GB2262314A - Air cooled gas turbine engine aerofoil. - Google Patents
Air cooled gas turbine engine aerofoil. Download PDFInfo
- Publication number
- GB2262314A GB2262314A GB9126187A GB9126187A GB2262314A GB 2262314 A GB2262314 A GB 2262314A GB 9126187 A GB9126187 A GB 9126187A GB 9126187 A GB9126187 A GB 9126187A GB 2262314 A GB2262314 A GB 2262314A
- Authority
- GB
- United Kingdom
- Prior art keywords
- aerofoil
- cooled
- interior
- cooling air
- local
- 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.)
- Withdrawn
Links
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
- F01D5/186—Film cooling
-
- 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
- F05D2260/60—Fluid transfer
- F05D2260/607—Preventing clogging or obstruction of flow paths by dirt, dust, or foreign particles
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A gas turbine engine aerofoil 21 is provided with a linear cooling air passage 38 interconnecting the hollow aerofoil interior 33 with its exterior surface 39. The passage 38 is generally normal with respect to part of the surface 41 defining the hollow aerofoil interior 33 local thereto. Blockage of the cooling air passage 38 by airborne particulate material is thereby substantially avoided. The passage entrance may be formed in a protection (46) supporting an air inlet tube (43) (Figure 4) or a flared recess (49) or channel (52) (Figures 5 and 6). There may be a plurality of passages. <IMAGE>
Description
AIR COOLED GAS TURBINE ENGINE AEROFOIL
This invention relates to an air cooled gas turbine engine aerofoil. It is particularly concerned with such an aerofoil provided with film cooling passages.
The turbines of gas turbine engines are required to operate at temperatures which are so high that some form of cooling of the aerofoils in the turbine is usually necessary. This is particularly so in the case of aerofoils at the upstream end of the turbine where the temperatures are highest. A common way of cooling such aerofoils is to direct cooling air, usually derived from the engine's compressor, into the aerofoil's interior.
The air is then exhausted from the aerofoil interior through appropriately pcsitioned passages which terminate at the external surface of the aerofoil.
Commonly, these cooling air passages are so configured that the cooling air which is exhausted from them forms a film of cool air over the aerofoil exterior surface. The cooling air therefore provides aerofoil cooling in two stages: a first stage as it flows within the aerofoil and a second stage as it flows in a film over the aerofoil exterior surface.
Under most operating conditions, such a method of cooling is effective in ensuring that the temperatures of turbine aerofoils remain within acceptable limits.
However if the gas turbine engine is required to operate in conditions in which the air used for cooling carries particulate material, blockage of the cooling passages by the particles can sometimes occur. The problem is particularly acute in desert areas when small sand particles can block the cooling passages. If some or all of the aerofoil cooling air passages become blocked there is a very great danger that the particular aerofoil concerned will overheat and thereby be damaged.
It is an object of the present invention to provide a cooled aerofoil in which the possibility of such cooling air passage blockage is substantially reduced.
According to the present invention, a cooled aerofoil for a gas turbine engine is provided with a hollow interior into which, in operation, a flow of cooling air is directed, said aerofoil being provided with at least one passage interconnecting said hollow interior of said aerofoil with the external surface thereof, said aerofoil being so configured that the axis of said at least one cooling passage is generally normal with respect to the part of the surface defining said hollow aerofoil interior which is local thereto, but inclined with respect to the part of said external surface of said aerofoil which is local thereto.
The present invention will now be described, by way of example, with reference to the accompanying drawings in which
Figure 1 is a sectioned side view of a ducted fan gas turbine engine which incorporates a plurality of cooled aerofoils in accordance with the present invention.
Figure 2 is a partially broken away perspective view of a portion of the ducted fan gas turbine engine shown in Figure 1 showing several of the cooled aerofoils of that engine.
Figure 3 is a sectional view of a portion of one of the aerofoils shown in Figure 2.
Figure 4 is a sectioned view of a portion of an aerofoil similar to that shown in Figure 3 which has an alternative form of internal configuration.
Figure 5 is a view on an enlarged scale of alternative configurations for the cooling air passages in the aerofoils shown in Figure 2.
Figure 6 is a view on arrow A of Figure 5.
Referring to Fig 1, a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 11, - a fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust nozzle 19.
The engine 10 functions in the conventional manner whereby air entering the inlet 11 is accelerated by the fan 12. The air exhausted from the fan is divided into two flows: a first flow which is directed to atmosphere to provide propulsive thrust and a second flow which is directed into the intermediate pressure compressor 13.
After compression by the intermediate pressure compressor 13, the air is compressed still further by the high pressure compressor 14. The ts compressed air is then mixed with fuel and the mixture combusted in the combustion equipment 15. The resultant hot combustion products are then exhausted from the combustion equipment 15 to expand through and thereby drive, the high, intermediate and low pressure turbInes 16, 17 and 18 respectively. Finally the exhaust gases pass through the exhaust nozzle 19 to provide additional propulsive thrust.
The high pressure turbine 16, a portion of which can be seen more clearly if reference is now made to Fig 2, comprises an annular array of stator nozzle guide vanes 20 and an annular array of rotor aerofoil blades 24. The annular array of nozzle guide vanes 20 is positioned immediately downstream of the combustion apparatus 15.
The guide vanes 20 are so configured as to direct the gaseous combustion products exhausted from the combustion apparatus 15 on to the rotor aerofoil blades 24 at the appropriate angle. The rotor aerofoil blades 24 are mounted for rotation on a rotatable disc 25 in the conventional manner.
Each of the nozzle guide vanes 20 comprises an aerofoil portion 21 and radially inner and outer platforms 22 and 23 respectively. The platforms 22 and 23 of adjacent nozzle guide vanes 20 cooperate to define annular walls radially inwardly and outwardly of the aerofoils 21. These thus defined walls constitute an axial portion of the gas passage through the high pressure turbine 16, specifically that portion which encloses the aerofoils 21.
Both the nozzle guide vanes 21 and the rotor aerofoil blades 24 are exposed to gaseous combustion products which are at very high temperature. In order for the vanes 21 and blades 24 to be able to withstand these temperatures and still continue to function effectively, it is necessary to cool them. In accordance with usual practice, this is achieved using relatively cool air tapped from the high pressure compressor 14.
Cooling air for cooling of the rotor aerofoil blades 24 is directed along passages (not shown) radially inwardly of the combustion apparatus 15. The air then flows as indicated by the arrows 26 over an array of pre-swirl nozzles 27 and into radial passages 28 provided in te disc 25. The radial passages 28 direct the cooling air into the aerofoil blades 24 to provide blade cooling.
The air is then exhausted from the blade tips 29 as well as through a plurality of film cooling passages 30 provided in the blade 24.
Cooling air for the cooling of the nozzle guide vanes 20 is also derived through passages (not shown) radially inwardly of the combustion apparatus 15. The cooling air is directed as shown by the arrow 31 to the radially inward regions of the nozzle guide vanes 20 to provide cooling of the radially inward platforms 22.
Film cooling passages 32 in the platforms 22 permit some of the cooling to be exhausted into the hot gasecus combustion products flowing over the platforms 22. The remainder of the cooling air flows into the hollow interiors of the nozzle guide vane aerofoils 21.
Cooling air from the high pressure compressor 14 is also directed through passages (not shown) to the radially outward regions of the nozzle guide vanes 20 as indicated by the arrow 32a. After providing cooling of the radially outer platforms 23, the cooling air also flows into the hollow interiors of the nozzle guide vane aerofoils 21.
The manner in which cooling of the nozzle guide vane aerofoils 21 is achieved can be seen more easily if reference is now made to Fig 3. Fig 3 is a cross-sectional view of a nozzle guide vane aerofoil 21.
However in the interests of clarity, some of the interior features of the aerofoil 21 have been omitted.
The aerofoil 21 is, as previously stated, hollow.
Its interior is divided into two separate compartments 33 and 34 by an internal wall 35: the first compartment 33 terminating adjacent the aerofoil leading edge 36 while the second compartment terminates adjacent the aerofoil trailing edge 37.
The present invention will be described with reference to the first compartment 33. However it will be understood that it could be applied to te second compartment 34 if so desired or indeed to other air cooled engine components such as combustor parts, aerofoil platforms etc.
As explained above, cooling air is operationally directed into the first compartment 33. The cooling air is exhausted from the first compartment 33 through a plurality of generally similar film cooling passages 38, only one of which is visible in Fig 3. The film cooling passage 38 is linear and interconnects the first compartment 33 with the external surface 39 of the aerofoil 21. In this particular case, the passage 38 interconnects the first compartment 33 with the convex, lower pressure portion of the external surface 39. It will be appreciated, however, that cooling passages 38 could, if desired, interconnect the first compartment 33 with the concave, higher pressure portion of the external surface 39.
The film cooling passage 38 is so positioned that its axis is inclined with respect to the part of the external aerofoil surface 39 local thereto. This is so as to ensure that cooling air exhausted from the first compartment 33 through the passage 38 forms a film of cooling air over part of the external surface 39, thereby locally cooling that surface part.
The wall 40 of the aerofoil 21 is of generally constant thickness. However, the region of the aerofoil wall 40 which defines the first compartment 33 is locally modified. Specifically the wall 40 is locally thickened at 42 so that it projects above the remainder of the internal surface 41 adjacent thereto and so that the internal surface 41 is locally generally normal with respect to the axis of the passage 38.
The modification of the internal surface 41 of the wall 40 by thickening at 42 serves to ensure that cooling air entering the passage 38 does so in a modified manner.
Specifically, if the air entering the passage 38 is carrying particulate material, that material tends to flow completely through the passage 38. This is in contrast to conventional cooling passages which are inclined with respect to the part of the internal surface 41 local hereto. With such inclined passages, there is a tendency for particulate material to build up around the cooling passage entrance and eventually block the passage.
Although in the embodiment of Fig 3, the wall 40 is specifically thickened at 42 in order to define the desired local shape of the internal surface 41, it may alternatively be possible to take advantage of surface modifications which are also used for other purposes. In this respect, reference should now be made to Fig 4.
In Fig 4 there is shown a cross-sectioned view of an aerofoil which is similar in certain respects to that shown in Fig 3. Accordingly common features are provided with common reference numerals. The major difference is that in the embodiment of Fig 4, a cooling air entry tube 43 is located within the first compartment 33. The cooling air entry tube 43 extends the length of the aerofoil 21 and is provided with a plurality of large openings 44 adjacent the aerofoil leading edge 36 as well as nurHerous smaller holes (not shown). Cooling air directed into the tube 43 is exhausted through the various openings 44 and holes to provide cooling of the aerofcil wall 40.
The cooling air entry tube 43 is supported in spaced apart relationship with the aerofoil internal surface 41 by a plurality of location features, two of which 45 and 46 ca be seen in Fig 4. The location feature 45 is constltuted by a hollow tube which extends the length of the aerofoil 21. It serves to maintain the tube 43 in spaced apart relationship with the aerofoil wall 40 and addit-onally functions as a seal between regions of the first compartment 33. Such a seal is necessary to preserve pressure differentials between various regions of the first compartment 33.
The location feature 46 is integral with the aerofoil wall 40 aerofoil 21 and is very similar in certain respects to the thickened wall region 42 in the embodiment of Fig 3. In particular, the location feature has a cooling air passage 47 therein which fulfils all of the requirements of the cooling air passage 38. It is, therefore, linear, generally normal to the part of the internal surface 41 local thereto and inclined with respect to the local external aerofoil surface 39. It exhausts cooling air as a film over the high pressure portion of the external surface 39.
It will be seen, therefore, that the location feature 46 serves a dual role of supporting the cooling air entry tube 43 and providing a suitable vehicle for the cooling air passages 47.
n Fig 5 there is shown, on an enlarged scale, a portion of the aerofoil wall 40 provided with two further alternative cooling passage configurations in accordance with he present invention. In the first configuration, a linear cooling passage 48 is inclined with respect to the part of the external aerofoil surface 39 local thereto in the same manner as in the the embodiments described with respect to Figs 3 and 4. However the internal surface 41 part local to the cooling air passage 48 is modified so that it defines a flared opening 49 as can be seen in Fig 6. The cooling air passage 48 is in communication with the flared opening 49 in such a manner that it is effectively normal to the portion of flared opening 49 which it intersects.Consequently, there is less tendency for airborne particulate material flowing into the passage 48 through the flared opening 49 to be deposited at the passage 48 inlet.
In the second configuration, as in the first, a linear air cooling passage 50 is inclined with respect to the part of the external aerofoil surface 39 local thereto. The internal surface 41 is locally modified by the provision of a ridge 51 and an adjacent channel 52, both of which extend the length of the aerofoil 21. The cooling air passage 50 is in communication with the channel 52 in such a manner that is is effectively normal to the portion of the channel 52 which it intersect.
Consequently, as in the case of the first configuration, there is less tendency for airborne particulate material flowing into the passage 50 to be deposited at the passage 50 inlet. In practice one or more cooling air passages 50 would be in communication with the channel 52.
It will be seen therefore that the present invention, by simple modification of the interior surface of a hollow aerofoil, provides an effective means for inhibiting the blockage of air cooling passages with airborne particulate material.
Although the present invention has been described with reference to the aerofoil portions of stator aerofoil vanes, it will be appreciated that it is also applicable to the aerofoils, including those of rotor blades and indeed to other air cooled components.
Moreover, although the cooling passages 38, 47, 48 and 50 are substantially linear, it will be appreciated that they need not necessarily be so. However it is important for the effective functioning of the present invention that they are generally normal to the parts of the hollow interior which are local to them but inclined with respect to the aerofoil external surface parts which are local to the.
Claims (12)
1. A cooled aerofoil for a gas turbine engine, said aerofoil being provided with a hollow interior into which, ifl operation, a flow of cooling air is directed, said aerofoil being provided with at least one passage interconnecting said hollow interior of said aerofoil with the external surface thereof, said aerofoil being so configured that the axis of said at least dne cooling air passage is generally normal with respect to the part of the surface defining said hollow aerofoil interior which is local thereto, but inclined with respect to the part of said external surface of said aerofoil which is local thereto.
2. A cooled aerofoil as claimed in claim 1 wherein said local part of said surface defining said hollow aerofoil interior projects above the remainder of said interior defining surface adjacent thereto.
3. A cooled aerofoil as claimed in claim 1 wherein said local part of said surface defining said hollow aerofoil interior projects below the remainder of said interior defining surface.
4. A cooled aerofoil as claimed in claim 3 wherein said local part of said surface defining said hollow aerofoil interior is in the form of a flared opening in said surface.
5. A cooled aerofoil as claimed in claim 4 wherein said flared opening in said surface is additionally in the form of an open channel.
6. A cooled aerofoil as claimed in claim 5 wherein said open channels extends lengthwise of said aerofoil.
7. A cooled aerofoil as claimed in claim 6 wherein a correspcnding lengthwise extending ridge is located adjacent said lengthwise extending channel.
8. A cooled aerofoil as claimed in claim 2 wherein said local part of said surface projecting above the remainder of said interior defining surface is additionally adapted to support a cooling air entry tube.
9. A cooled aerofoil as claimed in any one preceding claim wherein said cooling air passage is inclined with respect to the local external surface of said aerofoil in order to provide film cooling of said external surface.
10. A cooled aerofoil as claimed in any one preceding claim wherein the axis of said at least one cooling air passage is linear.
11. A cooled aerofoil as claimed in any one preceding claim wherein said aerofoil is a stator vane.
12. A cooled aerofoil substantially as hereinbefore described with reference to and as shown in the accompanying drawings.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB9126187A GB2262314A (en) | 1991-12-10 | 1991-12-10 | Air cooled gas turbine engine aerofoil. |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| GB9126187A GB2262314A (en) | 1991-12-10 | 1991-12-10 | Air cooled gas turbine engine aerofoil. |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| GB9126187D0 GB9126187D0 (en) | 1992-04-08 |
| GB2262314A true GB2262314A (en) | 1993-06-16 |
Family
ID=10705985
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB9126187A Withdrawn GB2262314A (en) | 1991-12-10 | 1991-12-10 | Air cooled gas turbine engine aerofoil. |
Country Status (1)
| Country | Link |
|---|---|
| GB (1) | GB2262314A (en) |
Cited By (34)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0677644A1 (en) * | 1994-04-14 | 1995-10-18 | Mitsubishi Jukogyo Kabushiki Kaisha | Cooled gas turbine blade |
| US5688104A (en) * | 1993-11-24 | 1997-11-18 | United Technologies Corporation | Airfoil having expanded wall portions to accommodate film cooling holes |
| WO1998037310A1 (en) * | 1997-02-20 | 1998-08-27 | Siemens Aktiengesellschaft | Turbine blade and its use in a gas turbine system |
| WO1999014465A1 (en) * | 1997-09-18 | 1999-03-25 | Siemens Aktiengesellschaft | Turbine bucket and use thereof |
| US6065932A (en) * | 1997-07-11 | 2000-05-23 | Rolls-Royce Plc | Turbine |
| EP1208290A1 (en) * | 1999-06-29 | 2002-05-29 | Allison Advanced Development Company, Inc. | Cooled airfoil |
| WO2003080998A1 (en) * | 2002-03-25 | 2003-10-02 | Alstom Technology Ltd | Cooled turbine blade |
| DE10236676A1 (en) * | 2002-08-09 | 2004-02-19 | Rolls-Royce Deutschland Ltd & Co Kg | Turbine paddle, for a gas turbine, has at least one cooling passage opening with a structured cooling air flow linking the inner zone with the outer surface |
| EP1936118A2 (en) | 2006-12-11 | 2008-06-25 | United Technologies Corporation | Turbine blade main core modifications for peripheral serpentine microcircuits |
| US7654795B2 (en) | 2005-12-03 | 2010-02-02 | Rolls-Royce Plc | Turbine blade |
| EP2584147A1 (en) * | 2011-10-21 | 2013-04-24 | Siemens Aktiengesellschaft | Film-cooled turbine blade for a turbomachine |
| EP2584148A1 (en) * | 2011-10-21 | 2013-04-24 | Siemens Aktiengesellschaft | Film-cooled turbine blade for a turbomachine |
| US8591190B2 (en) | 2008-01-10 | 2013-11-26 | Rolls-Royce Plc | Blade cooling |
| RU2506429C1 (en) * | 2012-05-31 | 2014-02-10 | Федеральное государственное унитарное предприятие "Научно-производственный центр газотурбостроения "Салют" (ФГУП "НПЦ газотурбостроения "Салют") | Gas turbine cooled working blade |
| US20140102684A1 (en) * | 2012-10-15 | 2014-04-17 | General Electric Company | Hot gas path component cooling film hole plateau |
| JP2015068340A (en) * | 2013-09-26 | 2015-04-13 | ゼネラル・エレクトリック・カンパニイ | Air foil having a low angle hole and its punching method |
| EP1760265B1 (en) | 2005-08-31 | 2015-07-15 | United Technologies Corporation | Turbine engine component with a cooling microcircuit and corresponding manufacturing method |
| US20160023275A1 (en) * | 2013-03-15 | 2016-01-28 | United Technologies Corporation | Additive manufacturing baffles, covers, and dies |
| EP2993303A1 (en) * | 2014-09-04 | 2016-03-09 | United Technologies Corporation | Gas turbine engine component with film cooling hole with pocket |
| EP2993304A1 (en) * | 2014-09-08 | 2016-03-09 | United Technologies Corporation | Gas turbine engine component with film cooling hole |
| EP3000974A1 (en) * | 2014-09-08 | 2016-03-30 | United Technologies Corporation | Casting optimized to improve suction side cooling shaped hole performance |
| EP3012407A1 (en) * | 2014-10-20 | 2016-04-27 | United Technologies Corporation | Film hole with protruding flow accumulator |
| US9328616B2 (en) | 2013-02-01 | 2016-05-03 | Siemens Aktiengesellschaft | Film-cooled turbine blade for a turbomachine |
| WO2016099663A3 (en) * | 2014-10-31 | 2016-08-11 | General Electric Company | Film cooled engine component for a gas turbine engine |
| EP3168422A1 (en) * | 2015-11-11 | 2017-05-17 | United Technologies Corporation | Low loss airflow port |
| US20170306764A1 (en) * | 2016-04-26 | 2017-10-26 | General Electric Company | Airfoil for a turbine engine |
| EP3375978A1 (en) * | 2017-03-15 | 2018-09-19 | Mitsubishi Hitachi Power Systems, Ltd. | Film cooled turbine blade |
| US10233775B2 (en) | 2014-10-31 | 2019-03-19 | General Electric Company | Engine component for a gas turbine engine |
| US10280785B2 (en) | 2014-10-31 | 2019-05-07 | General Electric Company | Shroud assembly for a turbine engine |
| US20190218940A1 (en) * | 2018-01-17 | 2019-07-18 | United Technologies Corporation | Dirt separator for internally cooled components |
| US10364684B2 (en) | 2014-05-29 | 2019-07-30 | General Electric Company | Fastback vorticor pin |
| US20190316472A1 (en) * | 2018-04-17 | 2019-10-17 | United Technologies Corporation | Double wall airfoil cooling configuration for gas turbine engine |
| US10563514B2 (en) | 2014-05-29 | 2020-02-18 | General Electric Company | Fastback turbulator |
| FR3111661A1 (en) * | 2020-06-22 | 2021-12-24 | Safran Aircraft Engines | Turbine blade with cooling system |
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| GB2184492A (en) * | 1985-12-23 | 1987-06-24 | United Technologies Corp | Film cooled vanes for turbines |
| GB2202907A (en) * | 1987-03-26 | 1988-10-05 | Secr Defence | Cooled aerofoil components |
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1991
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| GB1033759A (en) * | 1965-05-17 | 1966-06-22 | Rolls Royce | Aerofoil-shaped blade |
| GB1565361A (en) * | 1976-01-29 | 1980-04-16 | Rolls Royce | Blade or vane for a gas turbine engien |
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Cited By (50)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5688104A (en) * | 1993-11-24 | 1997-11-18 | United Technologies Corporation | Airfoil having expanded wall portions to accommodate film cooling holes |
| EP0677644A1 (en) * | 1994-04-14 | 1995-10-18 | Mitsubishi Jukogyo Kabushiki Kaisha | Cooled gas turbine blade |
| US5577889A (en) * | 1994-04-14 | 1996-11-26 | Mitsubishi Jukogyo Kabushiki Kaisha | Gas turbine cooling blade |
| WO1998037310A1 (en) * | 1997-02-20 | 1998-08-27 | Siemens Aktiengesellschaft | Turbine blade and its use in a gas turbine system |
| US6065932A (en) * | 1997-07-11 | 2000-05-23 | Rolls-Royce Plc | Turbine |
| WO1999014465A1 (en) * | 1997-09-18 | 1999-03-25 | Siemens Aktiengesellschaft | Turbine bucket and use thereof |
| EP1208290A1 (en) * | 1999-06-29 | 2002-05-29 | Allison Advanced Development Company, Inc. | Cooled airfoil |
| EP1208290A4 (en) * | 1999-06-29 | 2003-10-15 | Allison Advanced Dev Company I | Cooled airfoil |
| WO2003080998A1 (en) * | 2002-03-25 | 2003-10-02 | Alstom Technology Ltd | Cooled turbine blade |
| US7293962B2 (en) | 2002-03-25 | 2007-11-13 | Alstom Technology Ltd. | Cooled turbine blade or vane |
| DE10236676A1 (en) * | 2002-08-09 | 2004-02-19 | Rolls-Royce Deutschland Ltd & Co Kg | Turbine paddle, for a gas turbine, has at least one cooling passage opening with a structured cooling air flow linking the inner zone with the outer surface |
| EP1760265B1 (en) | 2005-08-31 | 2015-07-15 | United Technologies Corporation | Turbine engine component with a cooling microcircuit and corresponding manufacturing method |
| US7654795B2 (en) | 2005-12-03 | 2010-02-02 | Rolls-Royce Plc | Turbine blade |
| US7717676B2 (en) * | 2006-12-11 | 2010-05-18 | United Technologies Corporation | High aspect ratio blade main core modifications for peripheral serpentine microcircuits |
| EP1936118A3 (en) * | 2006-12-11 | 2011-10-05 | United Technologies Corporation | Turbine blade main core modifications for peripheral serpentine microcircuits |
| EP1936118A2 (en) | 2006-12-11 | 2008-06-25 | United Technologies Corporation | Turbine blade main core modifications for peripheral serpentine microcircuits |
| US8591190B2 (en) | 2008-01-10 | 2013-11-26 | Rolls-Royce Plc | Blade cooling |
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|---|---|
| GB9126187D0 (en) | 1992-04-08 |
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