invention is shown as applied to a turbojet engine of conventional design, the Savin US. Pat. No. 2,747,367 being a typical example. The engine includes a compressor, not shown herein, which discharges air into combustion chamber 2. Fuel is suitably supplied to combustion chamber 2 and the products of combustion are utilized to drive turbine 4.
Positioned around turbine 4 is turbine casing 6 within which is positioned a supply means for supplying a cooling fluid, hereinafter described in greater detail. The turbine 4 includes a turbine rotor assembly including turbine wheel or disk 8, on which is mounted a plurality of turbine blades 10. The center of the turbine disk 8 is closed whereby the disk 8 forms a solid wallfacing the forward compartment 60. Positioned within turbine disk 8 is a plurality of passageways 12, the passageways 12 having appropriate inlets l4 and outlets 16. The inlets l4 and outlets 16 of the disk open to the rearward side of the solid wall formed by the disk. Similarly positioned within each of the turbine blades 10 is a plurality of passageways 18, the passageways l2 and I8 being in communication with one another. Extending from supply means 80 in turbine casing 6 is first flow path means 20. As herein illustrated, first flow path means 20 comprises a double-wall construction, one wall being indicated by the reference character 22 and the second wall by the reference character 24. In the present embodiment, first flow path means are shown as passing through a passageway 26 in exit guide vane 28 positioned downstream of the turbine 4, it therefore being clear that at least exit guide vanes 28 may be cooled by any coolant flowing through the exit guide vanes.
As to the coolant, it has been found desirable from the standpoint of engine-thrust weight performance to employ a coolant with both a'coolant and a fuel. Genetically, it has been found that a cryogenic fuel satisfies this criteria. In the cooling construction of the present invention the coolant is supplied to the supply means 80 from where it flows through the first flow path means 20, through exit guide vanes 28 and hence connectively cools the exit guide vanes 28. The coolant continues to flow through first flow path means 20 to turbine disk inlet 14 and then internally of turbine disk 8 through passageways l2 and turbine blades 10 through passageways 18. As hereinbefore noted, the passageways l2 and 18 are in communication and it should be noted that for the sake of simplicity only one passageway in the dish, one turbine blade, one passageway in the turbine blade, one exit guide vane and one passageway in the exit guide vane have been illustrated, and that obviously in the usual engine construction, this is not the conventional construction.
It has been successfully demonstrated that by causing a cryogenic fuel to flow through first flow path means 20, and the turbine disk 8 and turbine blades 10, as hereinbefore described, that it is possible to run a stoichiometric turbine, or more specifically, a turbine with an inlet temperature in excess of 3000 F. The construction of the first flow path means 20 and other flow path means hereinafter described is of significance in that at the flow coefficient through these flow path means that actually controls or determines the level of the metal temperature of the turbine disk 8 and turbine blades 10. More specifically, metal temperatures throughout the entire turbine construction are controlled by the coolant velocity which is controlled as a function of the flow coefficient of the flow paths.
The coolant flow after convectively removing heat from the turbine 4 exits from disk 8 as at outlet 16. It enters second flow path means 30 which like first flow path 20 comprises a double-wall construction as indicated at 32 and 34. As illustrated in the present embodiment, second flow path means extend from outlet 16, through turbine casing 6 and passes forwardly to combustion chamber 2. Combustion chamber 2 includes injector means 36 through which air from the engine compressor (not shown) and the coolant are injected into combustion chamber 2. It should be clear that the present system by recovering the turbine coolant and utilizing it as a primary fuel significantly reduces engine fuel consumption and improves engine efficiency.
The cooling construction of the present invention includes one other flow path means, i.e., a third flow path means 40 which extends from the second flow path means 30 through turbine inlet nozzle 42 and into injector means 36. As hereinbefore explained, only one nozzle vane 42 and coolant passageway 46 are illustrated for purposes of simplicity. The construction of the third flow path means is similar to the first and second flow path means and comprises a double-wall passageway as indicated by reference characters 48 and 50.
A problem generally associated with a cooling system using fuel as the coolant is the fire hazard involved. The present embodiment of the invention minimized the fire hazard by compartmenting the engine with turbine disk 8. More specifically, all engine air remains forward of disk 8 as in compartment 60 and any free coolant or cryogenic fuel is contained aft or downstream of disk 8 as in compartment 62.
We claim:
1. In a gas turbine engine including a combustion section, a turbine section and an exit section, a double walled outer casing extending around said combustion section, turbine section and exit section, said turbine section including a turbine rotor mounted for rotation therein, said rotor including a rotor disk, a plurality of blades mounted on'said turbine disk, the center of said rotor disk being closed, said rotor disk thereby forming a solid wall, fixed means extending rearwardly from said disk with its outer wall forming an annular exit passageway with the casing, inlet means at the rear end of thete xi t section of the double walled outer casing for admitting "a cryogenic fluid, exit guide vanes extending between the inner wall of said doublewall casing and the outer wall of the fixed means within said annular exit passageway, said fixed means being double walled with a central passage means connecting the rear double wall to a location adjacent the rear of the disk and a concentric passage means connecting the front double wall to a second location adjacent the rear of the disk, said exit guide vanes and cooperating double walls being divided so that the double walled casing around said exit section is connected to the rear part of the double walled fixed means and the forward part of the double walled fixed means is connected to the double walled casing around the turbine section and combustion section, said disk having passageways therein which have inlets at one location and exits at another, said passageways being connected to passages in the blades, means connecting the inlets of said passageways to said central passage means and the exits to said concentric passage means, a double walled inner casing located radially inwardly from the. double walled casing and forwardly of the turbine rotor forming an annular combustion chamber with said double walled outer casing, turbine inlet nozzles extending between the double walled outer casing and the rear end of the double walled inner casing just upstream of the blades, said nozzles having a passageway thercthrough connecting .the interior of the two double walls, a plurality of injection means connecting the double walled outer casing to the forward end of the double walled inner casing, said injection means having openings for injecting a fluid into the combination chamber.

Claims (1)

1. In a gas turbine engine including a combustion section, a turbine section and an exit section, a double walled outer casing extending around said combustion section, turbine section and exit section, said turbine section including a turbine rotor mounted for rotation therein, said rotor including a rotor disk, a plurality of blades mounted on said turbine disk, the center of said rotor disk being closed, said rotor disk thereby forming a solid wall, fixed means extending rearwardly from said disk with its outer wall forming an annular exit passageway with the casing, inlet means at the rear end of the exit section of the double walled outer casing for admitting a cryogenic fluid, exit guide vanes extending between the inner wall of said double-wall casing and the outer wall of the fixed means within said annular exit passageway, said fixed means being double walled with a central passage means connecting the rear double wall to a location adjacent the rear of the disk and a concentric passage means connecting the front double wall to a second location adjacent the rear of the disk, said exit guide vanes and cooperating double walls being divided so that the double walled casing around said exit section is connected to the rear part of the double walled fixed means and the forward part of the double walled fixed means is connected to the double walled casing aroUnd the turbine section and combustion section, said disk having passageways therein which have inlets at one location and exits at another, said passageways being connected to passages in the blades, means connecting the inlets of said passageways to said central passage means and the exits to said concentric passage means, a double walled inner casing located radially inwardly from the double walled casing and forwardly of the turbine rotor forming an annular combustion chamber with said double walled outer casing, turbine inlet nozzles extending between the double walled outer casing and the rear end of the double walled inner casing just upstream of the blades, said nozzles having a passageway therethrough connecting the interior of the two double walls, a plurality of injection means connecting the double walled outer casing to the forward end of the double walled inner casing, said injection means having openings for injecting a fluid into the combination chamber.
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Cited By (22)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3756020A (en) * 1972-06-26 1973-09-04 Curtiss Wright Corp Gas turbine engine and cooling system therefor
US3841786A (en) * 1970-07-01 1974-10-15 Sulzer Ag Method and cooling system for cooling centrifugal pumps
FR2334826A1 (en) * 1975-12-10 1977-07-08 Stal Laval Turbin Ab GAS TURBINE
US4041699A (en) * 1975-12-29 1977-08-16 The Garrett Corporation High temperature gas turbine
WO1985000199A1 (en) * 1983-06-20 1985-01-17 Paul Marius A Process of intensification of the thermoenergetical cycle and air jet propulsion engines
US4522562A (en) * 1978-11-27 1985-06-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Turbine rotor cooling
US4571935A (en) * 1978-10-26 1986-02-25 Rice Ivan G Process for steam cooling a power turbine
US4845941A (en) * 1986-11-07 1989-07-11 Paul Marius A Gas turbine engine operating process
US5003766A (en) * 1984-10-10 1991-04-02 Paul Marius A Gas turbine engine
US5122033A (en) * 1990-11-16 1992-06-16 Paul Marius A Turbine blade unit
US5177954A (en) * 1984-10-10 1993-01-12 Paul Marius A Gas turbine engine with cooled turbine blades
WO1997044575A1 (en) * 1996-05-17 1997-11-27 Westinghouse Electric Corporation Closed loop air cooling system for combustion turbines
WO2000017504A1 (en) * 1998-09-24 2000-03-30 Siemens Aktiengesellschaft Fuel preheating in a gas turbine
US6988367B2 (en) 2004-04-20 2006-01-24 Williams International Co. L.L.C. Gas turbine engine cooling system and method
EP1875058A1 (en) * 2005-04-25 2008-01-09 Williams International Co., L.L.C. Gas turbine engine cooling system and method
US20110041509A1 (en) * 2008-04-09 2011-02-24 Thompson Jr Robert S Gas turbine engine cooling system and method
US20110209458A1 (en) * 2010-02-26 2011-09-01 Rolls-Royce Deutschland Ltd & Co Kg Aircraft gas turbine engine
US9453429B2 (en) 2013-03-11 2016-09-27 General Electric Company Flow sleeve for thermal control of a double-wall turbine shell and related method
US9464527B2 (en) 2008-04-09 2016-10-11 Williams International Co., Llc Fuel-cooled bladed rotor of a gas turbine engine
US10400674B2 (en) 2014-05-09 2019-09-03 United Technologies Corporation Cooled fuel injector system for a gas turbine engine and method for operating the same
US11466593B2 (en) 2020-01-07 2022-10-11 Raytheon Technologies Corporation Double walled stator housing
EP4056810A3 (en) * 2021-03-09 2022-12-28 Raytheon Technologies Corporation Fuel-cooled engine component(s)

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GB711985A (en) * 1951-01-18 1954-07-14 Helmut Philipp Georg Alexander Improvements in gas turbines
US3377803A (en) * 1960-08-10 1968-04-16 Gen Motors Corp Jet engine cooling system

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Publication number Priority date Publication date Assignee Title
GB711985A (en) * 1951-01-18 1954-07-14 Helmut Philipp Georg Alexander Improvements in gas turbines
US3377803A (en) * 1960-08-10 1968-04-16 Gen Motors Corp Jet engine cooling system

Cited By (33)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3841786A (en) * 1970-07-01 1974-10-15 Sulzer Ag Method and cooling system for cooling centrifugal pumps
US3756020A (en) * 1972-06-26 1973-09-04 Curtiss Wright Corp Gas turbine engine and cooling system therefor
FR2334826A1 (en) * 1975-12-10 1977-07-08 Stal Laval Turbin Ab GAS TURBINE
US4041699A (en) * 1975-12-29 1977-08-16 The Garrett Corporation High temperature gas turbine
US4102125A (en) * 1975-12-29 1978-07-25 The Garrett Corporation High temperature gas turbine
US4571935A (en) * 1978-10-26 1986-02-25 Rice Ivan G Process for steam cooling a power turbine
US4522562A (en) * 1978-11-27 1985-06-11 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Turbine rotor cooling
WO1985000199A1 (en) * 1983-06-20 1985-01-17 Paul Marius A Process of intensification of the thermoenergetical cycle and air jet propulsion engines
US5341636A (en) * 1984-10-10 1994-08-30 Paul Marius A Gas turbine engine operating method
US5003766A (en) * 1984-10-10 1991-04-02 Paul Marius A Gas turbine engine
US5177954A (en) * 1984-10-10 1993-01-12 Paul Marius A Gas turbine engine with cooled turbine blades
US4845941A (en) * 1986-11-07 1989-07-11 Paul Marius A Gas turbine engine operating process
US5122033A (en) * 1990-11-16 1992-06-16 Paul Marius A Turbine blade unit
WO1997044575A1 (en) * 1996-05-17 1997-11-27 Westinghouse Electric Corporation Closed loop air cooling system for combustion turbines
WO2000017504A1 (en) * 1998-09-24 2000-03-30 Siemens Aktiengesellschaft Fuel preheating in a gas turbine
US6381945B2 (en) 1998-09-24 2002-05-07 Siemens Aktiengesellschaft Fuel preheating in a gas turbine
US6988367B2 (en) 2004-04-20 2006-01-24 Williams International Co. L.L.C. Gas turbine engine cooling system and method
US20080199303A1 (en) * 2005-04-25 2008-08-21 Williams International Co., L.L.C. Gas Turbine Engine Cooling System and Method
EP1875058A4 (en) * 2005-04-25 2011-03-30 Williams Int Co Llc Gas turbine engine cooling system and method
US8057163B2 (en) 2005-04-25 2011-11-15 Williams International Co., L.L.C. Gas turbine engine cooling system and method
EP1875058A1 (en) * 2005-04-25 2008-01-09 Williams International Co., L.L.C. Gas turbine engine cooling system and method
US9464527B2 (en) 2008-04-09 2016-10-11 Williams International Co., Llc Fuel-cooled bladed rotor of a gas turbine engine
US20110041509A1 (en) * 2008-04-09 2011-02-24 Thompson Jr Robert S Gas turbine engine cooling system and method
US8820092B2 (en) 2008-04-09 2014-09-02 Williams International Co., L.L.C. Gas turbine engine cooling system and method
US20110209458A1 (en) * 2010-02-26 2011-09-01 Rolls-Royce Deutschland Ltd & Co Kg Aircraft gas turbine engine
US8720208B2 (en) * 2010-02-26 2014-05-13 Rolls-Royce Deutschland Ltd & Co Kg Aircraft gas turbine engine
US9453429B2 (en) 2013-03-11 2016-09-27 General Electric Company Flow sleeve for thermal control of a double-wall turbine shell and related method
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