CN110700894A - Turbine rotor blade of gas turbine and gas turbine adopting same - Google Patents
Turbine rotor blade of gas turbine and gas turbine adopting same Download PDFInfo
- Publication number
- CN110700894A CN110700894A CN201911074717.XA CN201911074717A CN110700894A CN 110700894 A CN110700894 A CN 110700894A CN 201911074717 A CN201911074717 A CN 201911074717A CN 110700894 A CN110700894 A CN 110700894A
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- Prior art keywords
- cooling
- blade
- turbine rotor
- edge
- rotor blade
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Classifications
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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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
- 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
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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
- 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
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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
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
A turbine rotor blade of a gas turbine comprises a blade profile, a blade root and a blade platform connecting the blade profile and the blade root; the outer surface of the blade profile of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively; the blade is internally provided with a cooling channel structure which comprises at least three cooling channels for cooling the tail edge of the blade. The invention divides the trailing edge into at least three areas along the blade height direction in the area close to the trailing edge of the blade, and the cooling air flow of the inlets of the cooling channels of the at least three cooling trailing edges can be independently adjusted, thereby easily achieving the expected cooling distribution effect, and particularly greatly reducing the cooling design difficulty of the areas for the blade root and the blade tip areas of the trailing edge.
Description
Technical Field
The invention relates to the technical field of gas turbine design, in particular to a turbine rotor blade of a gas turbine and the gas turbine adopting the turbine rotor blade.
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.
In the cooling design process of the turbine blade, due to the complex structure of the tail edge area of the blade, internal cooling gas and blade metal bear the action of centrifugal force in the radial direction, and the distribution of cold air quantity in the tail edge area is not uniform in the radial direction; in addition, the tail edge area is often positioned at the tail end of the cooling flow path, the temperature of cold air is higher, and the cooling effect is often poorer; the trailing edge region is also weaker in strength due to the thinner structure. Under the action of the factors, the tail edge area of the turbine rotor blade is easy to be subjected to high-temperature oxidation due to overhigh temperature or overlarge thermal stress caused by uneven distribution of cooling air, and the failure phenomena such as cracking, even ablation and the like occur.
Therefore, there is a need for more precise control of the flow of cooling air in the trailing edge region of the airfoil, with more reasonable and controllable radial distribution, without increasing the total cooling air volume, to reduce the temperature and thermal stress levels in the trailing edge region of the blade.
Disclosure of Invention
In view of the above, the main object of the present invention is to provide a turbine rotor blade of a gas turbine, which is intended to at least partially solve at least one of the above technical problems.
In order to achieve the above object, the present invention provides a turbine rotor blade of a gas turbine, comprising a blade profile, a blade root and a blade platform connecting the blade profile and the blade root; the outer surface of the blade profile of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively;
the blade is internally provided with a cooling channel structure which comprises at least three cooling channels for cooling the tail edge of the blade.
Among the cooling channels of the at least three cooling blade tail edges, at least one cooling channel is used for cooling the root part of the blade profile tail edge, at least one cooling channel is used for cooling the top part of the tail edge of the blade profile part, and at least one cooling channel is used for cooling the middle part of the blade profile tail edge.
Wherein the cooling air of the cooling passages of the at least three cooling blade trailing edges can be independently adjusted and supplied.
Wherein the cooling channels of the at least three cooling blade trailing edges are communicated internally, but each channel is also independently supplied with air.
At least one column rib structure is arranged inside the tail edge of the blade, at least two injection through holes are formed in the tail edge of the blade, and a partition plate of each injection through hole is arranged between every two injection through holes.
Wherein the blade further comprises an impingement cooling circuit for cooling the blade leading edge; at least one film cooling hole is arranged on the wall of the blade leading edge, and at least one impingement cooling hole is arranged in the impingement cooling loop; and cooling gas in the impingement cooling loop flows out through the film cooling hole after passing through the impingement cooling hole.
The cooling air flow turning part is provided with at least one flow deflector structure in the cooling channel of the tail edges of the at least three cooling blades, each flow deflector structure divides the turning area into two parts, and each flow deflector structure divides the cooling air entering the area where each flow deflector structure is located into two parts.
The inner wall surfaces of the cooling channels at the tail edges of the at least three cooling blades are provided with at least one fin structure or other turbulent flow structures, so that the heat exchange effect of the region is enhanced.
Wherein the cooling channels of the at least three cooling blade trailing edges are serpentine channels.
Based on the technical scheme, compared with the prior art, the turbine rotor blade disclosed by the invention has at least one of the following beneficial effects:
(1) in the area close to the trailing edge of the blade, the trailing edge is divided into at least three areas along the blade height direction.
(2) The cooling air flow rates at the cooling channel inlets of the at least three cooling tails are independently adjustable to more easily achieve the desired cooling distribution effect.
(3) Especially for the blade root and blade tip areas of the trailing edge, the cooling design difficulty of the areas can be greatly reduced.
Drawings
FIG. 1 is a three-dimensional view of a turbine rotor blade of a gas turbine engine of the present invention;
FIG. 2 is a schematic cross-sectional view of a turbine rotor blade internal cooling system of the gas turbine engine of the present invention;
FIG. 3 is a cross-sectional view A-A of FIG. 2;
FIG. 4 is a cross-sectional view B-B of FIG. 2;
FIG. 5 is a particular embodiment of a turbine rotor blade of a gas turbine engine of the present invention;
FIG. 6 is a cross-sectional view A-A of FIG. 5;
FIG. 7 is a cross-sectional view B-B of FIG. 5;
fig. 8 is a cross-sectional view C-C of fig. 5.
In the above drawings, the reference numerals have the following meanings:
1-root cooling channel inlet; 2-film cooling holes; 3-impingement cooling holes;
4-jet through holes; 5-top gas film hole;
11-blade root; 12-a blade platform; 13-blade profile;
15-a ribbed structure; 16-a column rib structure; 17-cooling channel partition plate;
18-a guide vane structure; 19-jet through hole partition plate;
20-profile camber line; 21-a second impingement cooling channel; 22-a first impingement cooling channel;
23-trailing edge first cooling channel; 24-trailing edge second lobe cooling channel; 25-trailing edge third lobe but channel;
26-a serpentine channel; 27-core support; 28-impingement cooling of the splitter plate;
31-suction surface; 32-pressure side; 33-blade leading edge; 34-the trailing edge of the blade;
101-first cooling channel inlet; 102-a second cooling channel inlet;
103-third cooling channel inlet; 104-a fourth cooling channel inlet;
a-high temperature fuel gas; b-cooling the air.
Detailed Description
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.
Specifically, the invention provides a turbine rotor blade of a gas turbine, which comprises a blade profile, a blade root and a blade platform for connecting the blade profile and the blade root; the outer surface of the blade profile of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively;
the blade is internally provided with a cooling channel structure which comprises at least three cooling channels for cooling the tail edge of the blade.
Among the cooling channels of the at least three cooling blade tail edges, at least one cooling channel is used for cooling the root part of the blade profile tail edge, at least one cooling channel is used for cooling the top part of the tail edge of the blade profile part, and at least one cooling channel is used for cooling the middle part of the blade profile tail edge.
Wherein the cooling air of the cooling passages of the at least three cooling blade trailing edges can be independently adjusted and supplied.
Wherein the cooling channels of the at least three cooling blade trailing edges are communicated internally, but each channel is also independently supplied with air.
At least one column rib structure is arranged inside the tail edge of the blade, at least two injection through holes are formed in the tail edge of the blade, and a partition plate of each injection through hole is arranged between every two injection through holes.
Wherein the blade further comprises an impingement cooling circuit for cooling the blade leading edge; at least one film cooling hole is arranged on the wall of the blade leading edge, and at least one impingement cooling hole is arranged in the impingement cooling loop; and cooling gas in the impingement cooling loop flows out through the film cooling hole after passing through the impingement cooling hole.
The cooling air flow turning part is provided with at least one flow deflector structure in the cooling channel of the tail edges of the at least three cooling blades, each flow deflector structure divides the turning area into two parts, and each flow deflector structure divides the cooling air entering the area where each flow deflector structure is located into two parts.
The inner wall surfaces of the cooling channels at the tail edges of the at least three cooling blades are provided with at least one fin structure or other turbulent flow structures, so that the heat exchange effect of the region is enhanced.
Wherein the cooling channels of the at least three cooling blade trailing edges are serpentine channels.
The accompanying drawings illustrate a specific embodiment of the present invention. Fig. 1 is a three-dimensional view of a high-temperature rotor blade of a gas turbine, comprising a blade profile 13, a blade root 11 and a blade platform 12 between the blade profile and the blade root. 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 interior has a plurality of cooling channels for the flow of cooling gas. The cooling gas is delivered to the root cooling channel inlets of the rotor blades, as shown in FIG. 3, the size and number of which are designed based on a combination of the desired cooling gas flow and the blade structural design.
At least three cooling flow paths inside the blade directly supply cooling air to the trailing edge region, as shown in fig. 2 and 4, and fig. 4 is a cross-sectional view B-B of fig. 2, which includes: the first cooling channel 23 at the tail edge, cooling air enters from the inlet 101 of the first cooling channel at the blade root, turns into axial flow after flowing along the radial direction, and is discharged from the tail edge after the tail edge of the cooling blade is close to the tail edge area of the blade tip; the tail edge second cooling channel 24 is used for cooling air entering from the blade root second cooling channel inlet 102, changing the cooling air into axial flow after flowing along the radial direction, and discharging the cooling air from the tail edge after the tail edge of the cooling blade is close to the middle area of the blade; and the third cooling channel 25 at the tail edge, cooling air enters from the third cooling channel inlet 103 at the blade root, changes into axial flow after flowing along the radial direction, and is discharged from the tail edge after cooling the tail edge of the blade close to the root area of the blade. The cooling air flow of the at least three channels for cooling the trailing edge may be independently adjusted according to the inlet geometry.
The at least three channels for cooling the trailing edge need to be internally communicated by the core support 27 if the internal air passages are not independent due to the process requirements of core support reinforcement, casting and the like, and the essence that each channel independently supplies air and performs zone adjustment on the trailing edge is not changed.
Fig. 5 shows another embodiment of the present invention. At least three cooling flow paths within the blade provide cooling air directly to the trailing edge region, FIG. 6 is a sectional view A-A of FIG. 5, FIG. 7 is a sectional view B-B of FIG. 5, and FIG. 8 is a sectional view C-C of FIG. 5, including: and a trailing edge first cooling channel 23, wherein channel cooling air enters from the root first cooling channel inlet 101, turns into axial flow after flowing along the radial direction, and is discharged from the trailing edge after the trailing edge of the cooling blade is close to the tail edge area of the blade tip. In the trailing edge first cooling channel 23, at least one fin structure 15 is arranged in the wall surface to enhance the convective cooling effect of the cooling channel wall surface. The geometry and arrangement of the fin structures is selected according to the requirements for heat transfer and pressure loss. The cooling gas should ensure sufficient pressure to be ejected from the at least one opening of the trailing edge first cooling channel 23 to the main flow of combustion gas and the at least one tip film hole 5 arranged at the tip of the blade.
The first cooling passage also includes an impingement cooling circuit. The impingement cooling circuit includes a first impingement cooling channel 22 and a second impingement cooling channel 21, the two impingement cooling channels separated by an impingement cooling partition plate 28; the impingement cooling partition plate has at least one impingement cooling hole 3 disposed therein. At least one fin structure 15 with an enhanced convection cooling effect is arranged on the wall surface of the first impingement cooling channel 22, and cooling gas is subjected to impingement cooling on the inner wall surface of a part of the blade leading edge in the second impingement cooling channel 21 through at least one impingement cooling hole 3 arranged on the impingement cooling partition plate 28 when the first impingement cooling channel 22 flows towards the blade top. At least one film cooling hole 2 is arranged on the wall of the blade leading edge 33, cooling gas passes through the impingement cooling hole 3 and then is sprayed out of the film cooling hole 2 to form a film to cover the blade suction surface 31 and the blade pressure surface 32, and therefore high-temperature combustion gas is separated. The geometrical dimensions such as the angle, the aperture and the like and the number of the film cooling holes 2 are determined according to the comprehensive influence effects of blade aerodynamics, heat transfer and the like.
The first cooling passage 22 also includes a baffle structure 18 at the intersection of the radial and axial cooling passages. When cooling air flows through the guide vane structure 18, the cooling air distribution of the two regions is automatically distributed according to the inlet areas of the upper and lower regions of the guide vane; the presence of the baffle structure 18 also reduces flow losses at the intersection of the radial and axial cooling channels, increasing the pressure margin in the aft region of the first cooling channel.
The first cooling channel is provided with at least one stud rib structure 16 in the region close to the trailing edge to enhance the convective cooling effect of the trailing edge of the blade. Meanwhile, the column rib structure 16 connects the suction surface 31 and the pressure surface 32 of the blade, and the effect of strengthening the structural strength of the blade is achieved. The geometrical dimensions of the column rib structure also need to be selected comprehensively according to the cooling effect and the strength requirement. At least two injection through holes 4 are arranged at the trailing edge of the blade, in which injection through holes the cooling gas is normally accelerated, thereby enhancing the convective cooling effect.
The second cooling channel 24 may be a serpentine channel structure, and after cooling air enters the serpentine channel from the second cooling channel 102 and the third cooling channel inlet 103, the cooling air cools the region of the trailing edge of the blade near the middle of the blade, and finally is discharged from the opening leading to the gas. In the trailing edge second cooling channel 24, at least one fin structure 15 is arranged in the wall surface to enhance the convective cooling effect of the cooling channel wall surface. The geometry and arrangement of the fin structures is selected according to the requirements for heat transfer and pressure loss. The cooling gas should ensure a sufficient pressure to be able to exit from the at least one opening of the trailing edge second cooling channel 24 to the main flow of gas.
The second cooling passage has at least one stud rib structure 16 disposed proximate the trailing edge region to enhance the convective cooling effect of the blade trailing edge. Meanwhile, the column rib structure is connected with the suction surface 31 and the pressure surface 32 of the blade, and the effect of strengthening the structural strength of the blade is achieved. The geometrical dimensions of the column rib structure also need to be selected comprehensively according to the cooling effect and the strength requirement. At least two injection through holes 4 are arranged at the trailing edge of the blade, in which injection through holes the cooling gas is normally accelerated, thereby enhancing the convective cooling effect.
And a third cooling channel 25 at the tail edge, wherein cooling air enters from a fourth cooling channel inlet 104 at the blade root, cools the area of the tail edge of the blade close to the root of the blade and finally is discharged from an opening leading to the fuel gas. The trailing edge tertiary cooling passage has at least one stud rib structure 16 disposed proximate the trailing edge region to enhance the convective cooling effect of the blade trailing edge. Meanwhile, the column rib structure is connected with the suction surface 31 and the pressure surface 32 of the blade, and the effect of strengthening the structural strength of the blade is achieved. The geometrical dimensions of the column rib structure also need to be selected comprehensively according to the cooling effect and the strength requirement. At least two injection through holes 4 are arranged at the trailing edge of the blade, in which injection through holes the cooling gas is normally accelerated, thereby enhancing the convective cooling effect.
The at least three channels for cooling the trailing edge need to be internally communicated by the core support 27 if the internal air passages are not independent due to the process requirements of core support reinforcement, casting and the like, and the essence that each channel independently supplies air and performs zone adjustment on the trailing edge is not changed.
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 of a gas turbine is characterized by comprising a blade profile, a blade root and a blade platform for connecting the blade profile and the blade root; the outer surface of the blade profile of the blade consists of a suction surface and a pressure surface, and the junction areas of the suction surface and the pressure surface are the front edge and the tail edge of the blade respectively;
the blade is internally provided with a cooling channel structure which comprises at least three cooling channels for cooling the tail edge of the blade.
2. The turbine rotor blade according to claim 1, wherein at least one of the at least three cooling passages cooling the trailing edge of the blade is configured to cool a root portion of the airfoil, at least one of the cooling passages is configured to cool a tip portion of the trailing edge of the airfoil, and at least one of the cooling passages is configured to cool a mid portion of the trailing edge of the airfoil.
3. The turbine rotor blade according to claim 1, wherein the cooling air of the cooling passages of the at least three cooling blade trailing edges is independently adjustable and feedable.
4. The turbine rotor blade according to claim 1 wherein the cooling passages of the at least three cooling blade trailing edges are in communication but each passage is independently fed.
5. The turbine rotor blade according to claim 1, wherein at least one post-rib structure is disposed inside the trailing edge of the blade and has at least two injection through-holes, and a partition plate for the injection through-holes is disposed between each two injection through-holes.
6. The turbine rotor blade according to claim 1, wherein the blade further comprises an impingement cooling circuit for cooling the leading edge of the blade; at least one film cooling hole is arranged on the wall of the blade leading edge, and at least one impingement cooling hole is arranged in the impingement cooling loop; and cooling gas in the impingement cooling loop flows out through the film cooling hole after passing through the impingement cooling hole.
7. The turbine rotor blade according to claim 1 wherein the cooling passages in the at least three cooling blade trailing edges have at least one baffle structure at the cooling gas flow turn, each baffle structure dividing the turn region into two and each baffle structure dividing the cooling gas entering the region of each baffle structure into two.
8. The turbine rotor blade according to claim 1, wherein the inner wall surfaces of the cooling passages of the at least three cooling blade trailing edges have at least one fin structure or other flow perturbation structure thereon to enhance heat exchange in that region.
9. The turbine rotor blade according to claim 1, wherein the cooling channels of the at least three cooling blade trailing edges are serpentine channels.
10. A gas turbine, characterized in that a turbine rotor blade according to any of claims 1-9 is used.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911074717.XA CN110700894A (en) | 2019-11-05 | 2019-11-05 | Turbine rotor blade of gas turbine and gas turbine adopting same |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911074717.XA CN110700894A (en) | 2019-11-05 | 2019-11-05 | Turbine rotor blade of gas turbine and gas turbine adopting same |
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| Publication Number | Publication Date |
|---|---|
| CN110700894A true CN110700894A (en) | 2020-01-17 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201911074717.XA Pending CN110700894A (en) | 2019-11-05 | 2019-11-05 | Turbine rotor blade of gas turbine and gas turbine adopting same |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111677557A (en) * | 2020-06-08 | 2020-09-18 | 清华大学 | Turbine guide blade and turbo machine with same |
-
2019
- 2019-11-05 CN CN201911074717.XA patent/CN110700894A/en active Pending
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN111677557A (en) * | 2020-06-08 | 2020-09-18 | 清华大学 | Turbine guide blade and turbo machine with same |
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