GB2119861A - Cooling high pressure turbine blade tips in a gas turbine engine - Google Patents
Cooling high pressure turbine blade tips in a gas turbine engine Download PDFInfo
- Publication number
- GB2119861A GB2119861A GB08234168A GB8234168A GB2119861A GB 2119861 A GB2119861 A GB 2119861A GB 08234168 A GB08234168 A GB 08234168A GB 8234168 A GB8234168 A GB 8234168A GB 2119861 A GB2119861 A GB 2119861A
- Authority
- GB
- United Kingdom
- Prior art keywords
- turbine
- combustor
- wall
- cooling air
- cooling
- 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
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
-
- 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
- F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
- F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
- F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
- F23R3/04—Air inlet arrangements
- F23R3/06—Arrangement of apertures along the flame tube
-
- 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
Abstract
Inlet air holes 36 are provided in a radially outer wall section of the combustor 16, just upstream of a turbine nozzle. Cooling air flows through these inlet air holes, into annulus regions protected from combustion gases, and then downstream along a radially outer wall 34 of the turbine 18. The cooling air forms a film that cools the turbine blade tips in a localized manner that adds to total engine power output. <IMAGE>
Description
SPECIFICATION
Cool tip combustor
This invention relates to means for directing cooling air to critical parts of hot section turbine blades in gas turbine engines.
In the course of gas turbine engine development, tremendous effort has been directed at raising the internal operating temperatures of such engines to improve thermodynamic efficiency. As turbine inlet temperatures have been increased in pursuit of this goal, it has become necessary to provide cooling air to hot section turbine blades and vanes in order to limit temperatures of those components to levels that can be accommodated by the blade and vane materials. The air that is used for this cooling function is usually compressed to pressures that meet or exceed the gas pressures inside the turbine section. Because the air has undergone the work necessary for compression, this cooling air must be used as efficiently as possible to limit the power required by the engine's compressor section in order to compress that air.To limit the amount of cooling air used, intricate cooling air flowpaths and passages are utilized that are intended to use the cooling air in a highly efficient manner.
In smaller airflow size engines, blade cooling configurations are generally restricted to fairly simple designs because of small dimensions and limitations of current manufacturing technologies. The implication is a typical smaller engine turbine blade or vane cannot be provided with the highly complex, internal air cooling passage configuration typically used today in larger gas turbine engines.
One particular problem with smaller engines is that tip sections of turbine blades are extremely difficult to cool efficiently. The cooling air used to internally cool turbine blade tips has increased its temperature by thermal pickup in the lower portion of the blade rendering it less effective for cooling purposes. In downstream sections of the turbine blade tips some of the cooling air has been bled out of the trailing edge cooling holes before it reaches the blade tip region, thereby reducing cooling air velocity and, consequently, its cooling effectiveness.
Adding to these difficulties of cooling small turbine blades, the downstream trailing edge of the blade tip region is usually very thin for aerodynamic performance reasons, which limits the ability to duct cooling air into this region.
As a result of these inherent limitations, design cycle temperatures of these small engines are restricted and engine performance is thereby limited. Further, the turbine blade tips often become a life-limiting engine component problem area. As the turbine tips deteriorate, due to oxidation and corrosion accumulating during engine use, the engine performance drops below minimum acceptable levels. The engine must then be removed from the aircraft and the turbine section refurbished. Maintenance and overhaul of the turbine section to correct deteriorated blade tips is both expensive and time consuming.
It is, therefore, an object of the present invention
to provide a means for cooling tips of turbine blades
in turbine sections of gas turbine engines with a
system that can be utilized in relatively small engine
configurations.
Another object of the present invention is to
provide a source of cooling air that can be directed
specifically to turbine blade tips in a turbine section
of a small gas turbine engine.
It is another object of the present invention to
provide a film of cooling air along a radially outer
most wall of a turbine section of a small gas turbine
engine for the purpose of cooling turbine blade tips with a limited amount of cooling air.
These and other objects will become more readily
apparent upon reference to the following description
in conjunction with the appended drawings.
Briefly, in accordance with one embodiment of the
present invention, means are provided for introduc
ing cooling air into a turbine section of a gas turbine engine in the region of tip sections of turbine blades.
The source of this cooling air is compressor discharge air that has bypassed the combustor. This compressor discharge air is introduced at an aft section of the combustor through inlet air holes just upstream of the turbine section. The air is introduced along a radially ouer section of the combustor only.
The cooling air flows initially into annulus regions within the combustor that are protected from hot combustion gases. From these annulus regions, the air flows downstream in the combustor and forms a thick film that blankets the combustor wall. Because it is introduced at the downstream section of the combustor, there is no combution of this cooling air, and it enters the turbine section at close to the same temperature as when the cooling air entered the combustor section. This temperature is much lower than the hot gases which have just undergone combustion. This thick, low-temperature cooling air film flows into the turbine section along a radially outer wall of the turbine flow path. The cooling air film provides a reiatively cooler gas flow along the tips of the turbine blades roating in the turbine section.It is primarily the tips of the turbine blades only that are thus cooled, and this limits the amount of cooling air employed.
Figure 1 is a schematic cross-sectional illustration of a central section of a gas turbine engine.
Figure 2 is a schematic cross-sectional illustration of a combustor and high pressure turbine section of a gas turbine engine with the present invention embodied therein.
Figure 3 is a cross-sectional illustration of a downstream portion of a combustorwall with one embodiment of part of the present invention incorporated therein.
Figure 4 is a graphical representation of test results of turbine blade temperatures.
Referring now to Figure 1, a central section of a typical gas turbine engine 10 is shown that involves substantial turbomachinery that rotates about an engine centerline 12. Components of this turbomachinery include, in serial flow relationship, a compressor 14, a combustor 16, a high-pressure turbine section 18, and a low-pressure turbine section 20. In conventional operation, inlet air is directed into and pressurized by the compressor 14 from which the air is discharged through a diffuser 22. A major portion of this compressor discharge air is then passed into the combustor 16 where it is mixed with fuel and vaporized to form highpressure, high-temperature combustion gases which flow downstream into the high-pressure turbine 18. The high-pressure gases cause turbine blades 24 in the high-pressure turbine 18 to rotate at high velocities thereby providing mechanical power.
These high-temperatures, high-pressure gases then continue to flow downstream into the low-pressure turbine 20 where they cause low-pressure turbine blades 26 to rotate thereby providing additional mechanical power. From the low-pressure turbine 20, the gases are discharged downstream so as to pass out of the engine 10.
A portion of the air discharged from the compressor 14 that passes through the diffuser 22 is circulated to cool a variety of hot parts of the engine 10. Some of that air used for cooling flows to the region of the combustor 16 and surrounds the combustor walls. In some engines, small cooling holes are provided in the combustor walls so that cooling air can enter the combustor to cool interior combustor surfaces.
Other portions of the cooling air are directed inter wally to hot temperature parts inside the highpressure turbine 18. A part of this air used to cool the high-pressure turbine is directed into the interior of a high-pressure turbine nozzle 28 so as to provide a internal cooling function by impingement and diffusion processes. Another part of the compressor discharge air is directed along other paths to cool interior regions of the turbine blades 24 of the high-pressure turbine 18. These cooling flowpaths are generally represented by the dark arrows in
Figure 1.
It is well-known in the art that during high-power, high-temperature operating conditions a substantial amount of cooling air is needed for these cooling processes. Because of the limitations of size and manufacturing processes, it is particularly difficult to cool tip sections 30 of the turbine blades 24. These tip sections 30 are usually very thin for aerodynamic performance reasons and this limits the ability to efficiently duct cooling air into the tip sections. In addition, the thin sections deteriorate due to oxidation and corrosion causing substantial problems in engine performance.
A prior art solution to the problem of cooling turine blade tips is a ducting of a small portion of the compressor discharge cooling air into the highpressure turbine 18 at a inlet location 32 just downstream of the turbine nozzle 28. Cooling air ducted in this manner would bypass the combustor and flow into the high-pressure turbine 18just upstream of the turbine blades 24. Studies have indicated that this proposal reduces turbine tip temperatures, but this approach also has a negative impact on engine performance, both in terms of thrust and fuel consumption. The detrimental effect on engine performance is caused because the cooling air enters the gas flow stream behind the first-stage turbine nozzle 28 and is, therefore, chargeable to the engine's thermodynamic cycle.As a result, the amount of air burned compared to the allowable tubine rotor inlet temperature level is reduced and engine performance decreases.
Referring now to Figure 2, a portion of a gas turbine engine 11 is shown that is generally similar to part of the engine shown in Figure 1, but, this time, incorporating an embodiment ofthe present invention. Again, as explained in relation to the engine shown in Figure 1, a portion of the cooling air discharged from the compressor does not enter the combustor 16, but, instead, flows downstream around the combustor as indicated by dark arrows in
Figure 2. This cooling air does not undergo the mixing and combustion processes occurring during engine operation inside the combustor 16. Because the air does not undergo combustion, it remains relatively cool and serves as a source of highpressure cooling air that can be utilized in the high-pressure turbine sections of the engine.Any cooling air used in the high-pressure turbine section must be at high pressure because the internal gases flowing through the high-pressure turbine area, as the name suggests, are at very high pressure. The cooling air introduced into the high-pressure turbine must be even higher in pessure than those gases flowing through the turbine so that the cooling air will be caused by its own pressure forces to flow into the turbine blades and vanes and from these into the combustion gas flow passage of the turbine section.
If the cooling air that was used for cooling in this region were lower in pressure than the combustion gases flowing through the turbine section, pressure forces would not permit the cooling air to flow from interior regions of the turbine blades and vanes out into the combustion gas flow passage.
Realizing that this compressor discharge air is the best available source of cooling air flow that can be utilized for cooling turbine blades, the problem becomes a matter of utilizing this air in the best manner possible to cool the turbine blades and the turbine blade tips. It is extremely important that the volume of cooling air used be kept as low as possible because the air has undergone a great deal of work in the compressor section in order to compress that air, and it is desirable to minimize the amount of air used in order to increase the efficiency of the engine.
It is also desirable to introduce this highly compressed cooling air in a location that permits highly pressurized air to be expanded and directed at the tubine blades in a manner such that the cooling air will not only cool the tips of the turbine blades but will also add to the effective gas forces that cause the turbine blades 24to rotate, thereby increasing the total power produced by the engine 10.
If cooling air is introduced at an inlet location 32 immediately downstream of the turbine nozzle 28, the air will tend to cool the turbine blade tips 20.
However, because the air has not been expanded and directed by the turbine nozzle 28, it will not be useful for providing appropriate gas forces for causing the turbine blades 24 to rotate.
The present invention comprises means for introducing cooling air forward or upstream of the first-stage turbine nozzle 28 so that there is no associated engine performance penalty. One embo diment of this means is shown in Figue 2, and a part of the invention is shown in larger scale in Figure 3.
Referring initially to Figure 2, a portion of the compressor discharge air that is flowing outside of the combustor 16 is directed into combustor inlet air holes 36 at a location just upstream of the turbine nozzle 28. The air is introduced at a location just upstream of the turbine nozzle 28, partly to prevent that cooling air from undergoing the normal combustion processes inside the tubine 16, and also to lessen heating of the cooling air from prolonged exposure to the hot combustion gases. If this cooling airwereto undergo combustion, it would climb dramatically in temperature and be rendered relatively useless for the purpose of cooling tips 30 of turbine blades.
Referring now to Figure 3, the inlet air holes 36 through which the cooling air is directed into a downstream section of the combustor 16 are shown in greater detail. A portion of a radially outer wall section 38 of the combustor 16 is shown in Figure 3.
This portion of the combustor wall section 38 is located just upstream of the turbine nozzle 28 (not shown). In the cross-sectional view shown, three inlet air holes 36 can be seen and their relative configuration can be appreciated. It should first be noted that the downstream portion of the combustor wall section 38 is actually double-walled. An outer wall section 40 connects to the turbine nozzle in a standard manner as would be the normal practice in many gas turbine engines. As inner combustor wall section 42 is provided and is protected from hot combustion gases at its upstream end by a flange 44.
At its downstream end, the inner wall section 42 extends almost to the turbine nozzle inlet. Cooling air from the compressor discharge is bled into annulus regions 46 that are open in a downstream direction and are generally protected from the combustion occurring inside the combustor 16.
Because the cooling air is bled into these protected annulus regions 46, the cooling air does not undergo combustion, and the air enters the turbine nozzle at substantially compressor discharge temperature, thereby forming a thick, low-temperature film along a radially outer wall of the turbine flow path.
As stated earlier, there are three inlet air holes 36 visible in Figure 3. Each of the holes 36, as shown, repesents one of a row of holes that extend around the entire circumference of the radially outer wall section 38 of the combustor 16. The total number of inlet air holes 36 could vary widely as could their general configuration.
A row of upstream inlet air holes 48 is provided to bleed cooling air into the annulus region between the flange 44 and the inner wall 42. A row of intermediate inlet air holes 50 is provided to bleed additional cooling air into the annulus region between the inner wall 42 and the outer wall 40. Finally, a row of downstream inlet air holes 52 is provided to direct additional cooling air into the annulus between the inner wall 42 and the outer wall 40. It can be readily appreciated by those skilled in the art that the size of these inlet air holes 36 can be varied for the purpose of introducing varying amounts of cooling air. To serve as a guide, in one embodiment of the present invention, these holes are varied from .026 inches (.066 cm.) in diameter to .035 inches (.089 cm.) in diameter.These dimensions, however, are simply a guideline and smaller or larger diameter holes could easily be utilized without departing from the scope of the present invention. Additionally, widely varying inlet air hole configurations would also be within the scope of the invention.
Referring again to Figure 2, small black arrows are shown entering the combustor 16 emanating from the annulus regions 46 within the combustor 16 and flowing downstream along the radially outer turbine wall 34, past the turbine nozzle 28 to the region of the turbine blade tips 30. This air tends to flow as a low temperature film in a manner that is ideal for cooling the turbine blade tips 30 without using excessive amounts of compressor discharge air thereby accomplishing the purpose of the present invention.
Referring now to Figure 4, a comparison of test results is shown that graphically represents turbine blade temperatures in a typical gas turbine engine and, additionally, represents turbine blade temperatures in a second gas turbine engine incorporating the present invention. The X (horizontal) coordinate in Figure 4 is marked off in degrees Fahrenheit. The
Y (vertical) coordinate in Figure 4 is a dimensionless representation of turbine blade height, beginning at a root of the turbine blade and ending at the tip of the turbine blade. The lines shown on the graph of
Figure 4 designated 54 represent turbine blade temperatures in two typical gas turbine engines, generally having an engine configuration similar two that shown in Figure 2 but without incorporating the present invention.The line designated 56 in Figure 4 represents turbine blade temperature, again within an engine having generally the same configuration as shown in Figure 2, but this time incorporating the present invention. It can be readily appreciated that tubine tip temperatures are significantly decreased in the engine incorporating the present invention.
Because of this temperature reduction at the turbine tip, the present invention has been commonly referred to as a "cool tip" engine. It is important to note that this reduction in turbine tip temperatures is achieved generally without utilizing excessive amounts of compressor discharge air and in a manner that directs the cooling effect at the turbine blade tips. It is desirable to obtain this "cool tip" effect in a localized manner as shown graphically in
Figure 4. .
Although the present invention has been described in terms of its preferred embodiment, it will be apparent to those skilled in the art that changes and modifications thereof may be made without departing from the scope of the appended claims which define the present invention.
Claims (7)
1. In a gas turbine engine having a compressor, a combustor, a turbine section with a turbine nozzle and turbine blades, all in serial flow relationship and disposed radially about an engine centerline, means for cooling tips of said turbine blades comprising:
means for directing high-pressure cooling air from said compressor into a downstream, radially outer wall section of said combustor in a manner that creates a cooling air film that extends downstream along a radially outer wall of said turbine section to cool said turbine blade tips.
2. The gas turbine engine recited in Claim 1 wherein said means includes a plurality of inlet air holes in a radially outer wall section of said combustor at a location just upstream of the turbine nozzle.
3. The gas turbine engine recited in Claim 2 wherein said inlet air holes direct said cooling air into annulus regions inside said combustorthereby preventing combustion of said cooling air in said combustor.
4. The gas turbine engine recited in Claim 3 wherein said cooling air is directed through some of said inlet air holes into an annulus formed between a flange and an inner wall of said combustor and, additionally, through other of said inlet air holes into another annulus formed between said inner wall and an outer wall of said combustor.
5. The gas turbine engine recited in Claim 4 wherein the cooling air that is directed into the annulus formed between said flange and said inner wall is directed through an upstream row of inlet air holes extending circumferentially around said combustor and the cooling air that is directed into the annulus formed between said inner wall and said outer wall is directed through an intermediate row of inlet air holes extending circumferentially around said combustor and, additionally, a downstream row of inlet air holes extending circumferentially around said combustor.
6. In a gas turbine engine having a compressor, a combustor, a turbine section with a turbine nozzle and turbine blades, all in serial flow relationship and disposed about an engine centerline, means for providing a film of cooling air along a radially outer wall of said turbine section for cooling tips of said turbine blades, said means comprising: :
an upstream row of inlet air holes extending circumferentially around a radially outer wall section of said combustor, wherein compressor discharge air is directed into an annulus between a flange and an inner wall of said combustor;
an intermediate row of inlet air holes and a downstream row of inlet air holes, both rows extending circumferentially around a radially outer wall section of said combustor, wherein compressor discharge air is directed from said intermediate row and said downstream row into another annulus between said inner wall and an outer wall of said combustor;
said annuli being open in a downstream direction such that said cooling air is directed into said combustor as a film of cooling air that extends downstream along a radially outer wall of said turbine section thereby cooling tips of said turbine blades.
7. A gas turbine as illustrated in Figures 2 and 3.
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US06/374,355 USH903H (en) | 1982-05-03 | 1982-05-03 | Cool tip combustor |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| GB2119861A true GB2119861A (en) | 1983-11-23 |
| GB2119861B GB2119861B (en) | 1985-07-31 |
Family
ID=23476435
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| GB08234168A Expired GB2119861B (en) | 1982-05-03 | 1982-12-01 | Cooling high pressure turbine blade tips in a gas turbine engine |
Country Status (8)
| Country | Link |
|---|---|
| US (1) | USH903H (en) |
| JP (1) | JPS58195027A (en) |
| CA (1) | CA1194803A (en) |
| DE (1) | DE3248439A1 (en) |
| FR (1) | FR2526083B1 (en) |
| GB (1) | GB2119861B (en) |
| IL (1) | IL67317A (en) |
| IT (1) | IT1153896B (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP0178242A1 (en) * | 1984-10-11 | 1986-04-16 | United Technologies Corporation | Cooling scheme for combustor vane interface |
| GB2184167A (en) * | 1983-02-22 | 1987-06-17 | Gen Electric | Cooling gas turbine engine components |
| US4785623A (en) * | 1987-12-09 | 1988-11-22 | United Technologies Corporation | Combustor seal and support |
| WO2009144300A1 (en) * | 2008-05-29 | 2009-12-03 | Snecma | Assembly including a turbine disc for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disc of such an assembly |
| EP2249003A1 (en) * | 2008-02-27 | 2010-11-10 | Mitsubishi Heavy Industries, Ltd. | Gas turbine and method of opening casing of gas turbine |
| EP2837769A1 (en) * | 2013-08-13 | 2015-02-18 | Alstom Technology Ltd | Rotor shaft for a turbomachine |
Families Citing this family (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5188510A (en) * | 1990-11-21 | 1993-02-23 | Thomas R. Norris | Method and apparatus for enhancing gas turbo machinery flow |
| US5178660A (en) * | 1991-06-26 | 1993-01-12 | Libbey-Owens-Ford Co. | Apparatus for bending glass sheets |
| GB9304994D0 (en) * | 1993-03-11 | 1993-04-28 | Rolls Royce Plc | Improvements in or relating to gas turbine engines |
| US5749218A (en) * | 1993-12-17 | 1998-05-12 | General Electric Co. | Wear reduction kit for gas turbine combustors |
| US6494678B1 (en) | 2001-05-31 | 2002-12-17 | General Electric Company | Film cooled blade tip |
| FR2920525B1 (en) * | 2007-08-31 | 2014-06-13 | Snecma | SEPARATOR FOR SUPPLYING THE COOLING AIR OF A TURBINE |
| US10954796B2 (en) * | 2018-08-13 | 2021-03-23 | Raytheon Technologies Corporation | Rotor bore conditioning for a gas turbine engine |
| US11371700B2 (en) * | 2020-07-15 | 2022-06-28 | Raytheon Technologies Corporation | Deflector for conduit inlet within a combustor section plenum |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB776847A (en) * | 1954-05-20 | 1957-06-12 | Sulzer Ag | Gas-cooled gas turbine |
| GB1193587A (en) * | 1968-04-09 | 1970-06-03 | Rolls Royce | Nozzle Guide Vanes for Gas Turbine Engines. |
| GB1516757A (en) * | 1975-10-14 | 1978-07-05 | United Technologies Corp | Turbomachinery vane or blade with cooled platforms |
Family Cites Families (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2541171A (en) * | 1947-01-25 | 1951-02-13 | Kellogg M W Co | Air inlet structure for combustion chambers |
| DE1270889B (en) * | 1964-05-13 | 1968-06-20 | Rolls Royce | Cooling device for combustion chambers of gas turbine engines |
| US3418808A (en) * | 1966-07-05 | 1968-12-31 | Rich David | Gas turbine engines |
| CA982828A (en) * | 1972-06-01 | 1976-02-03 | General Electric Company | Combustor casing cooling structure |
| US3965066A (en) * | 1974-03-15 | 1976-06-22 | General Electric Company | Combustor-turbine nozzle interconnection |
| GB1550368A (en) * | 1975-07-16 | 1979-08-15 | Rolls Royce | Laminated materials |
| GB1524956A (en) * | 1975-10-30 | 1978-09-13 | Rolls Royce | Gas tubine engine |
| DE2810240C2 (en) * | 1978-03-09 | 1985-09-26 | MTU Motoren- und Turbinen-Union München GmbH, 8000 München | Adjustable grille for turbines with axial flow, in particular high-pressure turbines for gas turbine engines |
| FR2490728A1 (en) * | 1980-09-25 | 1982-03-26 | Snecma | AIR FILM COOLING DEVICE FOR FLAME TUBE OF GAS TURBINE ENGINE |
-
1982
- 1982-05-03 US US06/374,355 patent/USH903H/en not_active Abandoned
- 1982-11-05 CA CA000414954A patent/CA1194803A/en not_active Expired
- 1982-11-21 IL IL67317A patent/IL67317A/en unknown
- 1982-12-01 GB GB08234168A patent/GB2119861B/en not_active Expired
- 1982-12-24 IT IT24980/82A patent/IT1153896B/en active
- 1982-12-28 FR FR8221913A patent/FR2526083B1/en not_active Expired
- 1982-12-28 JP JP57227827A patent/JPS58195027A/en active Granted
- 1982-12-29 DE DE19823248439 patent/DE3248439A1/en not_active Ceased
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB776847A (en) * | 1954-05-20 | 1957-06-12 | Sulzer Ag | Gas-cooled gas turbine |
| GB1193587A (en) * | 1968-04-09 | 1970-06-03 | Rolls Royce | Nozzle Guide Vanes for Gas Turbine Engines. |
| GB1516757A (en) * | 1975-10-14 | 1978-07-05 | United Technologies Corp | Turbomachinery vane or blade with cooled platforms |
Cited By (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2184167A (en) * | 1983-02-22 | 1987-06-17 | Gen Electric | Cooling gas turbine engine components |
| EP0178242A1 (en) * | 1984-10-11 | 1986-04-16 | United Technologies Corporation | Cooling scheme for combustor vane interface |
| US4785623A (en) * | 1987-12-09 | 1988-11-22 | United Technologies Corporation | Combustor seal and support |
| EP2249003A1 (en) * | 2008-02-27 | 2010-11-10 | Mitsubishi Heavy Industries, Ltd. | Gas turbine and method of opening casing of gas turbine |
| EP2249003A4 (en) * | 2008-02-27 | 2015-03-18 | Mitsubishi Heavy Ind Ltd | Gas turbine and method of opening casing of gas turbine |
| US9080464B2 (en) | 2008-02-27 | 2015-07-14 | Mitsubishi Hitachi Power Systems, Ltd. | Gas turbine and method of opening chamber of gas turbine |
| WO2009144300A1 (en) * | 2008-05-29 | 2009-12-03 | Snecma | Assembly including a turbine disc for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disc of such an assembly |
| FR2931873A1 (en) * | 2008-05-29 | 2009-12-04 | Snecma | A TURBINE DISK ASSEMBLY OF A GAS TURBINE ENGINE AND A BEARING BRIDGE SUPPORT CIRCUIT, COOLING CIRCUIT OF A TURBINE DISK OF SUCH AN ASSEMBLY. |
| US8899913B2 (en) | 2008-05-29 | 2014-12-02 | Snecma | Assembly including a turbine disk for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disk of such an assembly |
| EP2837769A1 (en) * | 2013-08-13 | 2015-02-18 | Alstom Technology Ltd | Rotor shaft for a turbomachine |
| US11105205B2 (en) | 2013-08-13 | 2021-08-31 | Ansaldo Energia Switzerland AG | Rotor shaft for a turbomachine |
Also Published As
| Publication number | Publication date |
|---|---|
| USH903H (en) | 1991-04-02 |
| FR2526083A1 (en) | 1983-11-04 |
| IT1153896B (en) | 1987-01-21 |
| GB2119861B (en) | 1985-07-31 |
| IT8224980A1 (en) | 1984-06-24 |
| DE3248439A1 (en) | 1983-11-03 |
| IL67317A0 (en) | 1983-03-31 |
| JPS58195027A (en) | 1983-11-14 |
| CA1194803A (en) | 1985-10-08 |
| JPH0415378B2 (en) | 1992-03-17 |
| IL67317A (en) | 1989-10-31 |
| FR2526083B1 (en) | 1987-11-27 |
| IT8224980A0 (en) | 1982-12-24 |
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