CN111852575A - Turbine rotor blade and gas turbine comprising same - Google Patents
Turbine rotor blade and gas turbine comprising same Download PDFInfo
- Publication number
- CN111852575A CN111852575A CN202010734235.9A CN202010734235A CN111852575A CN 111852575 A CN111852575 A CN 111852575A CN 202010734235 A CN202010734235 A CN 202010734235A CN 111852575 A CN111852575 A CN 111852575A
- Authority
- CN
- China
- Prior art keywords
- cold air
- impact
- air channel
- blade
- turbine rotor
- 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.)
- Pending
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
Abstract
A turbine rotor blade and a gas turbine comprising the same, wherein the turbine rotor blade comprises a blade body, the blade body comprises a cold air channel and an impact plate; the cold air channel is arranged inside the blade body; the impact plate is arranged inside the blade body; the impact plate is provided with impact holes, and the impact holes are communicated with the cold air channel and used for performing impact strengthening cooling on the top area of the blade body of the blade by cooling air in the cold air channel through the impact holes. The turbine rotor blade adopts a complex cold air channel structure inside the turbine rotor blade, and adopts the shock-enhanced cooling in the area of the cold air channel inside the blade body close to the blade tip, so that the metal temperature of the blade tip area of the rotor blade is reduced and the service life of the turbine rotor blade is prolonged by utilizing the characteristics of high air pressure margin and high shock cooling heat exchange coefficient of the rotor blade.
Description
Technical Field
The invention relates to the technical field of gas turbines, in particular to a turbine rotor blade and a gas turbine comprising the same.
Background
With the increasing level of gas turbine design technology, the gas turbine inlet gas temperature is increasing continuously, and the thermal load of turbine parts is extremely high, and the limit that high-temperature materials can bear is already exceeded. In order to ensure safe and reliable operation of the turbine blade, it is necessary to design the turbine blade with a complex cooling system to maintain the temperature and stress distribution of the blade body at a reasonable level.
Among various internal cooling heat exchange technologies, the impingement cooling heat exchange coefficient is the highest, the heat exchange effect is the best, but the pressure loss is also the largest, so the required cooling air pressure is higher.
In the design process of cooling the turbine blade, a three-dimensional flow structure such as leakage flow exists in a blade tip area, air used for cooling is heated and heated due to the fact that the blade tip area is far away from a cold air inlet, and in addition, the heat exchange coefficient of an internal convection cooling structure is relatively low, the cooling effect is often poor. Under the action of the factors, the blade tip area of the turbine rotor blade is easy to fail due to high temperature or high-temperature oxidation caused by high thermal stress due to poor cooling effect, and cracks, even ablation and the like occur.
Disclosure of Invention
In view of the above, the main object of the present invention is to provide a turbine rotor blade and a gas turbine including the same, which are intended to at least partially solve at least one of the above mentioned technical problems.
In order to achieve the purpose, the technical scheme of the invention is as follows:
as one aspect of the present invention, a turbine rotor blade is provided, comprising a blade body including a cold air channel and an impingement plate;
the cold air channel is arranged inside the blade body;
the impact plate is arranged inside the blade body;
the impact plate is provided with impact holes, and the impact holes are communicated with the cold air channel and used for performing impact strengthening cooling on the tip area of the blade body of the blade by cooling air in the cold air channel through the impact holes.
As another aspect of the present invention, there is also provided a gas turbine including the turbine rotor blade as described above.
Based on the technical scheme, compared with the prior art, the invention has at least one or one part of the following beneficial effects:
under the condition of not increasing the total cooling air quantity, the invention utilizes the characteristics of high cold air pressure and high heat load of the blade tip area and adopts an impact cooling structure to carry out enhanced heat exchange on the blade tip and the wall surface of the cavity after impact, thereby effectively reducing the metal temperature of the blade tip area.
Drawings
FIG. 1 is a perspective view of a turbine rotor blade according to embodiments 1-3 of the present invention;
FIG. 2 is a schematic cross-sectional view of the internal structure of a turbine rotor blade according to embodiment 1 of the present invention;
FIG. 3 is a sectional view taken along line A-A of FIG. 2;
FIG. 4 is a sectional view taken along line B-B of FIG. 2;
FIG. 5 is a cross-sectional view C-C of FIG. 2;
FIG. 6 is a schematic sectional view of the internal structure of a turbine rotor blade according to embodiment 2 of the present invention;
FIG. 7 is a cross-sectional view taken along line D-D of FIG. 6;
FIG. 8 is a cross-sectional view taken along line E-E of FIG. 6;
fig. 9 is a sectional view of fig. 6 taken along the line F-F.
FIG. 10 is a schematic sectional view of the inner structure of a turbine rotor blade according to embodiment 3 of the present invention;
FIG. 11 is a sectional view taken along line G-G of FIG. 10;
FIG. 12 is a sectional view taken along line H-H of FIG. 10;
fig. 13 is a sectional view taken along line J-J of fig. 10.
In the above figures, the reference numerals have the following meanings:
1-root cold air channel inlet; 2-film cooling holes; 3. 17-impingement holes; 4-trailing edge jet orifice; 5-air film hole; 8-top cover plate; 9-an impact plate; 11-blade root; 12-a blade platform; 13-blade body; 14-a cold air channel; 15-trailing edge cold air channel; 16-impingement cooling air channel; 18-a leading edge channel; 20-profile camber line; 21-a gas collection cavity; 22-post-impact chamber; 23-tip flute; 24-a root cavity; 31-suction surface; 32-pressure side; 33-blade leading edge; 34-the trailing edge of the blade; 41-fin structure; 42-column rib structure; 101-a first cold air channel inlet; 102-a second cold air channel inlet; 103-a third cold air channel inlet; a-high temperature fuel gas; b-cooling the air.
Detailed Description
Aiming at the characteristics that the top area of the blade has higher pressure margin and higher thermal load due to the action of centrifugal force, the invention adopts impingement cooling to strengthen heat exchange under the condition of not increasing the total cooling air quantity, strengthens the heat exchange of the tip area of the turbine blade and effectively reduces the temperature and the thermal stress level of the tip area of the blade.
In order that the objects, technical solutions and advantages of the present invention will become more apparent, the present invention will be further described in detail with reference to the accompanying drawings in conjunction with the following specific embodiments.
As one aspect of the present invention, a turbine rotor blade is provided, comprising a blade body including a cold air channel and an impingement plate;
the cold air channel is arranged inside the blade body;
the impact plate is arranged inside the blade body of the blade;
the impact plate is provided with impact holes, and the impact holes are communicated with the cold air channel and used for performing impact strengthening cooling on the blade tip area of the blade body of the blade by cooling air in the cold air channel through the impact holes.
In the embodiment of the invention, a plurality of impact holes are arranged on the impact plate, and the diameter of each impact hole is smaller than the inner diameter of the cold air channel.
In the embodiment of the invention, the blade body further comprises a top cover plate, and the top cover plate forms a top space communicated with the cold air channel on the top in the blade body; the impact plate is arranged in the top space, and the top space is divided into a gas collection cavity and an impact rear cavity;
wherein, the gas collection cavity is communicated with the cold air channel; the chamber after impact is communicated with the gas collection cavity through the impact hole;
and the wall surface of the impact rear cavity is provided with a cold air hole for cooling the tip area of the blade body of the blade by cooling air in the impact rear cavity through the cold air hole. Wherein the flow and pressure of the cooling air in the post-impingement cavity of the turbine rotor blade tip region is adjustable.
In the embodiment of the invention, the top cover plate is fixed with the blade body of the blade through a brazing process and blocks the cold air channel of the blade to form an impacted chamber; the impingement plate structure is secured to the blade body by a brazing process.
In the embodiment of the invention, the impact hole machining mode comprises casting, electric spark, laser or mechanical machining;
the hole pattern of the impact hole comprises one or more of a circle, an ellipse, a square and a rhombus;
the arrangement mode of the plurality of the impact holes on the impact plate is in a sequential or staggered arrangement.
The impact holes can be machined before brazing, and can also be machined after the impact plate and the blade body are brazed and fixed.
In the embodiment of the invention, the cold air hole is arranged on the top cover plate;
the cold air holes are arranged on the pressure surface and/or the suction surface of the blade body corresponding to the impacted chamber; and/or
The cold air hole is arranged on the tail edge of the blade body corresponding to the impacted chamber.
More specifically, in embodiments of the present invention, the cooling apertures include film holes disposed on the tip cover plate, trailing edge spray holes disposed on the trailing edge of the blade body, and film cooling holes disposed on the pressure side, the leading edge of the blade, and the suction side.
In an embodiment of the invention, the root of the turbine rotor blade may be a cavity structure, the inner part of the turbine rotor blade is connected with a plurality of cooling air channels along the radial direction, and cooling air can firstly enter the cavity and then enter the plurality of cooling air channels along the radial direction.
At least one complex cooling air channel is arranged inside the turbine rotor blade, and the complex cooling air channel is integrally formed in the precision casting process of the turbine rotor blade.
In the embodiment of the invention, the tail ends of a plurality of cold air channels are all communicated to a gas collecting cavity, and the gas collecting cavity is correspondingly provided with an impact plate and an impact rear cavity.
More specifically, a plurality of complex cold air channels in the turbine rotor blade are collected in one air collecting cavity, and cooling air cools the top cover plate and the wall surface of the cavity after impact through impact holes in the impact plate.
In other embodiments of the present invention, the ends of the plurality of cold air channels are respectively connected to the at least two air collecting chambers; each gas collecting cavity is independently provided with a corresponding impact plate and an impact rear cavity.
More specifically, the plurality of complex cold air channels inside the turbine rotor blade do not all correspond to one air collecting cavity; but can design for every air conditioning passageway corresponds solitary gas collection chamber respectively, perhaps a plurality of air conditioning passageways correspond two at least gas collection chambeies, and every gas collection chamber sets up the impingement plate alone and strikes back cavity, and cooling air cools off top apron and impact back cavity wall through the impact hole on the impingement plate.
In the embodiment of the invention, the wall surface corresponding to the cold air channel is provided with the cold air channel film cooling hole for discharging the cooling air in the cold air channel through the cold air channel film cooling hole to cool the blade tip area of the blade body.
In conclusion, after the turbine rotor blade impacts, cooling air in the cavity enters the blade tip groove through the air film hole and the dust removal hole in the top cover plate to further cool the blade tip area; or the main flow fuel gas enters through a tail edge jet hole; or through film cooling holes on the pressure side or suction side. In addition, the cooling air in at least one complex cooling air channel inside the turbine rotor blade can be directly discharged to the outside of the turbine rotor blade through the cooling air channel film cooling hole without entering the air collecting cavity.
In an embodiment of the invention, the cold air channel is arranged radially inside the blade body;
the processing mode of the cold air channel comprises casting, electrochemistry or electric spark;
the cold air channel comprises a section of through hole in the extending direction, but the cold air channel is not limited to the section of through hole, and a plurality of sections of through holes which are communicated with each other can be also used; wherein, the inner diameters of the multiple sections of through holes are different;
more specifically, when the turbine rotor blade is twisted more along the blade height direction, two sections of butt holes can be used for forming.
A fin structure and/or a column rib structure for strengthening heat exchange are/is arranged in the cold air channel;
the turbine rotor blade further includes a turbulence generating device disposed within the cold air channel. Used for strengthening convection cooling and improving the heat exchange capability of the area.
As another aspect of the present invention, there is also provided a gas turbine including the turbine rotor blade as described above.
The technical solution of the present invention is further described below with reference to specific examples, but it should be noted that the following examples are only for illustrating the technical solution of the present invention, but the present invention is not limited thereto.
Example 1
As shown in FIG. 1, a turbine rotor blade includes a blade airfoil 13, a blade root 11, and a blade platform 12 between the blade airfoil 13 and the blade root 11. The internal configuration of the blade is obtained by cutting the blade along the profile camber line 20, as shown in figure 2. The blade has a plurality of cold air passages 14 inside for the flow of cooling air. The cooling gas is delivered to the root cooling gas channel inlet 1 of the turbine rotor blade, the size and number of the root cooling gas channel inlets 1 are designed according to the required cooling gas flow and the blade structure, and as shown in fig. 3, the root cooling gas channel inlet 1 comprises a first cooling gas channel inlet 101, a second cooling gas channel inlet 102 and a third cooling gas channel inlet 103.
At least one cooling flow path within the blade interior provides cooling air to the tip region, as shown in fig. 2 and 4, and fig. 4 is a cross-sectional view B-B of fig. 2, including: three radial cold air channels 14 and a first cold air channel inlet 101, a second cold air channel inlet 102, a third cold air channel inlet 103 providing cooling air for the three cold air channels 14; after entering the cold air channel 14 from the first cold air channel inlet 101, the second cold air channel inlet 102, and the third cold air channel inlet 103, the cooling air enters the air collecting chamber 21 after having heat convection with the blades on the inner surface of the cold air channel 14.
As shown in fig. 2 and 5, the high-pressure cooling air in the air collecting cavity 21 performs impingement cooling on the top cover plate 8 through the impingement holes 3 on the impingement plate 9 and enters the impingement rear chamber 22, enters the blade top groove 23 from the air film holes 5 on the top cover plate 8, further cools the blade tip region and is mixed with the blade tip leakage flow; the cooling air entering the post impingement plenum 22 may be discharged directly into the mainstream combustion gases through the trailing edge injection holes 4 on the trailing edge 34 of the blade; the cooling air entering the post-impingement plenum 22 may enter the main flow of combustion gases through film cooling holes 2 disposed on the pressure side 32, suction side 31, and blade leading edge 33.
Example 2
As shown in fig. 6-9, a turbine rotor blade includes: three radial cooling air channels 14 and a root cavity 24 for supplying cooling air to the three cooling air channels 14; the cooling air enters the blade root cavity 24 from the first cooling air channel inlet 101, then enters the cooling air channel 14, and enters the gas collecting cavity 21 after the heat convection between the inner surface of the cooling air channel 14 and the blades occurs.
High-pressure cooling air in the gas collecting cavity 21 performs impact cooling on the top cover plate 8 through the impact holes 3 on the impact plate 9, enters the impact rear cavity 22, enters the blade top groove 23 from the air film holes 5 on the top cover plate 8, further cools the blade tip area and is mixed with blade tip leakage flow; the cooling air entering the post impingement plenum 22 may be discharged directly into the mainstream combustion gases through the trailing edge injection holes 4 on the trailing edge 34 of the blade; the cooling air entering the post-impingement plenum 22 may enter the main flow of combustion gases through film cooling holes 2 disposed on the pressure side 32, suction side 31, and blade leading edge 33.
Example 3
As shown in fig. 10-13, a turbine rotor blade includes: three complex cold air channels, a leading cold air channel 14, a trailing cold air channel 15, an impingement cold air channel 16, and a first cold air channel inlet 101, a second cold air channel inlet 102, a third cold air channel inlet 103 providing cooling air for the three cold air channels; after cooling air enters the cold air channel 14, the trailing edge cold air channel 15 and the impingement cold air channel 16 from the first cold air channel inlet 101, the second cold air channel inlet 102 and the third cold air channel inlet 103, heat convection occurs between the inner surfaces of the cold air channel 14, the trailing edge cold air channel 15 and the impingement cold air channel 16 and the blades.
After entering the impingement cold air channel 16, the cooling air in the first cold air channel inlet 101 enters the leading edge channel 18 along the impingement holes 17 and performs impingement cooling on the leading edge 33 of the blade, and the cooling air enters the main flow of combustion gas from the film cooling holes 2 of the leading edge (i.e., the cold air channel film cooling holes); after entering the trailing edge cold air channel 15, the cooling air in the third cold air channel inlet 103 performs enhanced heat exchange with the rib structure 41, enters the column rib structure 42 in the area of the trailing edge 34 of the blade to continue enhanced heat exchange, and then is ejected from the trailing edge ejection hole 4 and enters the main flow fuel gas.
After entering the cold air channel 14, the cooling air in the second cold air channel inlet 102 performs heat exchange with the fin structure 41, and then enters the air collecting chamber 21.
High-pressure cooling air in the gas collecting cavity 21 performs impact cooling on the top cover plate 8 through the impact holes 3 on the impact plate 9, enters the impact rear cavity 22, enters the blade top groove 23 from the air film holes 5 on the top cover plate 8, further cools the blade tip area and is mixed with blade tip leakage flow; the cooling air entering the post impingement plenum 22 may be discharged directly into the mainstream combustion gases through the trailing edge injection holes 4 on the trailing edge 34 of the blade; the cooling air entering the impingement rear chamber 22 may enter the main flow of combustion gases through film cooling holes 2 arranged on the pressure side 32, suction side 31 and leading edge 33.
The above-mentioned embodiments are intended to illustrate the objects, technical solutions and advantages of the present invention in further detail, and it should be understood that the above-mentioned embodiments are only exemplary embodiments of the present invention and are not intended to limit the present invention, and any modifications, equivalents, improvements and the like made within the spirit and principle of the present invention should be included in the protection scope of the present invention.
Claims (10)
1. A turbine rotor blade comprising a blade body, wherein the blade body comprises a cold air channel and an impingement plate;
the cold air channel is arranged inside the blade body;
the impact plate is arranged inside the blade body;
the impact plate is provided with impact holes, and the impact holes are communicated with the cold air channel and used for performing impact strengthening cooling on the tip area of the blade body of the blade by cooling air in the cold air channel through the impact holes.
2. The turbine rotor blade according to claim 1, wherein the impingement plate is provided with a plurality of impingement holes having a hole diameter smaller than an inner diameter of the cold air passage.
3. The turbine rotor blade according to claim 1 wherein said blade body further comprises a tip cover plate, said tip cover plate forming a tip space in communication with said cooling air passage at an inner tip of said blade body; the impact plate is arranged in the top space and divides the top space into a gas collection cavity and an impacted cavity;
the gas collection cavity is communicated with the cold air channel; the post-impact chamber is communicated with the gas collecting cavity through an impact hole;
and the wall surface of the impact rear cavity is provided with a cold air hole for cooling the blade tip area of the blade body by cooling air in the impact rear cavity through the cold air hole.
4. The turbine rotor blade according to claim 3,
the cold air hole is arranged on the top cover plate;
the cold air hole is arranged on the pressure surface and/or the suction surface of the blade body corresponding to the chamber after impact; and/or
The cold air hole is arranged on the tail edge of the blade body corresponding to the impact rear chamber.
5. The turbine rotor blade according to claim 4, wherein a plurality of said cold gas passages each terminate in a plenum, said plenums being provided with an impingement plate and an impingement post-chamber, respectively.
6. The turbine rotor blade according to claim 4, wherein a plurality of said cold gas channel ends are respectively connected to at least two gas collecting chambers; each gas collecting cavity is independently provided with a corresponding impact plate and an impact rear cavity.
7. The turbine rotor blade according to claim 1, wherein a cold air channel film cooling hole is provided on a corresponding wall surface of the cold air channel for discharging cooling air in the cold air channel through the cold air channel film cooling hole to cool a tip region of the blade body.
8. The turbine rotor blade according to claim 1,
the cold air channel is arranged in the blade body along the radial direction;
the processing mode of the cold air channel comprises casting, electrochemistry or electric spark;
the cold air channel comprises a section of through hole or a plurality of sections of through holes which are communicated with each other in the extending direction; wherein the inner diameters of the multiple sections of through holes are different;
a rib structure and/or a column rib structure for strengthening heat exchange are/is arranged in the cold air channel;
the turbine rotor blade further includes a turbulence generating device disposed within the cold air channel.
9. The turbine rotor blade according to claim 1,
the impact hole machining mode comprises casting, electric spark, laser or mechanical machining;
the hole pattern of the impact hole comprises one or more of a circle, an ellipse, a square and a rhombus;
the arrangement mode of the plurality of the impact holes on the impact plate is in a sequential or staggered arrangement.
10. A gas turbine comprising a turbine rotor blade according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010734235.9A CN111852575A (en) | 2020-07-27 | 2020-07-27 | Turbine rotor blade and gas turbine comprising same |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010734235.9A CN111852575A (en) | 2020-07-27 | 2020-07-27 | Turbine rotor blade and gas turbine comprising same |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN111852575A true CN111852575A (en) | 2020-10-30 |
Family
ID=72947415
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202010734235.9A Pending CN111852575A (en) | 2020-07-27 | 2020-07-27 | Turbine rotor blade and gas turbine comprising same |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN111852575A (en) |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112963862A (en) * | 2021-04-07 | 2021-06-15 | 西北工业大学 | Double-layer rhombic cross cooling structure |
| CN114278387A (en) * | 2021-12-22 | 2022-04-05 | 西安交通大学 | Blade top cooling structure and gas turbine movable blade adopting same |
| CN114412580A (en) * | 2022-02-09 | 2022-04-29 | 北京全四维动力科技有限公司 | Turbine blade air film cooling structure and gas turbine adopting same |
| CN114810216A (en) * | 2021-01-27 | 2022-07-29 | 中国航发商用航空发动机有限责任公司 | Aeroengine blade and aeroengine |
-
2020
- 2020-07-27 CN CN202010734235.9A patent/CN111852575A/en active Pending
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114810216A (en) * | 2021-01-27 | 2022-07-29 | 中国航发商用航空发动机有限责任公司 | Aeroengine blade and aeroengine |
| CN112963862A (en) * | 2021-04-07 | 2021-06-15 | 西北工业大学 | Double-layer rhombic cross cooling structure |
| CN114278387A (en) * | 2021-12-22 | 2022-04-05 | 西安交通大学 | Blade top cooling structure and gas turbine movable blade adopting same |
| CN114412580A (en) * | 2022-02-09 | 2022-04-29 | 北京全四维动力科技有限公司 | Turbine blade air film cooling structure and gas turbine adopting same |
| CN114412580B (en) * | 2022-02-09 | 2024-02-09 | 北京全四维动力科技有限公司 | Turbine blade air film cooling structure and gas turbine adopting same |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN111852575A (en) | Turbine rotor blade and gas turbine comprising same | |
| JP4509263B2 (en) | Backflow serpentine airfoil cooling circuit with sidewall impingement cooling chamber | |
| US6234754B1 (en) | Coolable airfoil structure | |
| US8753083B2 (en) | Curved cooling passages for a turbine component | |
| EP0416542B1 (en) | Turbine blade | |
| JP5599624B2 (en) | Turbine blade cooling | |
| US20070122282A1 (en) | Central cooling circuit for a moving blade of a turbomachine | |
| CN111927562A (en) | Turbine rotor blade and aircraft engine | |
| US9228440B2 (en) | Turbine blade airfoils including showerhead film cooling systems, and methods for forming an improved showerhead film cooled airfoil of a turbine blade | |
| US7001141B2 (en) | Cooled nozzled guide vane or turbine rotor blade platform | |
| CN113236370B (en) | Cooling structure of high-pressure moving blade of gas turbine | |
| CN103133040A (en) | Cooled turbine blade and corresponding method for cooling a turbine blade | |
| CN103089326A (en) | Method and apparatus for cooling gas turbine rotor blades | |
| CN210118169U (en) | Low-flow low-aspect-ratio high-pressure turbine cooling guide vane | |
| CN110700896B (en) | Gas turbine rotor blade with swirl impingement cooling structure | |
| CN113090335A (en) | Impact air-entraining film double-wall cooling structure for turbine rotor blade | |
| JP2015516539A (en) | Turbine airfoil trailing edge cooling slot | |
| EP3436669A1 (en) | Turbine airfoil with internal cooling channels having flow splitter feature | |
| CN210289846U (en) | Shrouded cooling turbine rotor blade leading edge cooling structure | |
| CN212535771U (en) | Turbine rotor blade and gas turbine comprising same | |
| EP1213442A1 (en) | Coolable airfoil structure | |
| CN112282860A (en) | Turbine rotor blade platform cooling structure | |
| CN112922676B (en) | Internal back basin rotary cooling channel of turbine blade | |
| CN113266429B (en) | Turbine guide vane end wall composite cooling structure | |
| KR20190082003A (en) | Gas turbine blade |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination |