WO2006115484A1 - Gas turbine engine cooling system and method - Google Patents
Gas turbine engine cooling system and method Download PDFInfo
- Publication number
- WO2006115484A1 WO2006115484A1 PCT/US2005/013950 US2005013950W WO2006115484A1 WO 2006115484 A1 WO2006115484 A1 WO 2006115484A1 US 2005013950 W US2005013950 W US 2005013950W WO 2006115484 A1 WO2006115484 A1 WO 2006115484A1
- Authority
- WO
- WIPO (PCT)
- Prior art keywords
- fuel
- rotor
- turbine engine
- gas turbine
- flow path
- Prior art date
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, 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
- F02C7/22—Fuel supply systems
- F02C7/224—Heating fuel before feeding to the burner
-
- 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
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
-
- 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
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
- F01D5/085—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
- F01D5/088—Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor in a closed cavity
-
- 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
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, 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
- F02C7/12—Cooling of plants
- F02C7/16—Cooling of plants characterised by cooling medium
-
- 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
- F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- FIG. 1 illustrates a cross-sectional view of gas turbine engine incorporating a system for cooling the turbine rotor and the associated blades thereof;
- FIG. 2 illustrates a isometric view of a portion of a bladed rotor and associated fragmentary sectional views thereof
- FIG. 3 illustrates a diagram of the relationship between fuel pressure and radial location within the bladed rotor of the gas turbine engine illustrated in Fig. 1;
- FIG. 4 illustrates a diagram of the density and state of fuel as a function of temperature and pressure;
- FIG. 5 illustrates a cross-sectional view of a portion of a bladed rotor and an associated thermosiphon process therein;
- FIG. 6 illustrates a cross-sectional view of gas turbine engine incorporating another embodiment of a system for cooling the turbine rotor and the associated blades thereof.
- a gas turbine engine 10 fuel 12 and air 14 are combusted in a combustion chamber 16 so as to generate relatively hot, relatively high pressure exhaust gases 18.1 which are directed through a turbine 20 comprising a bladed rotor 22, e.g. a rotor 24 incorporating a plurality of blades 26 on the periphery thereof.
- the turbine 20 is operatively coupled to a shaft assembly 28, e.g. with a bolt 30 through an associated flange 32, and the shaft assembly 28 is supported from the housing 34 of the gas turbine engine 10 by one or more bearings 35 that provide for rotation of the shaft assembly 28 and turbine 20 relative thereto.
- the action of the exhaust gases 18.1 against the blades 26 rotates the turbine 20 and the shaft assembly 28, which, for example, is operatively coupled to a compressor (not illustrated) that provides for pumping the air 14 into the combustion chamber 16.
- the exhaust gases 18.2 discharged from the turbine 20 are at a relatively lower pressure than the exhaust gases 18.1 upstream thereof as a result of the work done by the exhaust gases 18.1 on the turbine 20.
- the air 14 supplied to the gas turbine engine 10 is relatively hot, which contributes to increased temperature of the exhaust gases 18.1, and which is not sufficiently cool to otherwise provide for adequately cooling the turbine 20, so that the temperature of the associated blades 26 can become excessive.
- the fuel 12 is generally sufficiently cool to provide sufficient cooling capacity to cool the gas turbine engine 10, and particularly, to cool the turbine 20 thereof, which might otherwise be susceptible to thermally induced failure, whereby the gas turbine engine 10 is cooled by directing fuel 12 from a source of fuel 36 through the rotor 24 and blades 26 of the turbine 20 to cool the rotor 24 and the blades 26 of the turbine 20, and then combusting this fuel 12 — heated by the cooling process — in the combustion chamber 16.
- fuel 12 from a source of fuel 36 comprising a fuel tank and an associated fuel pump is supplied through a first control valve 37 to an orifice 38 that is relatively fixed with respect to the housing 34 of the gas turbine engine 10.
- the fuel 12 is discharged from the orifice 38 into an inlet 40 of a first rotary fluid trap 42 operatively coupled to the rotor 24 so as to rotate therewith.
- the outlet 44 of the first rotary fluid trap 42 is in fluid communication with a first portion 46.1 of a first cavity 46 that is bounded by a portion of a first side 48 of the rotor 24 and by a first bounding surface of an aft cover 50 of which the first rotary fluid trap 42 is a part.
- the first rotary fluid trap 42 comprises a passage 52 that provides for fluid communication between the inlet 40 and the outlet 44, wherein, in accordance with the teachings of U.S. Patent Nos. 4,870,825 and 6,269,647, and of U.S. Application No. 10/249,967, each of which is incorporated herein by reference, the passage 52 is adapted so the when the first rotary fluid trap 42 is rotated, a centrifugal acceleration at any point within the passage 52 is greater than a centrifugal acceleration at any point on either the inlet 40 or the outlet 44.
- the rotating passage 52 when the rotating passage 52 is filled with a relatively high density medium, such as liquid fuel 12.1, the radial levels of the inlet 40 and outlet 44 will be equal when there is no pressure differential therebetween, and will be otherwise unequal by an amount dependent upon the magnitude of the pressure differential and the speed of rotation.
- a relatively low pressure supply of liquid fuel 12.1 to an inlet 40 of a passage 52 feeding a relatively high pressure region at the outlet 44 the passage 52 can prevent backflow therethrough.
- the first rotary fluid trap 42 provides for isolating the pressure in the first cavity 46 ⁇ which can be relatively high — from the pressure at the inlet 40 of the passage 52 — which is relatively lower — thereby providing for supplying fuel 12 to the inlet 40 of the first rotary fluid trap 42 across a rotary junction 54 between the rotating inlet 40 and the relatively fixed orifice 38, whereby liquid fuel 12.1 sprayed from the relatively fixed orifice 38 becomes captured by an internal trough 56 associated with the inlet 40 of the first rotary fluid trap 42 as a result of centrifugal acceleration acting upon the liquid fuel 12.1 upon striking the internal trough 56 and rotating therewith.
- the aft cover 50 comprises an intermediate rim 58 and an outer rim 60 that engage respective first 62.1 and second 62.2 lips formed on the first side 48 of the rotor 24.
- the outer rim 60 is sealed to the second lip 62.2 so as to prevent leakage of fuel 12 from the joint therebetween.
- the intermediate rim 58 incorporates at least one passage 64 that provides for fluid communication between first 46.1 and second 46.2 portions of the first cavity 46.
- the second portion 46.2 of the first cavity 46 is in fluid communication with a plurality of first passages 66 that extend through the rotor 24.
- each first passage 66 has a first opening 68 on the first side 48 of the rotor 24, and a second opening 70 on a second side 72 of the rotor 24, the first 48 and second 72 sides being opposite to one another.
- the first passages 66 are in fluid communication with a second portion 74.2 of a second cavity 74 that is bounded by a portion of the second side 72 of the rotor 24 and by a second bounding surface of a forward cover 50, wherein the forward cover 50 comprises an intermediate rim 78 and an outer rim 80 that engage respective first 82.1 and second 82.2 lips formed on the second side 72 of the rotor 24.
- the outer rim 80 is sealed to the second lip 82.2 so as to prevent leakage of fuel 12 from the joint therebetween.
- the intermediate rim 78 incorporates at least one passage 84 that provides for fluid communication between the second portion 74.2 of the second cavity 74 and a first portion 74.1 thereof.
- the first portion 74.1 of the second cavity 74 is in fluid communication with the interior 86 of a shaft 88 of the shaft assembly 28 via at least one passage 90 through the shaft 88, and the interior 86 of the shaft 88 is in fluid communication with a first discharge orifice 92 through at least one other passage 94 through the shaft 88.
- the first discharge orifice 92 is in fluid communication with the combustion chamber 16, and thereby provides for a discharge of fuel 12 directly from the rotating shaft 88 to the combustion chamber 16.
- the first discharge orifice 92 is, for example, a part of a second rotary fluid trap 96 that provides for isolating the relatively high pressure of the combustion chamber 16 from the relatively lower pressure of the interior of the shaft 88 and the first portion 74.1 of the second cavity 74, whereby the principles of structure and operation of the second rotary fluid trap 96 are the same as those of the first rotary fluid trap 42 described hereinabove.
- the first passages 66 and associated first 68 and second 70 openings are substantially uniform in size and shape, and uniformly distributed so as to provide a mechanically balanced rotor 24.
- the axial shape 98 of the first passages 66 is adapted to at least partially conform to a profile of the associated blades 26.
- the first passages 66 have chevron axial shape 98.1 so as to at least partially conform to the camber of the blades 26.
- a first set 66.1 of first passages 66 extend through the rotor 24 at associated circumferential locations that are substantially between the associated circumferential locations of the associated blades 26, and a second set 66.2 of first passages 66 extend through the rotor 24 at associated circumferential locations that are substantially aligned with the associated circumferential locations of the associated blades 26, whereby the first 66.1 and second 66.2 sets of first passages 66 are interleaved with respect to one another.
- Each of the blades 26 incorporates a plurality of second passages 100 that extend substantially radially therewithin, each of which at a first end 102 thereof intersects an associated first passage 66 of the second set 66.2 that is aligned therewith.
- the second passages 100 are substantially linear along the length thereof.
- the diameter of the second passages 100 within a particular blade 26 can be adapted in accordance with the associated blade thickness proximate thereto, so as to provide sufficient heat transfer between the outer surface 104 of the blade 26 and the surface 106 of the associated second passage 100 while providing for adequate blade strength.
- the distal second ends 108 of the second passages 100 are terminated in a third cavity 110 proximate to a tip 112 of the blade 26, wherein the third cavity 110 provides for fluid communication amongst the second passages 100 within the associated blade 26.
- the third cavity 110 is formed by a end cap 114 that is separated from the second ends 108 of the second passages 100, and which is secured at its periphery to the edge 116 of the blade 26.
- the gas turbine engine 10 comprises a rotatable portion 118 that is rotatable with respect to a housing 34 of the gas turbine engine 10, wherein the rotatable portion 118 comprises the turbine 20 / bladed rotor 22, comprising the rotor 24 and the blades 26; the aft cover 50 and associated first rotary fluid trap 42; the forward cover 50; and the shaft assembly 28 / shaft 88 and associated first discharge orifice 92 / second rotary fluid trap 96, all of which rotate in unison with a rotating frame of reference.
- the fuel 12 After discharge from the relatively fixed orifice 38, the fuel 12 is contained within the rotatable portion 118 until discharge directly into the combustion chamber 16 from the first discharge orifice 92 of the rotatable portion 118 in the rotating frame of reference Accordingly, because all of the elements of the rotatable portion 118 rotate in unison with the rotating frame of reference, these elements can be readily sealed to one another as necessary to contain the fuel 12 therein, for example, at the junctions of the outer rims 60, 80 of the first 50 and second 76 bounding surfaces with the second lips 62.2, 82.2 of the rotor 24, which could otherwise be problematic if it were necessary to provide for sealing across a relatively moving junction of elements to be sealed to one another.
- liquid fuel 12.1 provided by the source of fuel 36 and regulated by the first control valve 37 is discharged from the relatively fixed orifice 38 into the internal trough 56 of the inlet 40 of the first rotary fluid trap 42.
- the discharged liquid fuel 12.1 is captured by the internal trough 56 as a result of the centrifugal acceleration acting upon the discharged liquid fuel 12.1 which commences rotation with the rotatable portion 118 upon impact with the internal trough 56 or the liquid fuel 12.1 contained therein.
- Liquid fuel 12.1 entering the inlet 40 of the first rotary fluid trap 42 is pumped through the associated passage 52 of the first rotary fluid trap 42 by the action of centrifugal acceleration forces acting upon the liquid fuel 12.1 contained within the first rotary fluid trap 42, and this action of centrifugal acceleration forces also isolates the relatively low pressure at the inlet 40 of the first rotary fluid trap 42 from a relatively high pressure at the outlet 44 thereof.
- the fuel 12 Upon exiting the outlet 44 of the first rotary fluid trap 42, the fuel 12 is accelerated radially outwards, whereby liquid fuel 12.1 ⁇ which is relatively dense in comparison with associated fuel vapor — tends to follow the inside of the aft cover 50.
- the hottest portion of the turbine 20 / bladed rotor 22 are the blades 26 which are directly exposed to the relatively hot exhaust gases 18.1 from the combustion chamber 16. Heat from the blades 26 is transferred to the rotor 24 and associated first 50 and second 76 bounding surfaces, which provides for heating any fuel 12 in the associated first 46 and second 74 cavities that are adjacent to the first 48 and second 72 sides of the rotor 24. Accordingly, the temperature of the rotor 24 and adjacent aft cover 50 increases with decreasing distance from the blades 26, so that fuel 12 within the first cavity 46 is heated as it flows radially outwards. Furthermore, referring to Fig.
- Fuel 12 in the first set 66.1 of first passages 66 flows therethrough, out of the second openings 70 thereof, and then into the second portion 74.2 of the second cavity 74, and in the process, provides for cooling the rim 120 of the rotor 24 in the regions between the blades 26.
- the centrifugal acceleration field causes relatively dense fuel 12 in the second set 66.2 of first passages 66 to flow into the second passages 100 intersecting therewith, which displaces fuel 12 therein that has become relatively more heated and less dense, responsive to a thermosiphon process that is driven by the centrifugal acceleration field and by the decrease in density as fuel 12 becomes heated as a result of heat transfer from the blades 26 which cools the blades 26.
- the thermosiphon flow 122 within the second passages 100 and between the first 66 and second 100 passages causes a continuous exchange of relatively cooler fuel 12.2 for relatively hotter fuel 12.3, which is also illustrated by the points "D", "E” and "F” in Figs. 4 and 5.
- the relatively hotter fuel 12.3 ultimately flows through the second opening 70 of the second set 66.2 of first passages 66 and into the second portion 74.2 of the second cavity.
- the second set 66.2 of first passages 66 provides for the flow of fuel 12 either directly therethrough from the first opening 68 to the second opening 70 along a first flow path 124, which provides for cooling the rotor 24 at the base of the associated blade 26; or indirectly after first flowing along a second flow path 126 which includes one or more second passages 100 responsive to a thermosiphon process, which provides for cooling the associated blade 26 of the turbine 20.
- the relatively less dense heated fuel 12.3 in the second portion 74.1 of the second cavity 74 flows through the passage 84 into the first portion 74.1 of the second cavity 74 after being displaced by relatively more dense less heated fuel 12 from the first passages 66.
- the pressure thereof is reduced, and the fuel 12 is cooled by exchange of heat with the relatively cooler surroundings, transforming from a superheated vapor to a saturated vapor then a saturated liquid, as indicated by the locus of points labeled "G" on Fig. 4 corresponding to the location similarly labeled in Fig. 1.
- the fuel 12 then flows through the passage 90 through the shaft 88, through the interior 86 of the shaft 88, out of a second passage through the shaft 88 and into the combustion chamber 16 through the first discharge orifice 92 which is part of a second rotary fluid trap 96.
- the above-described system and method of cooling the turbine 20 -- wherein fuel 12 is delivered by a first fuel distribution circuit 128 from the source of fuel 36 through the first control valve 37 to the rotor 24 and blades 26 ⁇ is beneficially used when the turbine 20 is at a temperature that is sufficient to vaporize the fuel 12 so as to mitigate against interfering with the mechanical balance of the turbine 20.
- liquid fuel 12.1 supplied from the source of fuel 36 is regulated by a second control valve 132 and delivered to a second discharge orifice 134, for example, a part of a third rotary fluid trap 136, for example, operatively coupled to the shaft 88, wherein fuel 12 is supplied from the second control valve 132 through a separate passage 138 in the interior of the shaft 88.
- the first 37 and second 130 control valves would be controlled so that all of the fuel 12 to the gas turbine engine 10 is delivered by the second fuel distribution circuit 130 during startup and warm- up conditions.
- the second fuel distribution circuit 130 provides for a sufficient amount of fuel 12 to maintain an idle operating condition, and the remaining fuel 12 is provided by the first control valve 38 via the first fuel distribution circuit 128 responsive to operationally dependent demand.
- all of the fuel 12 might be delivered by the first fuel distribution circuit 128 after the gas turbine engine 10 has warmed up.
- some other relative distribution of fuel 12 between the first 128 and second 130 fuel distribution circuits is used.
- the first discharge orifice 92 and associated second rotary fluid trap 96 are incorporated in the forward cover 76, so as to provide for injection of fuel 12 directly into the combustion chamber 16 therefrom, without involving the shaft 88 as an associated flow path.
- the first fuel distribution circuit 128 In addition to providing for cooling the blades 26 and rotor 24 of the turbine 20, the first fuel distribution circuit 128 also provides for a regenerative recovery of heat from the exhaust 18.1 so as to provide for improved operating efficiency, particularly for stationary applications.
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
Claims
Priority Applications (10)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
RU2007143518/06A RU2477379C2 (en) | 2005-04-25 | 2005-04-25 | Method of gas turbine engine cooling (versions), method of controlling gas turbine engine, gas turbine engine and its rotor |
CN2005800495910A CN101184912B (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
MX2007013030A MX2007013030A (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method. |
JP2008508807A JP2008538804A (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and cooling method |
US11/912,544 US8057163B2 (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
EP05738670A EP1875058A4 (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
CA002605391A CA2605391A1 (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
PCT/US2005/013950 WO2006115484A1 (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
IL186875A IL186875A0 (en) | 2005-04-25 | 2007-10-24 | Gas turbine engine cooling system and method |
NO20075814A NO20075814L (en) | 2005-04-25 | 2007-11-13 | Gas turbine engine cooling system and method |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
PCT/US2005/013950 WO2006115484A1 (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
Publications (1)
Publication Number | Publication Date |
---|---|
WO2006115484A1 true WO2006115484A1 (en) | 2006-11-02 |
Family
ID=37215026
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
PCT/US2005/013950 WO2006115484A1 (en) | 2005-04-25 | 2005-04-25 | Gas turbine engine cooling system and method |
Country Status (9)
Country | Link |
---|---|
US (1) | US8057163B2 (en) |
EP (1) | EP1875058A4 (en) |
JP (1) | JP2008538804A (en) |
CN (1) | CN101184912B (en) |
CA (1) | CA2605391A1 (en) |
IL (1) | IL186875A0 (en) |
MX (1) | MX2007013030A (en) |
RU (1) | RU2477379C2 (en) |
WO (1) | WO2006115484A1 (en) |
Cited By (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2009107418A1 (en) * | 2008-02-27 | 2009-09-03 | 三菱重工業株式会社 | Turbine disc and gas turbine |
Families Citing this family (13)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US8763405B2 (en) * | 2008-04-09 | 2014-07-01 | Williams International Co., L.L.C. | Gas turbine engine rotary injection system and method |
US8820092B2 (en) * | 2008-04-09 | 2014-09-02 | 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 |
US8671696B2 (en) * | 2009-07-10 | 2014-03-18 | Leonard M. Andersen | Method and apparatus for increasing thrust or other useful energy output of a device with a rotating element |
CN103133060B (en) * | 2011-11-25 | 2015-10-21 | 中航商用航空发动机有限责任公司 | Gas turbine engine and control the method in gap between turbine casing and rotor blade |
JP5868802B2 (en) * | 2012-07-20 | 2016-02-24 | 株式会社東芝 | Turbine |
US10273816B2 (en) * | 2013-02-12 | 2019-04-30 | United Technologies Corporation | Wear pad to prevent cracking of fan blade |
US9790859B2 (en) | 2013-11-20 | 2017-10-17 | United Technologies Corporation | Gas turbine engine vapor cooled centrifugal impeller |
EP3080400B1 (en) | 2013-12-12 | 2019-04-10 | United Technologies Corporation | Gas turbine engine rotor and corresponding method of cooling |
US10393382B2 (en) | 2016-11-04 | 2019-08-27 | General Electric Company | Multi-point injection mini mixing fuel nozzle assembly |
ES2950768T3 (en) | 2018-05-11 | 2023-10-13 | Carrier Corp | Surface Plasmon Resonance Detection System |
US11859535B2 (en) * | 2021-03-09 | 2024-01-02 | Rtx Corporation | Fuel-cooled engine component(s) |
CN113983497B (en) * | 2021-10-22 | 2022-08-19 | 北京航空航天大学 | Supercritical combustion chamber and aircraft engine |
Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3550372A (en) * | 1967-08-03 | 1970-12-29 | Ass Elect Ind | Method and apparatus for operating a gas turbine with gases including contaminants of a residual fuel |
US6357217B1 (en) * | 1999-07-26 | 2002-03-19 | Alstom (Switzerland) Ltd | Endothermic cooling of guide vanes and/or moving blades in a gas turbine |
Family Cites Families (46)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US2647368A (en) * | 1949-05-09 | 1953-08-04 | Hermann Oestrich | Method and apparatus for internally cooling gas turbine blades with air, fuel, and water |
US2866313A (en) * | 1950-04-14 | 1958-12-30 | Power Jets Res & Dev Ltd | Means for cooling turbine-blades by liquid jets |
US2783613A (en) * | 1951-01-18 | 1957-03-05 | Helmut P G A R Von Zborowski | Cooling system employing fuel for cooling the blades of gas turbine engines |
US2849210A (en) * | 1953-01-19 | 1958-08-26 | Gen Motors Corp | Turbine blade cooling system |
US2883151A (en) * | 1954-01-26 | 1959-04-21 | Curtiss Wright Corp | Turbine cooling system |
US2945671A (en) * | 1955-02-10 | 1960-07-19 | Rolls Royce | Bladed rotor constructions for fluid machines |
US3355883A (en) * | 1966-01-24 | 1967-12-05 | Gen Motors Corp | Closed loop heat exchanger for a gas turbine engine |
US3734639A (en) * | 1968-01-25 | 1973-05-22 | Gen Motors Corp | Turbine cooling |
US3600890A (en) * | 1968-11-29 | 1971-08-24 | United Aircraft Corp | Turbine cooling construction |
US3728042A (en) * | 1971-08-27 | 1973-04-17 | Westinghouse Electric Corp | Axial positioner and seal for cooled rotor blade |
US3756020A (en) * | 1972-06-26 | 1973-09-04 | Curtiss Wright Corp | Gas turbine engine and cooling system therefor |
US3902819A (en) * | 1973-06-04 | 1975-09-02 | United Aircraft Corp | Method and apparatus for cooling a turbomachinery blade |
US4134709A (en) * | 1976-08-23 | 1979-01-16 | General Electric Company | Thermosyphon liquid cooled turbine bucket |
US4259037A (en) * | 1976-12-13 | 1981-03-31 | General Electric Company | Liquid cooled gas turbine buckets |
US4156582A (en) * | 1976-12-13 | 1979-05-29 | General Electric Company | Liquid cooled gas turbine buckets |
US4190398A (en) * | 1977-06-03 | 1980-02-26 | General Electric Company | Gas turbine engine and means for cooling same |
US4179240A (en) * | 1977-08-29 | 1979-12-18 | Westinghouse Electric Corp. | Cooled turbine blade |
JPS5477820A (en) * | 1977-12-02 | 1979-06-21 | Hitachi Ltd | Method of cooling gas turbine blade |
JPS54101012A (en) * | 1978-01-26 | 1979-08-09 | Denriyoku Chuo Kenkyusho | Method of liquiddcooling gas turbine vane |
US4260336A (en) * | 1978-12-21 | 1981-04-07 | United Technologies Corporation | Coolant flow control apparatus for rotating heat exchangers with supercritical fluids |
US5003766A (en) * | 1984-10-10 | 1991-04-02 | Paul Marius A | Gas turbine engine |
US4845941A (en) * | 1986-11-07 | 1989-07-11 | Paul Marius A | Gas turbine engine operating process |
US4769996A (en) * | 1987-01-27 | 1988-09-13 | Teledyne Industries, Inc. | Fuel transfer system for multiple concentric shaft gas turbine engines |
US4870825A (en) * | 1988-06-09 | 1989-10-03 | Williams International Corporation | Rotary fuel injection system |
RU1802173C (en) * | 1990-08-01 | 1993-03-15 | Самарское Конструкторское Бюро Машиностроения | Gas turbine wheel |
US5125793A (en) * | 1991-07-08 | 1992-06-30 | The United States Of America As Represented By The Secretary Of The Air Force | Turbine blade cooling with endothermic fuel |
US5224713A (en) * | 1991-08-28 | 1993-07-06 | General Electric Company | Labyrinth seal with recirculating means for reducing or eliminating parasitic leakage through the seal |
US5299418A (en) * | 1992-06-09 | 1994-04-05 | Jack L. Kerrebrock | Evaporatively cooled internal combustion engine |
US5313790A (en) * | 1992-12-17 | 1994-05-24 | Alliedsignal Inc. | Endothermic fluid based thermal management system |
US5323602A (en) * | 1993-05-06 | 1994-06-28 | Williams International Corporation | Fuel/air distribution and effusion cooling system for a turbine engine combustor burner |
KR100389990B1 (en) * | 1995-04-06 | 2003-11-17 | 가부시끼가이샤 히다치 세이사꾸쇼 | Gas turbine |
US5568984A (en) * | 1995-09-05 | 1996-10-29 | Williams International Corporation | Fuel lubricated bearing |
JP3448145B2 (en) * | 1995-11-24 | 2003-09-16 | 三菱重工業株式会社 | Heat recovery type gas turbine rotor |
US5857836A (en) * | 1996-09-10 | 1999-01-12 | Aerodyne Research, Inc. | Evaporatively cooled rotor for a gas turbine engine |
JP3621523B2 (en) * | 1996-09-25 | 2005-02-16 | 株式会社東芝 | Gas turbine rotor blade cooling system |
JP3925985B2 (en) * | 1997-05-07 | 2007-06-06 | 株式会社東芝 | Combined cycle power plant |
JPH11210492A (en) * | 1998-01-20 | 1999-08-03 | Toshiba Corp | Fuel supply device for gas turbine plant, method of warming-up operation of the device and method of cooling-down operation of the same |
JP3426952B2 (en) * | 1998-03-03 | 2003-07-14 | 三菱重工業株式会社 | Gas turbine blade platform |
EP1117913B1 (en) * | 1998-09-24 | 2003-06-04 | Siemens Aktiengesellschaft | Fuel preheating in a gas turbine |
JP3752568B2 (en) * | 1999-01-22 | 2006-03-08 | 株式会社日立製作所 | Gas turbine fuel heating system |
DE60044733D1 (en) * | 1999-03-10 | 2010-09-02 | Williams Int Co Llc | rocket engine |
US6192670B1 (en) * | 1999-06-15 | 2001-02-27 | Jack L. Kerrebrock | Radial flow turbine with internal evaporative blade cooling |
US6672075B1 (en) * | 2002-07-18 | 2004-01-06 | University Of Maryland | Liquid cooling system for gas turbines |
US6925812B2 (en) * | 2003-05-22 | 2005-08-09 | Williams International Co., L.L.C. | Rotary injector |
US6988367B2 (en) * | 2004-04-20 | 2006-01-24 | Williams International Co. L.L.C. | Gas turbine engine cooling system and method |
US8763405B2 (en) * | 2008-04-09 | 2014-07-01 | Williams International Co., L.L.C. | Gas turbine engine rotary injection system and method |
-
2005
- 2005-04-25 EP EP05738670A patent/EP1875058A4/en not_active Withdrawn
- 2005-04-25 MX MX2007013030A patent/MX2007013030A/en not_active Application Discontinuation
- 2005-04-25 US US11/912,544 patent/US8057163B2/en active Active
- 2005-04-25 CA CA002605391A patent/CA2605391A1/en not_active Abandoned
- 2005-04-25 CN CN2005800495910A patent/CN101184912B/en not_active Expired - Fee Related
- 2005-04-25 WO PCT/US2005/013950 patent/WO2006115484A1/en active Application Filing
- 2005-04-25 RU RU2007143518/06A patent/RU2477379C2/en not_active Application Discontinuation
- 2005-04-25 JP JP2008508807A patent/JP2008538804A/en active Pending
-
2007
- 2007-10-24 IL IL186875A patent/IL186875A0/en unknown
Patent Citations (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3550372A (en) * | 1967-08-03 | 1970-12-29 | Ass Elect Ind | Method and apparatus for operating a gas turbine with gases including contaminants of a residual fuel |
US6357217B1 (en) * | 1999-07-26 | 2002-03-19 | Alstom (Switzerland) Ltd | Endothermic cooling of guide vanes and/or moving blades in a gas turbine |
Non-Patent Citations (1)
Title |
---|
See also references of EP1875058A4 * |
Cited By (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2009107418A1 (en) * | 2008-02-27 | 2009-09-03 | 三菱重工業株式会社 | Turbine disc and gas turbine |
JP2009203870A (en) * | 2008-02-27 | 2009-09-10 | Mitsubishi Heavy Ind Ltd | Turbine disk and gas turbine |
CN101960092A (en) * | 2008-02-27 | 2011-01-26 | 三菱重工业株式会社 | Turbine disc and gas turbine |
KR101245094B1 (en) * | 2008-02-27 | 2013-03-18 | 미츠비시 쥬고교 가부시키가이샤 | Turbine disc and gas turbine |
US8770919B2 (en) | 2008-02-27 | 2014-07-08 | Mitsubishi Heavy Industries, Ltd. | Turbine disk and gas turbine |
Also Published As
Publication number | Publication date |
---|---|
RU2477379C2 (en) | 2013-03-10 |
IL186875A0 (en) | 2008-02-09 |
CN101184912B (en) | 2010-05-12 |
EP1875058A1 (en) | 2008-01-09 |
RU2007143518A (en) | 2009-06-10 |
US20080199303A1 (en) | 2008-08-21 |
CN101184912A (en) | 2008-05-21 |
US8057163B2 (en) | 2011-11-15 |
EP1875058A4 (en) | 2011-03-30 |
MX2007013030A (en) | 2008-03-18 |
JP2008538804A (en) | 2008-11-06 |
CA2605391A1 (en) | 2006-11-02 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US8057163B2 (en) | Gas turbine engine cooling system and method | |
US6988367B2 (en) | Gas turbine engine cooling system and method | |
JP3607331B2 (en) | Seal structure of axial gas turbine engine | |
EP2546471B1 (en) | Tip clearance control for turbine blades | |
US7942009B1 (en) | Gas turbine engine with an air cooled bearing | |
US7775764B2 (en) | Gas turbine engine rotor ventilation arrangement | |
US9963994B2 (en) | Method and apparatus for clearance control utilizing fuel heating | |
US20040088998A1 (en) | Turbine | |
US8763405B2 (en) | Gas turbine engine rotary injection system and method | |
US8820092B2 (en) | Gas turbine engine cooling system and method | |
US6357999B1 (en) | Gas turbine engine internal air system | |
US20150098791A1 (en) | Method and system for passive clearance control in a gas turbine engine | |
US9464527B2 (en) | Fuel-cooled bladed rotor of a gas turbine engine | |
US7909565B2 (en) | Turbomachine, in particular a gas turbine | |
US11319830B2 (en) | Control of clearance between aircraft rotor blades and a casing | |
US4358926A (en) | Turbine engine with shroud cooling means | |
CN111022190B (en) | Heat pipe in turbine engine | |
EP3536933B1 (en) | Ring segment and gas turbine | |
JP2003520315A (en) | Gas turbine engine | |
US20200378307A1 (en) | Thermal management of a shaft | |
JPH1136983A (en) | Turbine frame structure of turbofan engine | |
CN101900032B (en) | Gas turbine engine cooling system and method |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
WWE | Wipo information: entry into national phase |
Ref document number: 200580049591.0 Country of ref document: CN |
|
121 | Ep: the epo has been informed by wipo that ep was designated in this application | ||
DPE1 | Request for preliminary examination filed after expiration of 19th month from priority date (pct application filed from 20040101) | ||
WWE | Wipo information: entry into national phase |
Ref document number: 2005738670 Country of ref document: EP |
|
WWE | Wipo information: entry into national phase |
Ref document number: 2605391 Country of ref document: CA |
|
WWE | Wipo information: entry into national phase |
Ref document number: MX/a/2007/013030 Country of ref document: MX |
|
ENP | Entry into the national phase |
Ref document number: 2008508807 Country of ref document: JP Kind code of ref document: A |
|
WWE | Wipo information: entry into national phase |
Ref document number: 186875 Country of ref document: IL |
|
WWE | Wipo information: entry into national phase |
Ref document number: 11912544 Country of ref document: US |
|
NENP | Non-entry into the national phase |
Ref country code: DE |
|
WWE | Wipo information: entry into national phase |
Ref document number: 2007143518 Country of ref document: RU |
|
WWP | Wipo information: published in national office |
Ref document number: 2005738670 Country of ref document: EP |