CN113266436A - Channel structure for cooling inside of gas turbine stationary blade and gas turbine stationary blade - Google Patents
Channel structure for cooling inside of gas turbine stationary blade and gas turbine stationary blade Download PDFInfo
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- CN113266436A CN113266436A CN202110528139.3A CN202110528139A CN113266436A CN 113266436 A CN113266436 A CN 113266436A CN 202110528139 A CN202110528139 A CN 202110528139A CN 113266436 A CN113266436 A CN 113266436A
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- gas turbine
- ribs
- channel structure
- channel
- rib
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
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- 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
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2212—Improvement of heat transfer by creating turbulence
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
- F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention belongs to the field of gas turbines, and discloses a channel structure for cooling the interior of a gas turbine stationary blade and the gas turbine stationary blade, which comprise a plurality of straight channels and a plurality of bent channels; a plurality of straight channels are arranged in parallel, and two adjacent straight channels are connected through a bent channel; defining the wall surface positioned at the upper side of the main flow direction in the straight channel as a suction surface; the wall surface positioned at the lower side of the main flow direction is a pressure surface; the suction surface and the pressure surface are both provided with a plurality of inclined ribs; a plurality of transverse ribs are arranged on the wall surface connected with the downstream ends of the inclined ribs in the straight channel, and the downstream ends of the inclined ribs are the downstream ends of the inclined ribs in the projection of the inclined ribs in the main flow direction; one end of the transverse rib is connected with the inclined rib on the suction surface, and the other end of the transverse rib is connected with the inclined rib on the pressure surface. The heat transfer of the low Reynolds number area at the downstream of the fins can be effectively improved, so that the temperature distribution is more uniform, the local hot spots of the stationary blades of the gas turbine are eliminated, the thermal stress is improved, the stability of the stationary blades of the gas turbine is improved, and the service life is prolonged.
Description
Technical Field
The invention belongs to the field of gas turbines, and relates to a channel structure for cooling the interior of a stationary blade of a gas turbine and the stationary blade of the gas turbine.
Background
The gas turbine is power equipment with high cycle efficiency, compact structure, small volume, light weight and capability of generating larger output power, is widely applied to important industrial fields of aviation power, ship propulsion, land power generation and the like, and has an important position in the fields of national defense construction and civil use.
Research shows that increasing the inlet temperature of the gas turbine is an effective method for increasing the cycle efficiency and the output power of the gas turbine. When the temperature of the gas inlet is increased by 100 ℃, the circulation specific work can be increased by 20-25%, and the fuel consumption is reduced by 6-7%. At present, with the continuous development of gas turbine technology, the turbine inlet temperature is continuously increased, for a heavy-duty gas turbine for land use, the turbine inlet temperature is developed from 1200-1400 ℃ of B/E/F stage to 1600 ℃ of G/H/J stage, the corresponding single cycle efficiency is increased from 32-34% to 42%, and the combined cycle efficiency is increased from 48-52% to 61%. In order to obtain higher heat-power conversion efficiency and output power, more advanced heavy-duty gas turbines with a turbine inlet temperature of 1700 ℃ are also under development.
However, as the temperature of the turbine inlet of the gas turbine is increased, the temperature far exceeds the allowable temperature of the metal material of the turbine blade, and under the erosion action of high-temperature gas, large thermal stress is generated, which not only reduces various mechanical properties of the metal blade, but also easily causes the blade to be burnt, thereby seriously affecting the economy and safety of the operation of the gas turbine. Therefore, gas turbine blade cooling technology is of great importance.
Disclosure of Invention
The invention aims to overcome the defects that the thermal stress of a gas turbine blade is increased along with the increase of the temperature of a turbine inlet of a gas turbine, so that the performance of the blade is reduced and the blade is even damaged in the prior art, and provides a channel structure for cooling the inside of a gas turbine static blade and the gas turbine static blade.
In order to achieve the purpose, the invention adopts the following technical scheme to realize the purpose:
in one aspect of the present invention, a channel structure for cooling the inside of a stationary blade of a gas turbine includes a plurality of straight channels and a plurality of curved channels;
a plurality of straight channels are arranged in parallel, and two adjacent straight channels are connected through a bent channel;
defining the wall surface positioned at the upper side of the main flow direction in the straight channel as a suction surface; the wall surface positioned at the lower side of the main flow direction is a pressure surface; the suction surface and the pressure surface are both provided with a plurality of inclined ribs;
a plurality of transverse ribs are arranged on the wall surface connected with the downstream ends of the inclined ribs in the straight channel, and the downstream ends of the inclined ribs are the downstream ends of the inclined ribs in the projection of the inclined ribs in the main flow direction; one end of the transverse rib is connected with the inclined rib on the suction surface, and the other end of the transverse rib is connected with the inclined rib on the pressure surface.
The channel structure for cooling the inside of the gas turbine static blade of the invention is further improved in that:
the included angle between the inclined ribs and the main flow direction is 30-150 degrees.
The width-to-height ratio of the straight channel is 0.25-2.
The height of the inclined ribs and the height of the transverse ribs are 1-3 mm.
The inclined ribs on the suction surface and the pressure surface are uniformly distributed, and the distance between every two adjacent inclined ribs is 10-20 mm.
The suction surface and the pressure surface are curved surfaces and are respectively consistent with the bending of the two outer wall surfaces of the stationary blade of the gas turbine.
The suction surface and the pressure surface are both provided with penetrating ribs along the main flow direction, and the penetrating ribs sequentially penetrate through a plurality of obliquely arranged ribs on the suction surface or the pressure surface.
And a plurality of transverse micro-teeth are arranged on the inclined ribs and the transverse ribs along the main flow direction.
The number of the straight channels is 3.
In a second aspect of the present invention, a gas turbine stationary blade is provided with the above-described passage structure for cooling inside the gas turbine stationary blade.
Compared with the prior art, the invention has the following beneficial effects:
the invention relates to a channel structure for cooling the interior of a stationary blade of a gas turbine, which is characterized in that a plurality of inclined ribs are arranged on a suction surface and a pressure surface, a plurality of transverse ribs are arranged on a wall surface connected with the downstream ends of the inclined ribs in a straight channel, and after a cooling working medium enters the channel, the top ends of the ribs of the inclined ribs are separated and attached to the area between the downstream ribs, so that longitudinal secondary flow is generated. The longitudinal secondary flow can develop along the direction of the fins and is converged with the main secondary flow at the downstream end of the fins, and because of the existence of the lateral ribs on the lateral wall surface, the boundary layer along the lateral wall surface can also generate separation and reattachment phenomena at the lateral ribs, so that the boundary layer on the lateral wall surface is continuously damaged and is difficult to develop and thicken, thereby strengthening the mixing of cooling working media in the channel, improving the temperature gradient of the cooling working media near the wall surface, further effectively improving the heat transfer of a low Reynolds number area at the downstream of the fins, ensuring more uniform temperature distribution, realizing better cooling effect, eliminating local hot spots on the stationary blades of the gas turbine, improving the thermal stress, improving the stability of the stationary blades of the gas turbine and prolonging the service life of the stationary blades of the gas turbine.
Furthermore, the suction surface and the pressure surface are curved surfaces and are respectively consistent with the bending of the two outer wall surfaces of the gas turbine stationary blade, and the suction surface and the pressure surface are set to be curved surfaces consistent with the bending of the two outer wall surfaces of the gas turbine stationary blade, so that the heat exchange area is effectively increased, the flow loss is reduced, and the heat exchange effect is finally improved.
Furthermore, the suction surface and the pressure surface are both provided with penetrating ribs along the main flow direction, the penetrating ribs sequentially penetrate through a plurality of obliquely arranged ribs on the suction surface or the pressure surface, and through the penetrating ribs, the heat exchange of the channel structure is more uniform, and the flow loss is effectively reduced.
Furthermore, a plurality of transverse micro-teeth are arranged on the inclined ribs and the transverse ribs along the main flow direction, and the longitudinal secondary flow of the cooling working medium is facilitated to be enhanced by the micro-teeth, so that the heat exchange capacity is improved.
Drawings
FIG. 1 is a schematic illustration of a channel configuration for internal cooling of a gas turbine vane in an embodiment of the present invention;
FIG. 2 is a top view of a channel structure for internal cooling of a gas turbine vane in an embodiment of the present invention;
FIG. 3 is a schematic diagram of an included angle between an inclined rib and a main flow direction according to the present invention;
FIG. 4 is a schematic illustration of a further channel configuration for internal cooling of a gas turbine vane in an embodiment of the present invention;
FIG. 5 is a temperature distribution cloud of a prior art channel structure for internal cooling of a gas turbine vane;
FIG. 6 is a temperature profile cloud for the channel structure for internal cooling of a gas turbine vane of the present invention.
Wherein: 1-a straight channel; 2-bending the channel; 3-obliquely arranging ribs; 4-transverse ribs; 5-through the ribs.
Detailed Description
In order to make the technical solutions of the present invention better understood, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that the terms "first," "second," and the like in the description and claims of the present invention and in the drawings described above are used for distinguishing between similar elements and not necessarily for describing a particular sequential or chronological order. It is to be understood that the data so used is interchangeable under appropriate circumstances such that the embodiments of the invention described herein are capable of operation in sequences other than those illustrated or described herein. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or elements is not necessarily limited to those steps or elements expressly listed, but may include other steps or elements not expressly listed or inherent to such process, method, article, or apparatus.
The invention is described in further detail below with reference to the accompanying drawings:
referring to fig. 1 to 3, in an embodiment of the present invention, a channel structure for cooling inside a stationary blade of a gas turbine is provided, which can effectively improve heat transfer in a low reynolds number region downstream of a fin, so that temperature distribution is more uniform, a local hot spot of the stationary blade of the gas turbine is eliminated, thermal stress is improved, stability of the stationary blade of the gas turbine is improved, and a service life of the stationary blade of the gas turbine is prolonged.
Specifically, the channel structure for cooling the inside of the static blade of the gas turbine comprises a plurality of straight channels 1 and a plurality of bent channels 2; a plurality of straight channels 1 are arranged in parallel, and two adjacent straight channels 1 are connected through a bent channel 2; defining the wall surface at the upper side of the main flow direction in the straight channel 1 as a suction surface; the wall surface positioned at the lower side of the main flow direction is a pressure surface; the suction surface and the pressure surface are both provided with a plurality of inclined ribs 3; a plurality of transverse ribs 4 are arranged on the wall surface connected with the downstream ends of the inclined ribs 3 in the straight channel 1, and the downstream ends of the inclined ribs 3 are the downstream ends of the inclined ribs 3 in the projection of the inclined ribs 3 in the main flow direction; one end of the transverse rib 4 is connected with the inclined rib 3 on the suction surface, and the other end is connected with the inclined rib 3 on the pressure surface. Wherein, the main flow direction is the direction from the inlet to the outlet of the whole channel structure, and the cooling working medium generally adopts compressed cold air.
In this embodiment, three straight channels 1 and two curved channels 2 are provided, and the three straight channels 1 are arranged in parallel and connected by the two curved channels 2 to form an s-shaped three-flow-path channel.
During specific work, after a cooling working medium enters the channel, boundary layer separation occurs at the top ends of the ribs of the inclined ribs 3, and the boundary layer separation occurs in the downstream intercostal area, so that longitudinal secondary flow is generated. The longitudinal secondary flow can develop along the direction of the fins and is converged with the main secondary flow at the downstream end of the fins, and because of the existence of the lateral wall surface transverse ribs 4, the boundary layer along the lateral wall surface can also generate separation and reattachment phenomena at the transverse ribs 4, so that the boundary layer of the lateral wall surface is continuously damaged and is difficult to develop and thicken, thereby strengthening the mixing of cooling working media in the channel, improving the temperature gradient of the cooling working media near the wall surface, further effectively improving the heat and mass transfer at the downstream of the fins, ensuring more uniform temperature distribution, realizing better cooling effect, eliminating local hot spots on the stationary blade of the gas turbine, improving the thermal stress, improving the stability of the stationary blade of the gas turbine and prolonging the service life of the stationary blade of the gas turbine.
Preferably, the angle α between the inclined rib 3 and the main flow direction is 30 ° to 150 °, and preferably, the angle α between the inclined rib 3 and the main flow direction is 60 °.
Preferably, the width-to-height ratio of the straight channel 1 is 0.25-2, and preferably, the width-to-height ratio of the straight channel 1 is 1.
Preferably, the height of the inclined ribs 3 and the transverse ribs 4 is 1-3 mm, and the height of the inclined ribs 3 and the transverse ribs 4 is 2 mm.
Preferably, a plurality of inclined ribs 3 on the suction surface and the pressure surface are uniformly distributed, the distance between every two adjacent inclined ribs 3 is 10-20 mm, and preferably, the distance between every two adjacent inclined ribs 3 is 16 mm.
Preferably, the suction surface and the pressure surface are curved surfaces, and are respectively consistent with the curves of the two outer wall surfaces of the gas turbine stationary blade. Through setting up suction surface and pressure surface into the curved surface unanimous with the bending of two outer wall surfaces of gas turbine quiet leaf, and then effectively promote heat transfer area and reduce the flow loss, finally improve the heat transfer effect.
Preferably, referring to fig. 4, the suction surface and the pressure surface are both provided with through ribs 5 along the main flow direction, and the through ribs 5 sequentially pass through a plurality of inclined ribs 3 on the suction surface or the pressure surface. Through setting up and running through rib 5, make channel structure's heat transfer more even, effectively reduce the flow loss.
Preferably, a plurality of transverse micro-teeth are arranged on the inclined ribs 3 and the transverse ribs 4 along the main flow direction. Wherein, the width of the microteeth is 0.87mm, the height is 0.5mm, and the space between the adjacent microteeth is 0.87 mm. The micro-teeth are arranged on the rib surfaces of the inclined ribs 3 and the transverse ribs 4, so that the longitudinal secondary flow of the cooling working medium is enhanced, and the heat exchange capacity is improved.
Referring to fig. 5 and 6, a simulation of the cooling effect of the prior channel structure for cooling the inside of a gas turbine stationary blade and the channel structure for cooling the inside of a gas turbine stationary blade according to the present invention is shown, in this embodiment, in the channel structure for cooling the inside of a gas turbine stationary blade according to the present invention, the inclined rib 3 forms an angle of 60 ° with the main flow direction, the width-to-height ratio of the straight channel 1 is 1, the height of the inclined rib 3 and the transverse rib 4 is 2mm, and the distance between adjacent inclined ribs 3 is 16 mm.
As can be seen from the figure, the lighter the color is, the higher the surface temperature is, the channel structure for cooling the inside of the static blade of the gas turbine in the prior art has obvious high-temperature areas at the positions of the three channels and the downstream of the fins, and the temperature is higher and is close to the temperature of a main flow working medium of the gas turbine. Such high temperature regions generate large thermal stress, which not only reduces various mechanical properties of the metal blades, but also easily causes blade burnout, thereby seriously affecting the economy and safety of the gas turbine operation. The channel structure for cooling the inside of the stationary blade of the gas turbine can greatly reduce the area of a high-temperature area at the downstream of the fins, and can reduce the area of the high-temperature area by more than 50% at the front part and the middle part of the three channels. It is worth noting that the boundary conditions and the working medium inlet parameters used by the two simulations are identical. The cooling effect of the existing channel structure for cooling the inside of the stationary blade of the gas turbine is limited, and the area of the high-temperature area is larger, so that the heat absorption capacity of a working medium is less than that of the channel structure for cooling the inside of the stationary blade of the gas turbine, and the temperature of the working medium at the outlet of the channel is lower than that of the channel structure, so that the area reduction effect of the high-temperature area is not obvious near the outlet of the channel.
Therefore, compared with the existing channel structure for cooling the inside of the stationary blade of the gas turbine, the channel structure for cooling the inside of the stationary blade of the gas turbine can obviously reduce the area of high-temperature hot spots, eliminate partial hot spots, reduce thermal stress, improve various mechanical properties of metal blades, increase the area of a low-temperature area, enable the temperature distribution of the wall surface to be more uniform and strengthen the heat exchange effect.
In an embodiment of the present invention, a gas turbine stationary blade is provided, in which the above-described channel structure for cooling the inside of the gas turbine stationary blade is provided. The invention relates to a gas turbine stator blade, which realizes the internal cooling of the gas turbine stator blade by digging the channel structure for the internal cooling of the gas turbine stator blade in the existing gas turbine stator blade.
The above-mentioned contents are only for illustrating the technical idea of the present invention, and the protection scope of the present invention is not limited thereby, and any modification made on the basis of the technical idea of the present invention falls within the protection scope of the claims of the present invention.
Claims (10)
1. A channel structure for cooling the inside of a stationary blade of a gas turbine is characterized by comprising a plurality of straight channels (1) and a plurality of curved channels (2);
a plurality of straight channels (1) are arranged in parallel, and two adjacent straight channels (1) are connected through a bent channel (2);
defining the wall surface at the upper side of the main flow direction in the straight channel (1) as a suction surface; the wall surface positioned at the lower side of the main flow direction is a pressure surface; the suction surface and the pressure surface are both provided with a plurality of inclined ribs (3);
a plurality of transverse ribs (4) are arranged on the wall surface connected with the downstream end of the inclined rib (3) in the straight channel (1), and the downstream end of the inclined rib (3) is the downstream end of the inclined rib (3) in the projection of the inclined rib (3) in the main flow direction; one end of the transverse rib (4) is connected with the inclined rib (3) on the suction surface, and the other end is connected with the inclined rib (3) on the pressure surface.
2. The channel structure for internal cooling of a gas turbine vane according to claim 1, characterized in that the angle of the slanted rib (3) to the main flow direction is 30 ° to 150 °.
3. The channel structure for internal cooling of a gas turbine vane according to claim 1, wherein the width to height ratio of the straight channel (1) is 0.25 to 2.
4. The channel structure for internal cooling of a gas turbine stationary blade according to claim 1, wherein the height of the diagonal ribs (3) and the transverse ribs (4) is 1 to 3 mm.
5. The channel structure for cooling inside a gas turbine vane according to claim 1, wherein the plurality of inclined ribs (3) on the suction surface and the pressure surface are uniformly distributed, and the distance between adjacent inclined ribs (3) is 10-20 mm.
6. The channel structure for internal cooling of a gas turbine vane as claimed in claim 1, wherein the suction surface and the pressure surface are curved surfaces respectively conforming to the curvature of both outer wall surfaces of the gas turbine vane.
7. The channel structure for the internal cooling of a gas turbine vane according to claim 1, characterized in that the suction side and the pressure side are each provided with a through-going rib (5) in the main flow direction, the through-going ribs (5) passing in turn through several slanted ribs (3) on the suction side or the pressure side.
8. The channel structure for the internal cooling of a gas turbine vane according to claim 1, characterized in that the inclined ribs (3) and the transverse ribs (4) are each provided with a number of transverse microteeth in the main flow direction.
9. The channel structure for internal cooling of a gas turbine vane according to claim 1, characterized in that the number of the straight channels (1) is 3.
10. A gas turbine vane characterized in that the passage structure for internal cooling of a gas turbine vane as claimed in any one of claims 1 to 9 is provided inside the gas turbine vane.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110528139.3A CN113266436B (en) | 2021-05-14 | 2021-05-14 | Channel structure for cooling inside of gas turbine stationary blade and gas turbine stationary blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110528139.3A CN113266436B (en) | 2021-05-14 | 2021-05-14 | Channel structure for cooling inside of gas turbine stationary blade and gas turbine stationary blade |
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| Publication Number | Publication Date |
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| CN113266436A true CN113266436A (en) | 2021-08-17 |
| CN113266436B CN113266436B (en) | 2022-10-25 |
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Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH1037704A (en) * | 1996-07-19 | 1998-02-10 | Mitsubishi Heavy Ind Ltd | Stator blade of gas turbine |
| US6257830B1 (en) * | 1997-06-06 | 2001-07-10 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade |
| US20020197161A1 (en) * | 2001-06-11 | 2002-12-26 | Norman Roeloffs | Gas turbine airfoill |
| US20130343872A1 (en) * | 2011-02-17 | 2013-12-26 | Rolls-Royce Plc | Cooled component for the turbine of a gas turbine engine |
| CN104791020A (en) * | 2015-04-23 | 2015-07-22 | 华能国际电力股份有限公司 | Gas turbine blade with longitudinal crossed rib cooling structure |
| CN110234840A (en) * | 2017-01-31 | 2019-09-13 | 西门子股份公司 | Turbine rotor blade or Turbomachinery for gas turbine |
| CN111271133A (en) * | 2020-03-09 | 2020-06-12 | 北京南方斯奈克玛涡轮技术有限公司 | Turbine guider blade with complex fin structure inner cooling channel |
-
2021
- 2021-05-14 CN CN202110528139.3A patent/CN113266436B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH1037704A (en) * | 1996-07-19 | 1998-02-10 | Mitsubishi Heavy Ind Ltd | Stator blade of gas turbine |
| WO1999036675A1 (en) * | 1996-07-19 | 1999-07-22 | Mitsubishi Heavy Industries, Ltd. | Stationary blade of gas turbine |
| US6257830B1 (en) * | 1997-06-06 | 2001-07-10 | Mitsubishi Heavy Industries, Ltd. | Gas turbine blade |
| US20020197161A1 (en) * | 2001-06-11 | 2002-12-26 | Norman Roeloffs | Gas turbine airfoill |
| US20130343872A1 (en) * | 2011-02-17 | 2013-12-26 | Rolls-Royce Plc | Cooled component for the turbine of a gas turbine engine |
| CN104791020A (en) * | 2015-04-23 | 2015-07-22 | 华能国际电力股份有限公司 | Gas turbine blade with longitudinal crossed rib cooling structure |
| CN110234840A (en) * | 2017-01-31 | 2019-09-13 | 西门子股份公司 | Turbine rotor blade or Turbomachinery for gas turbine |
| CN111271133A (en) * | 2020-03-09 | 2020-06-12 | 北京南方斯奈克玛涡轮技术有限公司 | Turbine guider blade with complex fin structure inner cooling channel |
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| CN113266436B (en) | 2022-10-25 |
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