GB2165315A - Improvements in or relating to hollow fluid cooled turbine blades - Google Patents
Improvements in or relating to hollow fluid cooled turbine blades Download PDFInfo
- Publication number
- GB2165315A GB2165315A GB08425092A GB8425092A GB2165315A GB 2165315 A GB2165315 A GB 2165315A GB 08425092 A GB08425092 A GB 08425092A GB 8425092 A GB8425092 A GB 8425092A GB 2165315 A GB2165315 A GB 2165315A
- Authority
- GB
- United Kingdom
- Prior art keywords
- passage
- wall
- turning
- end portion
- shaped end
- 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.)
- Granted
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
- F01D5/187—Convection cooling
-
- 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
Description
1 GB2165315A 1
SPECIFICATION
Improvements in or relating to hollow fluid cooled turbine blades The present invention relates to hollow fluid cooled turbine blades particularly for use in gas turbine engines.
Hollow cooled turbine blades generally com prise longitudinally extending passages which convey cooling fluid, generally air, although other fluids have been used ie, molten metals, through the turbine blade in order to pick up and dissipate heat from the walls of the tur bine blade.
In many designs of hollow cooled turbine blades the cooling fluid passes up and down adjacent longitudinally extending passages and the cooling flow must be turned through angles of the order of 180' in turning pas sages connecting the longitudinally extending passages. The present invention is concerned with this later type of turbine blade.
These types of turbine blades, generally called -multi pass blade", suffer from a prob- 90 lem in the turning passage and longitudinally extending passages. As the cooling fluid flows through a first longitudinally extending passage and enters the turning passage the cooling fluid must turn through an angle of 180' into 95 the adjacent longitudinally extending passage, and the boundary layer of cooling fluid formed on a wall separating the two adjacent longitu dinally extending passages becomes detached from the wall in the turning passages and the 100 second longitudinally extending passage caus ing generation of vortices and turbulence which gives rise to a pressure loss.
Attempts to overcome this problem have used the addition of one or more turning vanes positioned in the turning passage, but these have proved to be of only marginal be nefit, as the boundary layer still separates from the wall and vanes and, vortices and turbulence are produced in the turning passage 110 and second longitudinally extending passage.
The present invention seeks to reduce the turbulence generated in the turning passage and second longitudinally extending passage by reducing the separation of the boundary layer in order to reduce pressure losses in the cooling fluid flow through the turbine blade.
Accordingly the present invention provides a hollow fluid cooked turbine blade comprising a root, a platform and an aerofoil, the aerofoil 120 having at least two longitudinally extending passages for the flow of cooling fluid, at least on turning passage where the cooling fluid is turned from the first longitudinally extending passage into the second longitudinally extend- 125 ing passage, a wall being positioned between and separating the longitudinally extending passages, said wall having an aerodynamically shaped end portion adjacent said turning pas sage, said aerodynamically shaped end portion 130 of said wall being arranged at an angle with respect to the remainder of said wall to increase the turning radius and said aerodynamically shaped end portion of said wall extending towards a second wall defining an opposing boundary of said first passage, a turning vane having a relatively large turning radius being spaced from said aerodynamically shaped end portion of said wall to define a passage therebetween, said passage between said aerodynamically shaped end portion of said wall and said turning vane converging in the direction of flow from said first longitudinally extending passage to said second longi- tudinally extending passage, said turning vane extending around said aerodynamically shaped portion of said wall towards said second longitudinally extending passage so as to direct cooling fluid flowing from said first passage into said second passage through said converging passage formed between said aerodynamically shaped end portion of said wall and said turning vane substantially parallel to said wall, said turning vane directing the cooling fluid flowing around said turning vane substantially parallel to the cooling fluid leaving said converging passage so as to reduce separation of the boundary layer of the cooling fluid from said wall in order to reduce pressure losses in said turning passage.
The turning vane may have a tail extending longitudinally of the blade parallel to said wall in said second passage to direct cooling fluid leaving said converging passage formed between said aerodynamically shaped end portion of said wall and said turning vane substantially parallel to said wall.
The aerofoil is defined by a pressure and a suction sidewall, the suction sidewall may be provided with bleed holes, one bleed hole being provided in the converging passage adjacent the leading edge of the aerodynamically shaped end portion of the wall, one bleed hole being provided in the turning passage adjacent the leading edge of the turning vane.
The wall and the turning vane may form an angle of substantially 90' with the pressure and suction sidewalls of the aerofoil to reduce separation of the cooling fluid from said wall and turning vane at said suction wall.
The present invention will be more fully described by way of reference to the accompanying drawings, in which:- Figure 1. is a partially cut-away view of a gas turbine engine showing a hollow fluid cooled turbine blade according to the present invention.
Figure 2. is an enlarged view of the hollow fluid cooled turbine blade in Fig. 1.
Figure 3. is an enlarged view of a turning passage of a hollow fluid cooled turbine blade in Fig. 2.
Figure 4. is an enlarged view of an alternative embodiment of a turning passage of a hoillow fluid cooled turbine blade in Fig. 2.
2 GB2165315A 2 Figure 5. is an enlarged view of a further embodiment of a turning passage of a fluid control turbine blade in Fig. 2. and Figure 6. is a section along line A-A in Fig. 5 2.
Fig. 1. shows a gas turbine engine 10, which in this case is a turbofan engine, and comprises in flow series an intake 12, a fan and intermediate pressure compressor denoted jointly by numeral 14, a high pressure compressor 18, a combustion system 20, a turbine section 22 and an exhaust nozzle 24. The gas turbine engine works in a well known manner, in that air is compressed by the fan, intermediate and high pressure compressors 14 and 18 respectively, before being supplied into the combustion system 20. Fuel is injected into the combustion system and is burnt in the compressed air to produce hot gases which drive the turbines in the turbine section 22 before leaving the gas turbine engine 10 through exhaust nozzle 24. The turbines drive the fan, intermediate and high pressure compressors via shafts, not shown in the figure. The fan also provides additional thrust by driving air through a fan duct 16 around the core engine.
The turbine section 22 comprises a stator assembly 26 having a casing 28 which forms a boundary for the flow of hot gases, and which carries stator vanes 30 and forms shrouds for one or more sets of rotor blades. A rotor assembly 32 comprises one or more turbine discs 34 carrying circumferential arrays of radially extending turbine blades 36. The turbine discs 34 are connected to the compressor via shafts 38.
The turbine blades 36 are shown more clearly in Fig. 2---and are hollow cooled turbine blades of the multi pass type. The turbine blade 36 comprises a root 40, a platform 42 and an aerofoil portion 44. The aerofoil 44 comprises a leading edge 46 and a trailing edge 48, and a suction sidewall 50 and a pressure sidewall 52 extend from the leading edge to the trailing edge to define the aerofoil. The aerofoil also has a tip wall 54 at its longitudinal extremity remote from the platform 42. There are a plurality of internal walls 56, 58, 60, 62 and 64 respectively which extend longitudinally of the aerofoil from the platform 42 or the tip wall 54, to divide the aerofoil into a number of longitudinally extending passages 66, 68, 70, 72, 74 and 76 respectively. At the longitudinal extremities of the aerofoil 44 are turning passages 78, 80, 82 and 84 which interconnect passages 76, 74, 72 and 70 to form a flow passage for cooling fluid. The tip wall 54 has a number of apertures 86, 88 and 90 which extend from passages 66, 68 and 82 respectively for permitting cooling fluid to be blown over the tip of the aerofoil 44 for cooling purposes and to reduce leakage of motive fluid between the blade tip and a respective shroud structure.
The longitudinally extending internal wall 64 near the trailing edge of the aerofoil 44 has an aerodynamically shaped end portion 92 adjacent turning passage 78. The aerodynami- cally shaped end portion 92 of the wall 64 is arranged at an angle with respect to the remainder of the wall 64, and the aerodynamically shaped end portion 92 extends towards the trailing edge 48 of the pressure and suc- tion sidewalls 52 and 50 respectively.
A turning vane 94 of relatively large turning radius is positioned within the turning passage 78 and is spaced from the aerodynamically shaped end portion 92 of the wall 64 to de- fine a convergent passage 96 therebetween. The turning vane 94 extends around the aerodynamically shaped end portion 92 of the wall 64 so that the leading edge of the turning vane 94 is adjacent the leading edge of the aerodynamically shaped end portion 92 of the wall 64, and the trailing edge of the turning vane 94 extends into passage 74 and is adjacent the trailing edge of the aerodynamically shaped end portion 92 where the end portion 92 joins the remainder of the wall 64.
The convergent passage 96 converges in the direction from the leading edge to the trailing edge of the turning vane 94. The trailing edge 48 of the aerofoil 44 is provided with a plurality of discharge slots 98 or apertures for the discharge of cooling fluid from passage 76 over the trailing edge of the aerofoil 44 for film cooling thereof. The root 40 and platform 42 have passages 100, 102 and 104 which supply cooling fluid, generally air, supplied from the compressor 14 or 18 to the internal longitudinally extending passages 66 and 76 in the aerofoil 44 of the turbine blade 36.
Fig. 3. shows the turning passage 78 and the associated turning vane 94, and the aero dynamically shaped end portion 92 of the longitudinally extending wall 64 to a larger scale. The convergent passage 96 can be seen clearly and is formed between the smooth shaped aerodynamic end portion 92 and the turning vane 94.
In operation the cooling air supplied from the compressor 14 or 18 through passages 100, 102 and 104 flows into the lorigitudi- nally extending passages 66 and 76 respectively in the aerofoil 44. The cooling air in passage 66 at the leading edge 46 of the aerofoil 44 flows longitudinally of the aerofoil 44 to the tip where the cooling air flows through the aperture 86.
The cooling air in passage 76 flows longitudinally along the aerofoil 44 to the tip where the cooling air flows into the turning passage 78 and is turned through 180' to flow into passage 74. The cooling air then flow along passage 74 and is again turned through 180' in turning passage 80 into passage 72, likewise the cooling is turned through 180' in turning passages 82 and 84 into passages 70 3 GB2165315A 3 and 68 respectively. Some of the cooling air in turning passage 82 flows through aperture 90, and cooling air flowing through passage 68 flows through aperture 88 to cool the tip wall 54 of the aerofoil 44.
As discussed previously, the cooling air flowing through the turning passages becomes detached from the end portion of the longitudinally extending walls as it turns around the end portion of the walls, this gives rise to turbulence which leads to pressure losses in the passages. The use of prior art turning vanes in the turning passages of turbine blades has not been of significant benefit.
The invention reduces the pressure losses in the turning passage 78 by providing the longitudinally extending wall 64 with an aerodynamically shaped end portion 92 which is angled with respect to the remainder of the wall 64, this increases the turning radius for the flow of cooling air around the end portion 92 of the wall 64, and provides a smooth curving surface for the flow of the cooling air which aids in reducing separation of the boundary layer. A turning vane 94 extends around the aerodynamically shaped end portion 92 of the wall 64, and the turning vane 94 has a relatively large turning radius and a relatively large surface area in relation to the longitudinal width of passage 96 which assists in the retention of a boundary layer of cooling air on the turning vane 94. The cooling air flowing around the turning vane 94 is directed by the turning vane 94 to flow substantially parallel to the wall 64 as it enters passage 74.
The cooling air flowing through the convergent passage 96 between the aerodynamically shaped end portion 92 of the wall 64 and the turning vane 94 is also directed by the turning vane 94 to flow substantially parallel to the wall 64 as it enters passage 74. The use of a convergent passage 96 reduces the separation of the boundary layer of cooling air from the aerodynamically shaped end portion 92 of the wall 64.
It can be seen that by directing the cooling air leaving the convergent passage 96 and the cooling air flowing around the turning vane 94 substantially parallel to the wall 64, and to each other, into the passage 74 that turbulence and separation of the cooling air from the wall 64 is reduced because these two cooling air flows do not impinge upon and interfere with each other to direct the cooling air away form the wall 64.
The turning passage 78 also converges in the direction of flow of the cooling air from passage 76 to passage 74 and this also helps to prevent or limit separation.
Fig. 4. shows an alternative embodiment of the turning passage 78 shown in Fig. 3. The turning vane 94 is provided with a tail 106 which extends from the turning vane 94 parallel to the wall 64 in passage 74 to ensure that the cooling air flowing in the convergent passage 96 is directed along passage 74 parallel to wall 64.
Fig. 5. shows a further embodiment of the turning passage 78 shown in Fig. 3. The turn- ing passage 78 is provided with bleed holes 108 and 110 positioned in the suction sidewall 50 of the aerofoil 44. Bleed hole 108 is positioned in the convergent passage 96 between the turning vane 94 and the aero- dynamically shaped end portion 92, close to the leading edge of the end portion 92, and bleed hole 110 is positioned above the turning vane 94 close to its leading edge. The bleed holes 108 and 110 reduce the separation of the cooling air from the aerodynamically shaped end portion 92 and the turning vane 94 at their leading edges where separation commences to give a smoother flow.
Fig. 6. shows the longitudinally extending passages 66, 68, 70, 72, 74 and 76, and walls 56, 58, 60, 62, and 64 in cross-section. The walls 56, 58, 60, 62 and 64 are substantially perpendicular to the suction and pressure walls 50 and 52 respectively. The walls are arranged preferably at this angle because if the walls 56, 58, 60, 62 and 64 are angled with respect to the perpendicular to the sidewalls at angles substantially different separation of the cooling air occurs. If the walls 56, 58, 60, 62 and 64 are angled in the region of 27' to 300 with respect to the sidewalls the cooling air flowing over the turning vane 94 and the end portion 92 and other walls 56, 58, 60 and 62 flows into the pres- sure sidewall 52 at the pressure side, and away form the suction sidewall 50 at the suction side. This can be considered as a converging flow on the pressure sidewall and a diverging flow on the suction sidewall, hence separation of the cooling air flow from the suction sidewail 50 over the vane 94 and the aerodynamically shaped end portion 92 of wall 64 in passage 102.
The use of a turning vane and aerodynami- cally shaped end portion angled with respect to the remainder of an internal wall to produce a convergent passage may be applied to other internal walls of the aerofoil and not just the internal wall adjacent the trailing edge as shown in the example.
The turning vane and aerodynamically shaped end portion angled with respect to the remainder of an internal wall may be applied to the aerofoils of any hollow turbine blade having at least two longitudinally extending passages separated by a wall.
Claims (6)
1. A hollow fluid cooled turbine blade comprising a root, a platform and an aerofoil, the aerofoil having at least two longitudinally extending passages for the flow of cooling fluid, at least one turning passage where the cooling fluid is turned from the first longitudi- nally extending passage into the second longi- 4 GB2165315A 4 tudinally extending passage, a wall being positioned between and separating the longitudinally extending passages, said wall having an aerodynamically shaped end portion adjacent said turning passage, said aerodynamically shaped end portion of said wall being arranged at an angle with respect to the remainder of said wall to increase the turning radius and said aerodynamically shaped end portion of said wall extending towards a second wall defining an opposing boundary of said first passage, a turning vane having a relatively large turning radius being spaced from said aerodynamically shaped end portion of said wall to define a passage therebetween, said passage between said aerodynamically shaped end portion of said wall and said turning vane converging in the direction of flow from said first longitudinally extending passage to said second longitudinally extending passage, said turning vane extending around said aerodynamically shaped portion of said wall towards said second longitudinally extending passage so as to direct cooling fluid flowing from said first passage into said second passage through said converging passage formed between said aerodynamically shaped end portion of said wall and said turning vane substantially parallel to said wall, said turning vane directing the cooling fluid flowing around said turbine vane substantially parallel to the cooling fluid leaving said converging passage so as to reduce separation of the boundary layer of the cooling fluid from said wall in order to reduce pressure losses in said turning passage.
2. A hollow fluid cooled turbine blade as claimed in claim 1 in which the turning vane has a tail extending longitudinally of the blade parallel to said wall in said second passage to direct cooling fluid leaving said passage formed between said aerodynamically shaped end portion of said wall and said turning vane substantially parallel to said wall.
3. A hollow fluid cooled turbine blade as claimed in any of claims 1 to 2 in which the aerofoil is defined by a pressure and a suction sidewall, the suction sidewall being provided with bleed holes, one bleed hole being pro- vided in the converging passage adjacent the leading edge of the aerodynamically shaped end portion of the wail, one bleed hole being provided in the turning passage adjacent the leading edge of the turning vane.
4. A hollow fluid cooled turbine blade as claimed in any of claims 1 to 3 in which said wall and said turning vane form an angle of substantially 90' with a pressure and a suction wall of said aerofoil to reduce separation of the cooling fluid from said wall and turning vane at said suction wall.
5. A hollow fluid cooled turbine blade constructed and arranged substantially as herein described with reference to and as illustrated in the accompanying drawings.
6. A gas turbine engine comprising at least one hollow fluid cooled turbine blade as claimed in any of claims 1 to 5.
Printed in the United Kingdom for Her Majesty's Stationery Office, Dd 8818935, 1986, 4235. Published at The Patent Office, 25 Southampton Buildings, London, WC2A l AY, from which copies may be obtained.
Priority Applications (5)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB08425092A GB2165315B (en) | 1984-10-04 | 1984-10-04 | Improvements in or relating to hollow fluid cooled turbine blades |
US06/769,702 US4604031A (en) | 1984-10-04 | 1985-08-27 | Hollow fluid cooled turbine blades |
JP60208543A JPS6189901A (en) | 1984-10-04 | 1985-09-20 | Hollow fluid coling turbine blade |
FR8514256A FR2571428B1 (en) | 1984-10-04 | 1985-09-26 | HOLLOW BLADES OF TURBINES COOLED BY A FLUID AND MOTOR EQUIPPED WITH SUCH BLADES |
DE3534905A DE3534905C2 (en) | 1984-10-04 | 1985-09-30 | Cooled hollow turbine blade |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
GB08425092A GB2165315B (en) | 1984-10-04 | 1984-10-04 | Improvements in or relating to hollow fluid cooled turbine blades |
Publications (2)
Publication Number | Publication Date |
---|---|
GB2165315A true GB2165315A (en) | 1986-04-09 |
GB2165315B GB2165315B (en) | 1987-12-31 |
Family
ID=10567701
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
GB08425092A Expired GB2165315B (en) | 1984-10-04 | 1984-10-04 | Improvements in or relating to hollow fluid cooled turbine blades |
Country Status (5)
Country | Link |
---|---|
US (1) | US4604031A (en) |
JP (1) | JPS6189901A (en) |
DE (1) | DE3534905C2 (en) |
FR (1) | FR2571428B1 (en) |
GB (1) | GB2165315B (en) |
Cited By (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0465004A2 (en) * | 1990-07-03 | 1992-01-08 | ROLLS-ROYCE plc | Cooled aerofoil blade |
US5156526A (en) * | 1990-12-18 | 1992-10-20 | General Electric Company | Rotation enhanced rotor blade cooling using a single row of coolant passageways |
EP0672821A1 (en) * | 1994-02-09 | 1995-09-20 | ROLLS-ROYCE plc | Air cooled gas turbine aerofoil |
US5507621A (en) * | 1995-01-30 | 1996-04-16 | Rolls-Royce Plc | Cooling air cooled gas turbine aerofoil |
US5934874A (en) * | 1996-08-23 | 1999-08-10 | Asea Brown Boveri Ag | Coolable blade |
GB2349920A (en) * | 1999-05-10 | 2000-11-15 | Abb Alstom Power Ch Ag | Cooling arrangement for turbine blade |
US6241468B1 (en) | 1998-10-06 | 2001-06-05 | Rolls-Royce Plc | Coolant passages for gas turbine components |
US7137784B2 (en) | 2001-12-10 | 2006-11-21 | Alstom Technology Ltd | Thermally loaded component |
EP1793089A2 (en) | 2005-11-30 | 2007-06-06 | General Electric Company | Method and apparatus for reducing axial compressor blade tip flow |
EP2692991A1 (en) * | 2012-08-01 | 2014-02-05 | Siemens Aktiengesellschaft | Cooling of turbine blades or vanes |
EP2832953A1 (en) * | 2013-07-29 | 2015-02-04 | Siemens Aktiengesellschaft | Turbine blade |
DE102004011151B4 (en) * | 2003-03-19 | 2015-11-26 | Alstom Technology Ltd. | turbine blade |
EP2489838A3 (en) * | 2011-02-17 | 2017-12-27 | Rolls-Royce plc | Cooled component for the turbine of a gas turbine engine |
WO2022238051A1 (en) * | 2021-05-11 | 2022-11-17 | Siemens Energy Global GmbH & Co. KG | Rotor-blade tip including cooling configuration |
Families Citing this family (64)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US4753575A (en) * | 1987-08-06 | 1988-06-28 | United Technologies Corporation | Airfoil with nested cooling channels |
US4767268A (en) * | 1987-08-06 | 1988-08-30 | United Technologies Corporation | Triple pass cooled airfoil |
US4820122A (en) * | 1988-04-25 | 1989-04-11 | United Technologies Corporation | Dirt removal means for air cooled blades |
US4820123A (en) * | 1988-04-25 | 1989-04-11 | United Technologies Corporation | Dirt removal means for air cooled blades |
US5002460A (en) * | 1989-10-02 | 1991-03-26 | General Electric Company | Internally cooled airfoil blade |
FR2798421B1 (en) * | 1990-01-24 | 2002-10-11 | United Technologies Corp | COOLED BLADES FOR GAS TURBINE ENGINES |
FR2798423B1 (en) * | 1990-01-24 | 2002-10-11 | United Technologies Corp | GAME CONTROL FOR GAS TURBINE ENGINE TURBINE |
US5281084A (en) * | 1990-07-13 | 1994-01-25 | General Electric Company | Curved film cooling holes for gas turbine engine vanes |
US5704763A (en) * | 1990-08-01 | 1998-01-06 | General Electric Company | Shear jet cooling passages for internally cooled machine elements |
US5165852A (en) * | 1990-12-18 | 1992-11-24 | General Electric Company | Rotation enhanced rotor blade cooling using a double row of coolant passageways |
JPH0510102A (en) * | 1991-07-02 | 1993-01-19 | Hitachi Ltd | Gas turbine blade and gas turbine device |
US5299418A (en) * | 1992-06-09 | 1994-04-05 | Jack L. Kerrebrock | Evaporatively cooled internal combustion engine |
US5688107A (en) * | 1992-12-28 | 1997-11-18 | United Technologies Corp. | Turbine blade passive clearance control |
US5488825A (en) * | 1994-10-31 | 1996-02-06 | Westinghouse Electric Corporation | Gas turbine vane with enhanced cooling |
CA2262701C (en) * | 1997-06-06 | 2003-02-18 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade |
US5975851A (en) * | 1997-12-17 | 1999-11-02 | United Technologies Corporation | Turbine blade with trailing edge root section cooling |
JPH11241602A (en) * | 1998-02-26 | 1999-09-07 | Toshiba Corp | Gas turbine blade |
US6474947B1 (en) | 1998-03-13 | 2002-11-05 | Mitsubishi Heavy Industries, Ltd. | Film cooling hole construction in gas turbine moving-vanes |
EP1022435B1 (en) * | 1999-01-25 | 2009-06-03 | General Electric Company | Internal cooling circuit for a gas turbine bucket |
US6192670B1 (en) | 1999-06-15 | 2001-02-27 | Jack L. Kerrebrock | Radial flow turbine with internal evaporative blade cooling |
US6247896B1 (en) * | 1999-06-23 | 2001-06-19 | United Technologies Corporation | Method and apparatus for cooling an airfoil |
EP1132575A4 (en) * | 1999-09-16 | 2002-04-10 | Mitsubishi Heavy Ind Ltd | Film cooling hole construction in gas turbine moving-vanes |
US6257831B1 (en) * | 1999-10-22 | 2001-07-10 | Pratt & Whitney Canada Corp. | Cast airfoil structure with openings which do not require plugging |
DE50111949D1 (en) * | 2000-12-16 | 2007-03-15 | Alstom Technology Ltd | Component of a turbomachine |
US6672836B2 (en) * | 2001-12-11 | 2004-01-06 | United Technologies Corporation | Coolable rotor blade for an industrial gas turbine engine |
WO2003080998A1 (en) * | 2002-03-25 | 2003-10-02 | Alstom Technology Ltd | Cooled turbine blade |
US6746209B2 (en) * | 2002-05-31 | 2004-06-08 | General Electric Company | Methods and apparatus for cooling gas turbine engine nozzle assemblies |
US6955523B2 (en) * | 2003-08-08 | 2005-10-18 | Siemens Westinghouse Power Corporation | Cooling system for a turbine vane |
US6939102B2 (en) * | 2003-09-25 | 2005-09-06 | Siemens Westinghouse Power Corporation | Flow guide component with enhanced cooling |
US7185662B2 (en) * | 2003-11-14 | 2007-03-06 | United Technologies Corporation | Methods of preparing, cleaning and repairing article and article repaired |
EP1577497A1 (en) * | 2004-03-01 | 2005-09-21 | ALSTOM Technology Ltd | Internally cooled turbomachine blade |
US7059825B2 (en) * | 2004-05-27 | 2006-06-13 | United Technologies Corporation | Cooled rotor blade |
US7137780B2 (en) * | 2004-06-17 | 2006-11-21 | Siemens Power Generation, Inc. | Internal cooling system for a turbine blade |
US7198467B2 (en) * | 2004-07-30 | 2007-04-03 | General Electric Company | Method and apparatus for cooling gas turbine engine rotor blades |
US7144215B2 (en) * | 2004-07-30 | 2006-12-05 | General Electric Company | Method and apparatus for cooling gas turbine engine rotor blades |
US7131817B2 (en) * | 2004-07-30 | 2006-11-07 | General Electric Company | Method and apparatus for cooling gas turbine engine rotor blades |
CN101213373B (en) * | 2005-07-04 | 2012-05-09 | 贝洱两合公司 | Impeller |
US7632071B2 (en) * | 2005-12-15 | 2009-12-15 | United Technologies Corporation | Cooled turbine blade |
US10180074B2 (en) * | 2005-12-16 | 2019-01-15 | Mehmet Arik | Wireless monitoring system |
US8083485B2 (en) * | 2007-08-15 | 2011-12-27 | United Technologies Corporation | Angled tripped airfoil peanut cavity |
FR2954798B1 (en) * | 2009-12-31 | 2012-03-30 | Snecma | AUBE WITH INTERNAL VENTILATION |
US9206696B2 (en) | 2011-08-16 | 2015-12-08 | General Electric Company | Components with cooling channels and methods of manufacture |
US8864467B1 (en) * | 2012-01-26 | 2014-10-21 | Florida Turbine Technologies, Inc. | Turbine blade with serpentine flow cooling |
US8985940B2 (en) * | 2012-03-30 | 2015-03-24 | Solar Turbines Incorporated | Turbine cooling apparatus |
US9314838B2 (en) | 2012-09-28 | 2016-04-19 | Solar Turbines Incorporated | Method of manufacturing a cooled turbine blade with dense cooling fin array |
US9206695B2 (en) | 2012-09-28 | 2015-12-08 | Solar Turbines Incorporated | Cooled turbine blade with trailing edge flow metering |
US9228439B2 (en) | 2012-09-28 | 2016-01-05 | Solar Turbines Incorporated | Cooled turbine blade with leading edge flow redirection and diffusion |
WO2014120565A1 (en) | 2013-02-04 | 2014-08-07 | United Technologies Corporation | Bell mouth inlet for turbine blade |
US9638041B2 (en) | 2013-10-23 | 2017-05-02 | General Electric Company | Turbine bucket having non-axisymmetric base contour |
US9528379B2 (en) | 2013-10-23 | 2016-12-27 | General Electric Company | Turbine bucket having serpentine core |
US9670784B2 (en) | 2013-10-23 | 2017-06-06 | General Electric Company | Turbine bucket base having serpentine cooling passage with leading edge cooling |
US9797258B2 (en) * | 2013-10-23 | 2017-10-24 | General Electric Company | Turbine bucket including cooling passage with turn |
US9551226B2 (en) | 2013-10-23 | 2017-01-24 | General Electric Company | Turbine bucket with endwall contour and airfoil profile |
US9551229B2 (en) | 2013-12-26 | 2017-01-24 | Siemens Aktiengesellschaft | Turbine airfoil with an internal cooling system having trip strips with reduced pressure drop |
EP3059394B1 (en) * | 2015-02-18 | 2019-10-30 | Ansaldo Energia Switzerland AG | Turbine blade and set of turbine blades |
US10107108B2 (en) | 2015-04-29 | 2018-10-23 | General Electric Company | Rotor blade having a flared tip |
DE102015112643A1 (en) * | 2015-07-31 | 2017-02-02 | Wobben Properties Gmbh | Wind turbine rotor blade |
WO2017103650A1 (en) * | 2015-12-15 | 2017-06-22 | Technion Research & Development Foundation Limited | Acoustic resonance excited heat exchange |
US10260355B2 (en) | 2016-03-07 | 2019-04-16 | Honeywell International Inc. | Diverging-converging cooling passage for a turbine blade |
EP3673152A1 (en) * | 2017-08-24 | 2020-07-01 | Siemens Aktiengesellschaft | Turbine rotor airfoil and corresponding method for reducing pressure loss in a cavity within a blade |
US10837291B2 (en) * | 2017-11-17 | 2020-11-17 | General Electric Company | Turbine engine with component having a cooled tip |
US10830059B2 (en) * | 2017-12-13 | 2020-11-10 | Solar Turbines Incorporated | Turbine blade cooling system with tip flag transition |
US10655476B2 (en) * | 2017-12-14 | 2020-05-19 | Honeywell International Inc. | Gas turbine engines with airfoils having improved dust tolerance |
EP3862537A1 (en) * | 2020-02-10 | 2021-08-11 | General Electric Company Polska sp. z o.o. | Cooled turbine nozzle and nozzle segment |
Family Cites Families (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE379162C (en) * | 1923-08-17 | Hans Holzwarth Dipl Ing | Blading for explosion turbines | |
GB1202895A (en) * | 1967-03-01 | 1970-08-19 | John Alan Charles Kentfield | Improvements in or relating to fluid rectifiers |
BE794194A (en) * | 1972-01-18 | 1973-07-18 | Bbc Sulzer Turbomaschinen | COOLED MOBILE BLADE FOR GAS TURBINES |
US4162136A (en) * | 1974-04-05 | 1979-07-24 | Rolls-Royce Limited | Cooled blade for a gas turbine engine |
CA1087527A (en) * | 1977-02-10 | 1980-10-14 | George A. Durgin | Cooled gas turbine blade |
US4292008A (en) * | 1977-09-09 | 1981-09-29 | International Harvester Company | Gas turbine cooling systems |
US4236870A (en) * | 1977-12-27 | 1980-12-02 | United Technologies Corporation | Turbine blade |
US4278400A (en) * | 1978-09-05 | 1981-07-14 | United Technologies Corporation | Coolable rotor blade |
FR2468727A1 (en) * | 1979-10-26 | 1981-05-08 | Snecma | IMPROVEMENT TO COOLED TURBINE AUBES |
US4416585A (en) * | 1980-01-17 | 1983-11-22 | Pratt & Whitney Aircraft Of Canada Limited | Blade cooling for gas turbine engine |
US4474532A (en) * | 1981-12-28 | 1984-10-02 | United Technologies Corporation | Coolable airfoil for a rotary machine |
US4515526A (en) * | 1981-12-28 | 1985-05-07 | United Technologies Corporation | Coolable airfoil for a rotary machine |
US4775296A (en) * | 1981-12-28 | 1988-10-04 | United Technologies Corporation | Coolable airfoil for a rotary machine |
GB2119028B (en) * | 1982-04-27 | 1985-02-27 | Rolls Royce | Aerofoil for a gas turbine engine |
-
1984
- 1984-10-04 GB GB08425092A patent/GB2165315B/en not_active Expired
-
1985
- 1985-08-27 US US06/769,702 patent/US4604031A/en not_active Expired - Lifetime
- 1985-09-20 JP JP60208543A patent/JPS6189901A/en active Pending
- 1985-09-26 FR FR8514256A patent/FR2571428B1/en not_active Expired - Lifetime
- 1985-09-30 DE DE3534905A patent/DE3534905C2/en not_active Expired - Lifetime
Cited By (20)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0465004A2 (en) * | 1990-07-03 | 1992-01-08 | ROLLS-ROYCE plc | Cooled aerofoil blade |
EP0465004A3 (en) * | 1990-07-03 | 1992-12-02 | Rolls-Royce Plc | Cooled aerofoil blade |
US5156526A (en) * | 1990-12-18 | 1992-10-20 | General Electric Company | Rotation enhanced rotor blade cooling using a single row of coolant passageways |
EP0672821A1 (en) * | 1994-02-09 | 1995-09-20 | ROLLS-ROYCE plc | Air cooled gas turbine aerofoil |
US5507621A (en) * | 1995-01-30 | 1996-04-16 | Rolls-Royce Plc | Cooling air cooled gas turbine aerofoil |
US5934874A (en) * | 1996-08-23 | 1999-08-10 | Asea Brown Boveri Ag | Coolable blade |
US6241468B1 (en) | 1998-10-06 | 2001-06-05 | Rolls-Royce Plc | Coolant passages for gas turbine components |
GB2349920A (en) * | 1999-05-10 | 2000-11-15 | Abb Alstom Power Ch Ag | Cooling arrangement for turbine blade |
US6347923B1 (en) | 1999-05-10 | 2002-02-19 | Alstom (Switzerland) Ltd | Coolable blade for a gas turbine |
GB2349920B (en) * | 1999-05-10 | 2003-06-25 | Abb Alstom Power Ch Ag | Coolable blade for a gas turbine |
US7137784B2 (en) | 2001-12-10 | 2006-11-21 | Alstom Technology Ltd | Thermally loaded component |
DE102004011151B4 (en) * | 2003-03-19 | 2015-11-26 | Alstom Technology Ltd. | turbine blade |
EP1793089A2 (en) | 2005-11-30 | 2007-06-06 | General Electric Company | Method and apparatus for reducing axial compressor blade tip flow |
EP1793089A3 (en) * | 2005-11-30 | 2007-10-24 | General Electric Company | Method and apparatus for reducing axial compressor blade tip flow |
EP2489838A3 (en) * | 2011-02-17 | 2017-12-27 | Rolls-Royce plc | Cooled component for the turbine of a gas turbine engine |
EP2692991A1 (en) * | 2012-08-01 | 2014-02-05 | Siemens Aktiengesellschaft | Cooling of turbine blades or vanes |
EP2832953A1 (en) * | 2013-07-29 | 2015-02-04 | Siemens Aktiengesellschaft | Turbine blade |
WO2015014613A1 (en) * | 2013-07-29 | 2015-02-05 | Siemens Aktiengesellschaft | Turbine blade |
CN105431614B (en) * | 2013-07-29 | 2017-09-15 | 西门子股份公司 | Turbo blade |
WO2022238051A1 (en) * | 2021-05-11 | 2022-11-17 | Siemens Energy Global GmbH & Co. KG | Rotor-blade tip including cooling configuration |
Also Published As
Publication number | Publication date |
---|---|
US4604031A (en) | 1986-08-05 |
DE3534905C2 (en) | 2000-10-12 |
FR2571428A1 (en) | 1986-04-11 |
FR2571428B1 (en) | 1993-03-12 |
GB2165315B (en) | 1987-12-31 |
DE3534905A1 (en) | 1986-04-10 |
JPS6189901A (en) | 1986-05-08 |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
US4604031A (en) | Hollow fluid cooled turbine blades | |
EP0716217B1 (en) | Trailing edge ejection slots for film cooled turbine blade | |
US6264428B1 (en) | Cooled aerofoil for a gas turbine engine | |
EP0942150B1 (en) | A stator vane assembly for a turbomachine | |
US7371046B2 (en) | Turbine airfoil with variable and compound fillet | |
EP1091092B1 (en) | Coolable gas turbine airfoil | |
US5927946A (en) | Turbine blade having recuperative trailing edge tip cooling | |
JP4801513B2 (en) | Cooling circuit for moving wing of turbomachine | |
EP1126135B1 (en) | Ceramic turbine airfoils with cooled trailing edge blocks | |
EP1041247B1 (en) | Gas turbine airfoil comprising an open cooling circuit | |
JP4659206B2 (en) | Turbine nozzle with graded film cooling | |
JP4137507B2 (en) | Apparatus and method for airfoil film cooling | |
JP4785507B2 (en) | Turbine nozzle with bull nose step | |
JP4152184B2 (en) | Turbine platform with descending stage | |
CA2491619C (en) | Bifurcated outlet guide vanes | |
JP4311919B2 (en) | Turbine airfoils for gas turbine engines | |
JP4245873B2 (en) | Turbine airfoils for gas turbine engines | |
CA2944474A1 (en) | Turbine blade | |
GB2110767A (en) | A shrouded rotor for a gas turbine engine | |
US10724391B2 (en) | Engine component with flow enhancer | |
JPH11193701A (en) | Turbine wing | |
US6544001B2 (en) | Gas turbine engine system | |
JP2818266B2 (en) | Gas turbine cooling blade | |
US20220220855A1 (en) | Airfoils for gas turbine engines | |
GB2112869A (en) | Cooled airfoil |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
PE20 | Patent expired after termination of 20 years |