US20100303610A1 - Cooled gas turbine stator assembly - Google Patents
Cooled gas turbine stator assembly Download PDFInfo
- Publication number
- US20100303610A1 US20100303610A1 US12/474,822 US47482209A US2010303610A1 US 20100303610 A1 US20100303610 A1 US 20100303610A1 US 47482209 A US47482209 A US 47482209A US 2010303610 A1 US2010303610 A1 US 2010303610A1
- Authority
- US
- United States
- Prior art keywords
- stator assembly
- gas path
- disposed
- annular cavity
- path platform
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- 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
- 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
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
-
- 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
- F01D9/00—Stators
- F01D9/06—Fluid supply conduits to nozzles or the like
- F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
-
- 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
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
- F05D2240/81—Cooled platforms
Definitions
- This disclosure relates generally to stators in a gas turbine engine and, in particular, to cooled stator fairings.
- Stator fairing assemblies also known in the art as stator vane assemblies and stator assemblies, are used to direct fluid flow entering or exiting rotor assemblies within a gas turbine engine.
- Each stator fairing assembly typically includes a plurality of stator fairings extending radially between an inner platform and an outer platform. The temperature of a core gas flow passing through the stator fairing assembly typically requires cooling within the stator fairing assembly.
- the inner radial region receives a circumferentially non-uniform distribution of cooling air which can lead to thermal distortion within the inner radial region and problems associated therewith. Therefore, there is a need for a stator fairing assembly with an internal structure that promotes uniform cooling within the assembly.
- a stator assembly for a gas turbine engine having an annular body, an inner gas path platform, a plurality of fairings, and at least one nozzle.
- the annular body has an outer gas path platform and a circumferentially extending annular cavity disposed radially outside of the outer gas path platform.
- the fairings extend radially between the inner gas path platform and the outer gas path platform.
- Each fairing includes a gas passage extending from the annular cavity through the inner gas path platform.
- the at least one nozzle has an inlet orifice disposed outside of the annular cavity and an exit orifice disposed within the annular cavity. The exit orifice is oriented within the annular cavity such that cooling air exiting the nozzle travels in a substantially circumferential direction within the annular cavity.
- FIG. 1 is a side sectional view illustrating one example of a stator assembly
- FIG. 2 is a cross-sectional view illustrating the stator assembly in FIG. 1 ;
- FIG. 3A illustrates the net flow through a prior art stator assembly
- FIG. 3B illustrates the net flow through the stator assembly in FIG. 1 .
- a stator assembly 100 for a gas turbine engine is provided.
- the gas turbine engine has a plurality of rotor assemblies rotatable about an axial centerline 40 of the gas turbine engine.
- the stator assembly is shown as a mid turbine stator assembly disposed between a first rotor assembly 90 and a second rotor assembly 95 within the turbine section of the gas turbine engine.
- the present invention stator assembly 100 can be utilized in a plurality of positions within a gas turbine engine and is not, therefore, limited to the aforesaid mid turbine stator position.
- the stator assembly 100 has an annular body 10 , an inner gas path platform 30 , a plurality of fairings 20 , and at least one nozzle 50 .
- the annular body 10 has an outer gas path platform 12 , an outer body panel 16 , and a circumferentially extending annular cavity 14 disposed between the outer gas path platform 12 and the outer body panel (illustrated in FIG. 2 ).
- the annular body 10 is formed as a single annular structure.
- the annular body 10 is manufactured in circumferentially defined sections that are combined to form the annular body 10 .
- seals are typically disposed between adjacent sections to reduce leakage flow therebetween.
- the outer gas path platforms 12 and the outer body panel 16 are attached to one another by mechanical fasteners. In the embodiment shown in FIG.
- the annular cavity 14 has a height that is substantially uniform around the circumference of the annular body 10 .
- the forward and aft sections of the annular body include one or more apertures that function as cooling gas leakage paths to permit flow of cooling air outside of the annular cavity 14 (not shown).
- Each of the plurality of fairings 20 extends radially between the inner gas path platform 30 and the outer gas path platform 12 .
- the annular body 10 is disposed radially outside of the fairings 20 .
- Each fairing 20 has a pair of faces extending between a leading edge 24 and a trailing edge 26 .
- each fairing 20 includes a gas passage 22 extending radially through the fairing 20 .
- the passage 22 provides a cooling air gas path from the cavity 14 , through the inner annular platform 30 , and into an inner cavity 32 defined in part by the inner gas path platform 30 .
- one or more tie rods 80 and/or service lines are disposed within one or more of the plurality of fairings 20 .
- the at least one nozzle 50 is mounted relative to the annular body 10 , and extends through the outer body panel 16 .
- the nozzle has an inlet orifice 52 disposed outside of the annular body 10 and an exit orifice 54 disposed within the cavity 14 .
- the nozzle 50 is shaped and oriented within the cavity 14 such that cooling gas passing through the nozzle 50 and into the cavity 14 exits the nozzle 50 in a substantially circumferential direction.
- FIG. 2 diagrammatically illustrates a plurality of nozzles 50 , each having an approximately ninety degree (90°) turn.
- the present invention is not limited to nozzles 50 of this configuration, however.
- the geometry of the nozzles 50 may vary from nozzle to nozzle, for example, to accommodate structure within the cavity 14 , improve circumferential gas flow, etc.
- the number of nozzles 50 may vary between applications, and is not limited.
- the nozzles are uniformly disposed circumferentially around the annular body 10 .
- the cooling gas flow within the cavity 14 collectively exiting from the plurality of nozzles 50 creates a circumferentially directed cooling gas flow within the cavity 14 (illustrated in FIG. 2 ).
- a metering plate 28 having at least one orifice is disposed within the passage 22 of at least one of the fairings 20 and is configured to create a pressure drop across the orifices in the metering plate.
- the use of a metering plate 28 within a particular passage 22 , and the characteristics of the metering plate (e.g., size and number of orifices), are varied to suit particular applications.
- a metering plate 28 may be disposed in each fairing 20 and the characteristics of each metering plate are “tuned” to create uniform cooling gas flow through each of the fairings 20 .
- the position of the metering plate 28 within each fairing 20 is the same.
- FIG. 2 illustrates the metering plates 28 disposed at or near a radially inner end of the fairings 20 . In other embodiment, the position of the metering plate 28 is varied between fairings 20 .
- hot core gas 200 flows through the first rotor stage 90 , between the outer gas path platform 12 and the inner gas path platform 30 , around each of the fairings 20 , and through to the second rotor stage 95 .
- gas flow 200 travels through the assembly 100 , it causes each fairing 20 to increase in temperature.
- cooling air flow 60 is injected into the cavity 14 through the nozzles 50 .
- the cooling air exiting the nozzles 50 is directed in a substantially circumferential direction within the cavity 14 .
- the circumferentially traveling cooling air flow 60 created by the cooling air exiting the nozzles 50 increases the uniformity of the cooling around the circumference of the annular body 10 .
- the increased uniformity of the cooling air flow within the annular body 10 also increases the uniformity of the cooling air flow through the fairings 20 (illustrated in FIG. 3B ), as compared to the prior art (illustrated in FIG. 3A ).
- the metering plates 28 disposed within the passages 22 further increase the uniformity of the cooling air flow through the fairing passages 22 , and thereby increase the uniformity of cooling gas flow into the region radially inside of the stator assembly 100 .
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A stator assembly for a gas turbine engine is provided having an annular body, an inner gas path platform, a plurality of fairings, and at least one nozzle. The annular body has an outer gas path platform and a circumferentially extending annular cavity disposed radially outside of the outer gas path platform. The fairings extend radially between the inner gas path platform and the outer gas path platform. Each fairing includes a gas passage extending from the annular cavity through the inner gas path platform. The at least one nozzle has an inlet orifice disposed outside of the annular cavity and an exit orifice disposed within the annular cavity. The exit orifice is oriented within the annular cavity such that cooling air exiting the nozzle travels in a substantially circumferential direction within the annular cavity.
Description
- 1. Technical Field
- This disclosure relates generally to stators in a gas turbine engine and, in particular, to cooled stator fairings.
- 2. Background Information
- Stator fairing assemblies, also known in the art as stator vane assemblies and stator assemblies, are used to direct fluid flow entering or exiting rotor assemblies within a gas turbine engine. Each stator fairing assembly typically includes a plurality of stator fairings extending radially between an inner platform and an outer platform. The temperature of a core gas flow passing through the stator fairing assembly typically requires cooling within the stator fairing assembly.
- It is known to radially impinge cooling air against the outer platform to cool the outer platform. A problem associated with impinging cooling air against the outer platform is that impingement cooling creates discrete cooling zones around the circumference of the outer platform. As a result, the outer platform is non-uniformly cooled around its circumference and consequently susceptible to thermal distortion. It is also known to provide hollow fairings extending between the inner and outer platforms, for the purpose of providing a cooling air gas path between the regions radially inside of the stator assembly and radially outside of the stator assembly. A problem with the hollow fairings in combination with the impingement cooling is that the cooling air flow rate through the fairings can vary from fairing to fairing. As a result, the inner radial region receives a circumferentially non-uniform distribution of cooling air which can lead to thermal distortion within the inner radial region and problems associated therewith. Therefore, there is a need for a stator fairing assembly with an internal structure that promotes uniform cooling within the assembly.
- According to the present invention, a stator assembly for a gas turbine engine is provided having an annular body, an inner gas path platform, a plurality of fairings, and at least one nozzle. The annular body has an outer gas path platform and a circumferentially extending annular cavity disposed radially outside of the outer gas path platform. The fairings extend radially between the inner gas path platform and the outer gas path platform. Each fairing includes a gas passage extending from the annular cavity through the inner gas path platform. The at least one nozzle has an inlet orifice disposed outside of the annular cavity and an exit orifice disposed within the annular cavity. The exit orifice is oriented within the annular cavity such that cooling air exiting the nozzle travels in a substantially circumferential direction within the annular cavity.
- These and other features and advantages of the present invention method and apparatus will become apparent in light of the detailed description of the invention provided below and the accompanying drawings.
-
FIG. 1 is a side sectional view illustrating one example of a stator assembly; -
FIG. 2 is a cross-sectional view illustrating the stator assembly inFIG. 1 ; -
FIG. 3A illustrates the net flow through a prior art stator assembly; and -
FIG. 3B illustrates the net flow through the stator assembly inFIG. 1 . - Now referring to
FIG. 1 , astator assembly 100 for a gas turbine engine is provided. The gas turbine engine has a plurality of rotor assemblies rotatable about anaxial centerline 40 of the gas turbine engine. InFIG. 1 , the stator assembly is shown as a mid turbine stator assembly disposed between afirst rotor assembly 90 and asecond rotor assembly 95 within the turbine section of the gas turbine engine. The presentinvention stator assembly 100 can be utilized in a plurality of positions within a gas turbine engine and is not, therefore, limited to the aforesaid mid turbine stator position. Thestator assembly 100 has anannular body 10, an innergas path platform 30, a plurality offairings 20, and at least onenozzle 50. - The
annular body 10 has an outergas path platform 12, anouter body panel 16, and a circumferentially extendingannular cavity 14 disposed between the outergas path platform 12 and the outer body panel (illustrated inFIG. 2 ). In some embodiments, theannular body 10 is formed as a single annular structure. In other embodiments theannular body 10 is manufactured in circumferentially defined sections that are combined to form theannular body 10. In those embodiments where theannular body 10 is formed from a plurality of sections, seals are typically disposed between adjacent sections to reduce leakage flow therebetween. In some embodiments, the outergas path platforms 12 and theouter body panel 16 are attached to one another by mechanical fasteners. In the embodiment shown inFIG. 1 , theannular cavity 14 has a height that is substantially uniform around the circumference of theannular body 10. The forward and aft sections of the annular body include one or more apertures that function as cooling gas leakage paths to permit flow of cooling air outside of the annular cavity 14 (not shown). - Each of the plurality of
fairings 20 extends radially between the innergas path platform 30 and the outergas path platform 12. Theannular body 10 is disposed radially outside of thefairings 20. Eachfairing 20 has a pair of faces extending between a leadingedge 24 and atrailing edge 26. Further, eachfairing 20 includes agas passage 22 extending radially through thefairing 20. Thepassage 22 provides a cooling air gas path from thecavity 14, through the innerannular platform 30, and into aninner cavity 32 defined in part by the innergas path platform 30. In some embodiments, one ormore tie rods 80 and/or service lines (not shown) are disposed within one or more of the plurality offairings 20. - The at least one
nozzle 50 is mounted relative to theannular body 10, and extends through theouter body panel 16. The nozzle has aninlet orifice 52 disposed outside of theannular body 10 and anexit orifice 54 disposed within thecavity 14. Thenozzle 50 is shaped and oriented within thecavity 14 such that cooling gas passing through thenozzle 50 and into thecavity 14 exits thenozzle 50 in a substantially circumferential direction.FIG. 2 diagrammatically illustrates a plurality ofnozzles 50, each having an approximately ninety degree (90°) turn. The present invention is not limited tonozzles 50 of this configuration, however. In those embodiments having a plurality ofnozzles 50, the geometry of thenozzles 50 may vary from nozzle to nozzle, for example, to accommodate structure within thecavity 14, improve circumferential gas flow, etc. The number ofnozzles 50 may vary between applications, and is not limited. In one embodiment having a plurality of nozzles, the nozzles are uniformly disposed circumferentially around theannular body 10. The cooling gas flow within thecavity 14 collectively exiting from the plurality ofnozzles 50 creates a circumferentially directed cooling gas flow within the cavity 14 (illustrated inFIG. 2 ). - In some embodiments, a
metering plate 28 having at least one orifice is disposed within thepassage 22 of at least one of thefairings 20 and is configured to create a pressure drop across the orifices in the metering plate. The use of ametering plate 28 within aparticular passage 22, and the characteristics of the metering plate (e.g., size and number of orifices), are varied to suit particular applications. For example, in some embodiments ametering plate 28 may be disposed in eachfairing 20 and the characteristics of each metering plate are “tuned” to create uniform cooling gas flow through each of thefairings 20. In some embodiments, the position of themetering plate 28 within eachfairing 20 is the same. For example,FIG. 2 illustrates themetering plates 28 disposed at or near a radially inner end of thefairings 20. In other embodiment, the position of themetering plate 28 is varied betweenfairings 20. - During the operation of the engine,
hot core gas 200 flows through thefirst rotor stage 90, between the outergas path platform 12 and the innergas path platform 30, around each of thefairings 20, and through to thesecond rotor stage 95. As thegas flow 200 travels through theassembly 100, it causes each fairing 20 to increase in temperature. To counteract the temperature rise, coolingair flow 60 is injected into thecavity 14 through thenozzles 50. The cooling air exiting thenozzles 50 is directed in a substantially circumferential direction within thecavity 14. The circumferentially travelingcooling air flow 60 created by the cooling air exiting thenozzles 50 increases the uniformity of the cooling around the circumference of theannular body 10. In addition, the increased uniformity of the cooling air flow within theannular body 10 also increases the uniformity of the cooling air flow through the fairings 20 (illustrated inFIG. 3B ), as compared to the prior art (illustrated inFIG. 3A ). Themetering plates 28 disposed within thepassages 22 further increase the uniformity of the cooling air flow through thefairing passages 22, and thereby increase the uniformity of cooling gas flow into the region radially inside of thestator assembly 100. - While various embodiments of the stator assembly have been disclosed, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of this disclosure. Accordingly, this disclosure is not to be restricted except in light of the attached claims and their equivalents.
Claims (12)
1. A stator assembly for a gas turbine engine, comprising:
an annular body having an outer gas path platform and a circumferentially extending annular cavity disposed radially outside of the outer gas path platform;
an inner gas path platform;
a plurality of fairings extending radially between the inner gas path platform and the outer gas path platform, each fairing including a gas passage extending from the annular cavity through the inner gas path platform; and
at least one nozzle having an inlet orifice disposed outside of the annular cavity and an exit orifice disposed within the annular cavity, wherein the exit orifice is oriented within the annular cavity such that gas exiting the nozzle travels in a substantially circumferential direction within the annular cavity.
2. The stator assembly of claim 1 , further comprising a metering plate disposed within at least one of the plurality of gas passages.
3. The stator assembly of claim 2 , where metering plate is disposes at an inner end of at least one of the plurality of gas passages.
4. The stator assembly of claim 1 , further comprising a metering plate disposed within each of the plurality of gas passages.
5. The stator assembly of claim 4 , wherein each metering plate has at least one orifice operable to meter the flow of cooling air through the gas passage in which the metering plate is disposed.
6. The stator assembly of claim 1 , wherein the annular body has a plurality of circumferentially defined sections.
7. The stator assembly of claim 6 , wherein the plurality of circumferentially defined sections include the outer gas path platform and an outer body panel
8. The stator assembly of claim 6 , further comprising a seal configured between at least two of the sections.
9. The stator assembly of claim 6 , further comprising at least one nozzle disposed in each section of the annular body.
10. The stator assembly of claim 1 , further comprising a plurality of nozzles uniformly disposed around the annular body.
11. The stator assembly of claim 1 , further comprising at least one of a tie rod and a service line disposed within at least one of the fairings.
12. The stator assembly of claim 1 , wherein the circumferentially extending annular cavity has a uniform height.
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/474,822 US20100303610A1 (en) | 2009-05-29 | 2009-05-29 | Cooled gas turbine stator assembly |
EP10250997.3A EP2264282A3 (en) | 2009-05-29 | 2010-05-28 | Cooled gas turbine stator assembly |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/474,822 US20100303610A1 (en) | 2009-05-29 | 2009-05-29 | Cooled gas turbine stator assembly |
Publications (1)
Publication Number | Publication Date |
---|---|
US20100303610A1 true US20100303610A1 (en) | 2010-12-02 |
Family
ID=42729427
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US12/474,822 Abandoned US20100303610A1 (en) | 2009-05-29 | 2009-05-29 | Cooled gas turbine stator assembly |
Country Status (2)
Country | Link |
---|---|
US (1) | US20100303610A1 (en) |
EP (1) | EP2264282A3 (en) |
Cited By (26)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20130078080A1 (en) * | 2011-09-22 | 2013-03-28 | Eric Durocher | Air system architecture for a mid-turbine frame module |
US20140013771A1 (en) * | 2012-07-13 | 2014-01-16 | United Technologies Corporation | Mid-turbine frame with threaded spokes |
WO2014105598A1 (en) * | 2012-12-29 | 2014-07-03 | United Technologies Corporation | Slider seal |
US8932011B2 (en) | 2011-10-06 | 2015-01-13 | United Technologies Corporation | Shaft assembly for a gas turbine engine |
US9010122B2 (en) | 2012-07-27 | 2015-04-21 | United Technologies Corporation | Turbine engine combustor and stator vane assembly |
EP2955336A1 (en) | 2014-06-12 | 2015-12-16 | MTU Aero Engines GmbH | Intermediate housing for a gas turbine and gas turbine with such an intermediate housing |
US9217371B2 (en) | 2012-07-13 | 2015-12-22 | United Technologies Corporation | Mid-turbine frame with tensioned spokes |
US20160032775A1 (en) * | 2013-03-13 | 2016-02-04 | United Technologies Corporation | Engine mid-turbine frame transfer tube for low pressure turbine case cooling |
US9316117B2 (en) | 2012-01-30 | 2016-04-19 | United Technologies Corporation | Internally cooled spoke |
US20160208648A1 (en) * | 2015-01-16 | 2016-07-21 | United Technologies Corporation | Cooling passages for a mid-turbine frame |
US20160245114A1 (en) * | 2014-02-14 | 2016-08-25 | United Technologies Corporation | Engine mid-turbine frame distributive coolant flow |
US20160273383A1 (en) * | 2015-03-20 | 2016-09-22 | United Technologies Corporation | Cooling passages for a mid-turbine frame |
WO2017005781A1 (en) * | 2015-07-06 | 2017-01-12 | Siemens Aktiengesellschaft | Turbine stator vane and/or turbine rotor vane with a cooling flow adjustment feature and corresponding method of adapting a vane |
US9587514B2 (en) | 2012-07-13 | 2017-03-07 | United Technologies Corporation | Vane insertable tie rods with keyed connections |
US9617870B2 (en) | 2013-02-05 | 2017-04-11 | United Technologies Corporation | Bracket for mounting a stator guide vane arrangement to a strut in a turbine engine |
US20170138209A1 (en) * | 2015-08-07 | 2017-05-18 | MTU Aero Engines AG | Device and method for influencing the temperatures in inner ring segments of a gas turbine |
US10087775B2 (en) | 2013-03-26 | 2018-10-02 | Rolls-Royce Plc | Gas turbine engine cooling arrangement |
US10247035B2 (en) | 2015-07-24 | 2019-04-02 | Pratt & Whitney Canada Corp. | Spoke locking architecture |
US10273812B2 (en) | 2015-12-18 | 2019-04-30 | Pratt & Whitney Canada Corp. | Turbine rotor coolant supply system |
US10392969B2 (en) | 2014-12-02 | 2019-08-27 | United Technologies Corporation | Moment accommodating fastener assembly |
US10443449B2 (en) | 2015-07-24 | 2019-10-15 | Pratt & Whitney Canada Corp. | Spoke mounting arrangement |
US10914193B2 (en) | 2015-07-24 | 2021-02-09 | Pratt & Whitney Canada Corp. | Multiple spoke cooling system and method |
CN114198153A (en) * | 2020-09-17 | 2022-03-18 | 中国航发商用航空发动机有限责任公司 | Turbine blade cooling system and aircraft engine |
US11473439B1 (en) * | 2021-09-23 | 2022-10-18 | General Electric Company | Gas turbine engine with hollow rotor in fluid communication with a balance piston cavity |
US11549396B2 (en) | 2019-11-12 | 2023-01-10 | Pratt & Whitney Canada Corp. | Mid-turbine frame for gas turbine engine |
US11698005B2 (en) | 2020-02-07 | 2023-07-11 | Raytheon Technologies Corporation | Flow diverter for mid-turbine frame cooling air delivery |
Citations (16)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4478551A (en) * | 1981-12-08 | 1984-10-23 | United Technologies Corporation | Turbine exhaust case design |
US4767268A (en) * | 1987-08-06 | 1988-08-30 | United Technologies Corporation | Triple pass cooled airfoil |
US4770608A (en) * | 1985-12-23 | 1988-09-13 | United Technologies Corporation | Film cooled vanes and turbines |
US4890978A (en) * | 1988-10-19 | 1990-01-02 | Westinghouse Electric Corp. | Method and apparatus for vane segment support and alignment in combustion turbines |
US5117626A (en) * | 1990-09-04 | 1992-06-02 | Westinghouse Electric Corp. | Apparatus for cooling rotating blades in a gas turbine |
US5141394A (en) * | 1990-10-10 | 1992-08-25 | Westinghouse Electric Corp. | Apparatus and method for supporting a vane segment in a gas turbine |
US5387086A (en) * | 1993-07-19 | 1995-02-07 | General Electric Company | Gas turbine blade with improved cooling |
US5498126A (en) * | 1994-04-28 | 1996-03-12 | United Technologies Corporation | Airfoil with dual source cooling |
US5609467A (en) * | 1995-09-28 | 1997-03-11 | Cooper Cameron Corporation | Floating interturbine duct assembly for high temperature power turbine |
US5702214A (en) * | 1996-11-07 | 1997-12-30 | Avibank Mfg., Inc. | Non-removable structural fastener assembly |
US5741117A (en) * | 1996-10-22 | 1998-04-21 | United Technologies Corporation | Method for cooling a gas turbine stator vane |
US5848874A (en) * | 1997-05-13 | 1998-12-15 | United Technologies Corporation | Gas turbine stator vane assembly |
US6060776A (en) * | 1995-12-30 | 2000-05-09 | Robert Bosch Gmbh | Rectifier diode |
US6508620B2 (en) * | 2001-05-17 | 2003-01-21 | Pratt & Whitney Canada Corp. | Inner platform impingement cooling by supply air from outside |
US7273345B2 (en) * | 2003-11-07 | 2007-09-25 | Alstom Technology Ltd | Method for operating a turbo engine and turbo engine |
US20080232953A1 (en) * | 2007-03-20 | 2008-09-25 | Snecma | Inter-turbine casing with cooling circuit, and turbofan comprising it |
Family Cites Families (6)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4979872A (en) * | 1989-06-22 | 1990-12-25 | United Technologies Corporation | Bearing compartment support |
US5273397A (en) * | 1993-01-13 | 1993-12-28 | General Electric Company | Turbine casing and radiation shield |
JP4040556B2 (en) * | 2003-09-04 | 2008-01-30 | 株式会社日立製作所 | Gas turbine equipment and cooling air supply method |
US6929445B2 (en) * | 2003-10-22 | 2005-08-16 | General Electric Company | Split flow turbine nozzle |
JP2006037855A (en) * | 2004-07-28 | 2006-02-09 | Mitsubishi Heavy Ind Ltd | Cylinder casing and gas turbine |
US7798765B2 (en) * | 2007-04-12 | 2010-09-21 | United Technologies Corporation | Out-flow margin protection for a gas turbine engine |
-
2009
- 2009-05-29 US US12/474,822 patent/US20100303610A1/en not_active Abandoned
-
2010
- 2010-05-28 EP EP10250997.3A patent/EP2264282A3/en not_active Withdrawn
Patent Citations (16)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4478551A (en) * | 1981-12-08 | 1984-10-23 | United Technologies Corporation | Turbine exhaust case design |
US4770608A (en) * | 1985-12-23 | 1988-09-13 | United Technologies Corporation | Film cooled vanes and turbines |
US4767268A (en) * | 1987-08-06 | 1988-08-30 | United Technologies Corporation | Triple pass cooled airfoil |
US4890978A (en) * | 1988-10-19 | 1990-01-02 | Westinghouse Electric Corp. | Method and apparatus for vane segment support and alignment in combustion turbines |
US5117626A (en) * | 1990-09-04 | 1992-06-02 | Westinghouse Electric Corp. | Apparatus for cooling rotating blades in a gas turbine |
US5141394A (en) * | 1990-10-10 | 1992-08-25 | Westinghouse Electric Corp. | Apparatus and method for supporting a vane segment in a gas turbine |
US5387086A (en) * | 1993-07-19 | 1995-02-07 | General Electric Company | Gas turbine blade with improved cooling |
US5498126A (en) * | 1994-04-28 | 1996-03-12 | United Technologies Corporation | Airfoil with dual source cooling |
US5609467A (en) * | 1995-09-28 | 1997-03-11 | Cooper Cameron Corporation | Floating interturbine duct assembly for high temperature power turbine |
US6060776A (en) * | 1995-12-30 | 2000-05-09 | Robert Bosch Gmbh | Rectifier diode |
US5741117A (en) * | 1996-10-22 | 1998-04-21 | United Technologies Corporation | Method for cooling a gas turbine stator vane |
US5702214A (en) * | 1996-11-07 | 1997-12-30 | Avibank Mfg., Inc. | Non-removable structural fastener assembly |
US5848874A (en) * | 1997-05-13 | 1998-12-15 | United Technologies Corporation | Gas turbine stator vane assembly |
US6508620B2 (en) * | 2001-05-17 | 2003-01-21 | Pratt & Whitney Canada Corp. | Inner platform impingement cooling by supply air from outside |
US7273345B2 (en) * | 2003-11-07 | 2007-09-25 | Alstom Technology Ltd | Method for operating a turbo engine and turbo engine |
US20080232953A1 (en) * | 2007-03-20 | 2008-09-25 | Snecma | Inter-turbine casing with cooling circuit, and turbofan comprising it |
Cited By (40)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20130078080A1 (en) * | 2011-09-22 | 2013-03-28 | Eric Durocher | Air system architecture for a mid-turbine frame module |
US9279341B2 (en) * | 2011-09-22 | 2016-03-08 | Pratt & Whitney Canada Corp. | Air system architecture for a mid-turbine frame module |
US8932011B2 (en) | 2011-10-06 | 2015-01-13 | United Technologies Corporation | Shaft assembly for a gas turbine engine |
US10502095B2 (en) | 2012-01-30 | 2019-12-10 | United Technologies Corporation | Internally cooled spoke |
US9512738B2 (en) | 2012-01-30 | 2016-12-06 | United Technologies Corporation | Internally cooled spoke |
US9316117B2 (en) | 2012-01-30 | 2016-04-19 | United Technologies Corporation | Internally cooled spoke |
US9222413B2 (en) * | 2012-07-13 | 2015-12-29 | United Technologies Corporation | Mid-turbine frame with threaded spokes |
US9217371B2 (en) | 2012-07-13 | 2015-12-22 | United Technologies Corporation | Mid-turbine frame with tensioned spokes |
US20140013771A1 (en) * | 2012-07-13 | 2014-01-16 | United Technologies Corporation | Mid-turbine frame with threaded spokes |
US9587514B2 (en) | 2012-07-13 | 2017-03-07 | United Technologies Corporation | Vane insertable tie rods with keyed connections |
US9010122B2 (en) | 2012-07-27 | 2015-04-21 | United Technologies Corporation | Turbine engine combustor and stator vane assembly |
WO2014105598A1 (en) * | 2012-12-29 | 2014-07-03 | United Technologies Corporation | Slider seal |
US9617870B2 (en) | 2013-02-05 | 2017-04-11 | United Technologies Corporation | Bracket for mounting a stator guide vane arrangement to a strut in a turbine engine |
US20160032775A1 (en) * | 2013-03-13 | 2016-02-04 | United Technologies Corporation | Engine mid-turbine frame transfer tube for low pressure turbine case cooling |
US10087782B2 (en) * | 2013-03-13 | 2018-10-02 | United Technologies Corporation | Engine mid-turbine frame transfer tube for low pressure turbine case cooling |
US10087775B2 (en) | 2013-03-26 | 2018-10-02 | Rolls-Royce Plc | Gas turbine engine cooling arrangement |
US20160245114A1 (en) * | 2014-02-14 | 2016-08-25 | United Technologies Corporation | Engine mid-turbine frame distributive coolant flow |
US9803501B2 (en) * | 2014-02-14 | 2017-10-31 | United Technologies Corporation | Engine mid-turbine frame distributive coolant flow |
EP2955336A1 (en) | 2014-06-12 | 2015-12-16 | MTU Aero Engines GmbH | Intermediate housing for a gas turbine and gas turbine with such an intermediate housing |
US9938858B2 (en) | 2014-06-12 | 2018-04-10 | MTU Aero Engines AG | Mid-frame for a gas turbine and gas turbine having such a mid-frame |
US10392969B2 (en) | 2014-12-02 | 2019-08-27 | United Technologies Corporation | Moment accommodating fastener assembly |
US20160208648A1 (en) * | 2015-01-16 | 2016-07-21 | United Technologies Corporation | Cooling passages for a mid-turbine frame |
US9856750B2 (en) * | 2015-01-16 | 2018-01-02 | United Technologies Corporation | Cooling passages for a mid-turbine frame |
US9732628B2 (en) * | 2015-03-20 | 2017-08-15 | United Technologies Corporation | Cooling passages for a mid-turbine frame |
US20160273383A1 (en) * | 2015-03-20 | 2016-09-22 | United Technologies Corporation | Cooling passages for a mid-turbine frame |
CN107849925A (en) * | 2015-07-06 | 2018-03-27 | 西门子股份公司 | Turbine stator blade and/or turbine rotor blade and the corresponding method for adapting to blade with cooling flow regulation feature |
WO2017005781A1 (en) * | 2015-07-06 | 2017-01-12 | Siemens Aktiengesellschaft | Turbine stator vane and/or turbine rotor vane with a cooling flow adjustment feature and corresponding method of adapting a vane |
US10669859B2 (en) | 2015-07-06 | 2020-06-02 | Siemens Aktiengesellschaft | Turbine stator vane and/or turbine rotor vane with a cooling flow adjustment feature and corresponding method of adapting a vane |
US10920612B2 (en) | 2015-07-24 | 2021-02-16 | Pratt & Whitney Canada Corp. | Mid-turbine frame spoke cooling system and method |
US10247035B2 (en) | 2015-07-24 | 2019-04-02 | Pratt & Whitney Canada Corp. | Spoke locking architecture |
US10914193B2 (en) | 2015-07-24 | 2021-02-09 | Pratt & Whitney Canada Corp. | Multiple spoke cooling system and method |
US10443449B2 (en) | 2015-07-24 | 2019-10-15 | Pratt & Whitney Canada Corp. | Spoke mounting arrangement |
US10590788B2 (en) * | 2015-08-07 | 2020-03-17 | MTU Aero Engines AG | Device and method for influencing the temperatures in inner ring segments of a gas turbine |
US20170138209A1 (en) * | 2015-08-07 | 2017-05-18 | MTU Aero Engines AG | Device and method for influencing the temperatures in inner ring segments of a gas turbine |
US10907490B2 (en) | 2015-12-18 | 2021-02-02 | Pratt & Whitney Canada Corp. | Turbine rotor coolant supply system |
US10273812B2 (en) | 2015-12-18 | 2019-04-30 | Pratt & Whitney Canada Corp. | Turbine rotor coolant supply system |
US11549396B2 (en) | 2019-11-12 | 2023-01-10 | Pratt & Whitney Canada Corp. | Mid-turbine frame for gas turbine engine |
US11698005B2 (en) | 2020-02-07 | 2023-07-11 | Raytheon Technologies Corporation | Flow diverter for mid-turbine frame cooling air delivery |
CN114198153A (en) * | 2020-09-17 | 2022-03-18 | 中国航发商用航空发动机有限责任公司 | Turbine blade cooling system and aircraft engine |
US11473439B1 (en) * | 2021-09-23 | 2022-10-18 | General Electric Company | Gas turbine engine with hollow rotor in fluid communication with a balance piston cavity |
Also Published As
Publication number | Publication date |
---|---|
EP2264282A2 (en) | 2010-12-22 |
EP2264282A3 (en) | 2013-10-23 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US20100303610A1 (en) | Cooled gas turbine stator assembly | |
US7527470B2 (en) | Stator turbine vane with improved cooling | |
US9759092B2 (en) | Casing cooling duct | |
US8667682B2 (en) | Method of fabricating a nearwall nozzle impingement cooled component for an internal combustion engine | |
US20200240638A1 (en) | Film-cooled multi-walled structure with one or more indentations | |
EP3124746B1 (en) | Method for cooling a turbo-engine component and turbo-engine component | |
JP5058897B2 (en) | Device for cooling the slot of a turbomachine rotor disk | |
US10317079B2 (en) | Cooling an aperture body of a combustor wall | |
US10794595B2 (en) | Stepped heat shield for a turbine engine combustor | |
US10502422B2 (en) | Cooling a quench aperture body of a combustor wall | |
JP2013083250A (en) | Gas turbine | |
US20130011238A1 (en) | Cooled ring segment | |
CN104929694A (en) | Components with compound angled cooling features and methods of manufacture | |
CN105102765A (en) | Impingement ring element attachment and sealing | |
US10233757B2 (en) | Rotor for a thermal turbomachine | |
US20040208748A1 (en) | Turbine vane cooled by a reduced cooling air leak | |
EP3044440B1 (en) | Fluid injector for cooling a gas turbine engine component | |
US10385727B2 (en) | Turbine nozzle with cooling channel coolant distribution plenum | |
US9995172B2 (en) | Turbine nozzle with cooling channel coolant discharge plenum | |
JP4234650B2 (en) | Cooled gas turbine engine blades | |
US10731855B2 (en) | Combustor panel cooling arrangements | |
US9599020B2 (en) | Turbine nozzle guide vane assembly in a turbomachine | |
CN115349047B (en) | Assembly of turbine ring and stator of turbine | |
US10590788B2 (en) | Device and method for influencing the temperatures in inner ring segments of a gas turbine | |
US11300296B2 (en) | Combustion chamber of a turbomachine |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
AS | Assignment |
Owner name: UNITED TECHNOLOGIES CORPORATION, CONNECTICUT Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:WANG, CHENG-ZHANG;DANIELS, WILLIAM A.;SONNTAG, ROBERT M.;REEL/FRAME:022772/0431 Effective date: 20090528 |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |