CN114109514B - Turbine blade pressure surface cooling structure - Google Patents
Turbine blade pressure surface cooling structure Download PDFInfo
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- CN114109514B CN114109514B CN202111342660.4A CN202111342660A CN114109514B CN 114109514 B CN114109514 B CN 114109514B CN 202111342660 A CN202111342660 A CN 202111342660A CN 114109514 B CN114109514 B CN 114109514B
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- throttle
- impact
- throttling
- section
- blade pressure
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- 238000001816 cooling Methods 0.000 title claims abstract description 72
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 20
- 230000004888 barrier function Effects 0.000 claims abstract description 4
- 238000005192 partition Methods 0.000 claims description 2
- 239000007789 gas Substances 0.000 description 8
- 239000002131 composite material Substances 0.000 description 4
- 239000011159 matrix material Substances 0.000 description 4
- 230000000694 effects Effects 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000005266 casting Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 239000002356 single layer Substances 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the field of aero-engine blades, and particularly relates to a turbine blade pressure surface cooling structure. Comprising the following steps: the cooling unit is arranged a plurality of at blade pressure face, and the cooling unit separates into multistage through the barrier rib, and the cooling unit is provided with: the impact area is provided with an impact hole and an impact cavity communicated with the impact hole; the cross flow area is provided with turbulent flow columns, turbulent flow channels are arranged between the turbulent flow columns, and the turbulent flow channels are communicated with the impact cavity; the outlet section is provided with the throttle post, the throttle post includes first throttle post and second throttle post, the upstream section of two adjacent wall surfaces of first throttle post and second throttle post is the concentric circle form, the downstream section is the expansion form, form annular throttle passageway upstream section between two wall surface upstream sections of concentric circle form, form the throttle passageway downstream section of expansion form between two wall surface downstream sections of expansion form, throttle passageway upstream section constitutes the throttle passageway together with throttle passageway downstream section, throttle passageway and vortex passageway intercommunication, the air film hole has still been seted up to the outlet section.
Description
Technical Field
The application belongs to the field of aero-engine blades, and particularly relates to a turbine blade pressure surface cooling structure.
Background
With the development of aviation technology, the performance of an aeroengine is continuously improved, the temperature before a turbine is continuously increased, the temperature before the turbine is improved from 1700K level of a third-generation machine to 2000K level, and the heat load of turbine blades is greatly increased; meanwhile, the amount of cold air used for turbine blades is continuously reduced in pursuit of improvement of engine efficiency. The composite cooling structure commonly adopted in the three-fourth generation machine cannot meet the cooling requirement of the turbine blade at the inlet temperature of above 2000K.
The current most composite cooling structure adopts a cooling mode of impact convection and air film, and cold air forms impact cooling to the inner wall surface of the matrix through impact holes on the guide pipe, and then flows out through air film holes on the blades to form air film cooling to the outer wall surface of the matrix. The composite cooling structure has the following disadvantages:
a) Limited by the amount of cold gas and the processing capacity, the comprehensive cooling capacity basically reaches the limit;
b) The utilization rate of the cold air is low, the consumption of the cold air is high, the cold air is directly discharged from the air film holes after impacting the wall surface, the residence time of the cold air in the blades is short, and the temperature increase is limited;
c) The single-layer wall structure has thicker wall thickness and large heat conduction resistance, and is not beneficial to cooling the blade;
d) Circular air film holes are commonly adopted, and the air film cooling efficiency is low.
It is therefore desirable to have a solution that overcomes or at least alleviates at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
It is an object of the present application to provide a turbine blade pressure surface cooling structure that solves at least one of the problems of the prior art.
The technical scheme of the application is as follows:
a turbine blade pressure side cooling structure comprising:
a plurality of cooling units arranged on the blade pressure surface, the cooling units being divided into a plurality of sections by barrier ribs in the height direction, the cooling units being provided with an impingement zone, a cross flow zone and an outlet section, wherein,
the impact area is provided with an impact hole and an impact cavity communicated with the impact hole;
the cross flow area is provided with turbulent flow columns, turbulent flow channels are arranged between the turbulent flow columns, and the turbulent flow channels are communicated with the impact cavity;
the outlet section is provided with a throttling column, the throttling column comprises a first throttling column and a second throttling column, the upstream sections of two adjacent wall surfaces of the first throttling column and the second throttling column are in concentric circles, the downstream sections are in expanded forms, an annular throttling channel upstream section is formed between the two concentric wall surface upstream sections, an expanded throttling channel downstream section is formed between the two wall surface downstream sections in expanded forms, the throttling channel upstream section and the throttling channel downstream section form a throttling channel together, the throttling channel is communicated with the turbulence channel, and the outlet section is also provided with a gas film hole communicated with the throttling channel;
the cooling air flow inside the blade flows to the outside of the blade through the air film hole after sequentially passing through the impact hole, the impact cavity, the turbulent flow channel and the throttling channel.
In at least one embodiment of the application, the cooling units are arranged in 4 rows in sequence along the leading edge to the trailing edge of the blade.
In at least one embodiment of the present application, the cooling unit is divided into three sections in the height direction by two of the barrier ribs.
In at least one embodiment of the present application, the spoiler columns are diamond-shaped, and a plurality of spoiler columns are staggered.
In at least one embodiment of the present application, the gas film holes are in an expanded form.
In at least one embodiment of the present application, the gas film holes are inclined holes.
In at least one embodiment of the present application, the inclination angle of the air film hole is 45 degrees.
The application has at least the following beneficial technical effects:
the turbine blade pressure surface cooling structure can greatly improve the comprehensive cooling capacity of the blade; the utilization rate of the cold air is improved, and the consumption of the cold air is reduced; the heat conduction resistance of the matrix is reduced, and the heat transfer efficiency is improved; and the film cooling efficiency is remarkably improved.
Drawings
FIG. 1 is a schematic illustration of a turbine blade pressure side cooling structure arrangement in accordance with an embodiment of the present application;
FIG. 2 is a schematic diagram of a cooling unit according to one embodiment of the application;
fig. 3 is a cross-sectional view of a cooling unit according to one embodiment of the present application.
Wherein:
1-a cooling unit; 2-barrier ribs; 3-an impingement hole; 4-an impingement cavity; 5-turbulent flow column; 6-turbulent flow channel; 7-throttling the channel; 71-a throttle passage upstream section; 72-a throttle passage downstream section; 8-air film holes.
Detailed Description
In order to make the objects, technical solutions and advantages of the present application become more apparent, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the accompanying drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of the application. The embodiments described below by referring to the drawings are illustrative and intended to explain the present application and should not be construed as limiting the application. All other embodiments, which can be made by those skilled in the art based on the embodiments of the application without making any inventive effort, are intended to be within the scope of the application. Embodiments of the present application will be described in detail below with reference to the accompanying drawings.
In the description of the present application, it should be understood that the terms "center," "longitudinal," "lateral," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," and the like indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings, merely to facilitate describing the present application and simplify the description, and do not indicate or imply that the devices or elements being referred to must have a specific orientation, be configured and operated in a specific orientation, and therefore should not be construed as limiting the scope of the present application.
The application is described in further detail below with reference to fig. 1 to 3.
The application provides a turbine blade pressure surface cooling structure, which comprises: a plurality of cooling units 1.
Specifically, as shown in fig. 1, a plurality of cooling units 1 are arranged on the blade pressure surface, and in the height direction, the cooling units 1 are divided into a plurality of stages by partition ribs 2, and the cooling units 1 are provided with an impact zone, a cross flow zone, and an outlet section. As shown in fig. 2, the impact zone is provided with an impact hole 3 and an impact cavity 4 communicated with the impact hole 3; the cross flow area is provided with turbulent flow columns 5, turbulent flow channels 6 are arranged between the turbulent flow columns 5, and the turbulent flow channels 6 are communicated with the impact cavity 4; the outlet section is provided with the throttle post, the throttle post includes first throttle post and second throttle post, the upstream section of two wall adjacent of first throttle post and second throttle post is the concentric circle form, the downstream section is the expansion form, form annular throttle passageway upstream section 71 between two wall upstream sections of concentric circle form, form expansion type throttle passageway downstream section 72 between two wall downstream sections of expansion form, throttle passageway upstream section 71 constitutes throttle passageway 7 with throttle passageway downstream section 72 jointly, throttle passageway 7 communicates with vortex passageway 6, the air film hole 8 with throttle passageway 7 intercommunication has still been seted up to the outlet section.
In the turbine blade pressure surface cooling structure, as shown in fig. 3, cooling air flow inside the blade flows to the outside of the blade through the air film hole 8 after passing through the impact hole 3, the impact cavity 4, the turbulence channel 6 and the throttling channel 7 in sequence.
According to the turbine blade pressure surface cooling structure, a cooling unit 1 is composed of an impact area, a cross flow area and an outlet section, and cold air enters an impact cavity 4 in the cooling unit 1 through an impact hole 3 to form impact cooling on the inner wall; then, the cool air enters into the turbulent flow channels 6 between the turbulent flow columns 5, so that the heat exchange area is increased, the disturbance is enhanced, and the heat exchange is enhanced; finally, the cool air is discharged out of the blades through the air film holes 8 after passing through the throttling channel 7 of the outlet section, so as to form air film cooling on the outer wall surface.
In a preferred embodiment of the application, the cooling units 1 are arranged in 4 rows in sequence along the leading edge to the trailing edge of the blade. The number of units can be adjusted according to the cooling requirements in practical engineering applications. In the height direction, the cooling unit 1 is divided into three sections by 2 ribs 2, in order to reduce the risk of deformation during casting due to excessively long cores and to reduce manufacturing difficulties.
In the preferred embodiment of the present application, the spoiler posts 5 are diamond-shaped, and the spoiler posts 5 are staggered. In this embodiment, the air film hole 8 is an expansion type inclined hole, and the inclination angle is 45 degrees.
According to the turbine blade pressure surface cooling structure, the whole cooling unit 1 throttles at an outlet section, a throttling channel 7 comprises a throttling channel upstream section 71 and a throttling channel downstream section 72, the throttling channel upstream section 71 is an annular channel formed by concentric circular wall surfaces of two throttling column upstream sections, after passing through the throttling channel 7, the flow speed is obviously reduced, and finally, the throttling channel is discharged into a main channel through a gas film hole 8, so that slit type gas film cooling is formed on the outer surface of a blade, and the gas film cooling efficiency is effectively increased. Simultaneously, this design can effectively reduce back air film outflow speed, improves the air film and covers the effect, avoids high-speed air film outflow to mix with the gas and brings great aerodynamic loss, reduces turbine efficiency.
The turbine blade pressure surface cooling structure adopts the design of impact, cross flow and air film, so that the heat exchange strength of the cold air is enhanced, the residence time of the cold air in the blades is prolonged, and the utilization rate of the cold air is remarkably improved. Different from the composite cooling structure, the whole cooling channel is arranged inside the blade matrix, and the cooling channel is not required to be constructed through a conduit, so that the wall thickness between cold air and fuel gas can be obviously reduced, the heat conduction resistance is reduced, and the heat transfer efficiency is improved.
The turbine blade pressure surface cooling structure has the advantages of high comprehensive cooling effect, strong temperature bearing capacity, high cold air utilization rate, small cold air consumption and obviously improved air film cooling efficiency.
The foregoing is merely illustrative of the present application, and the present application is not limited thereto, and any changes or substitutions easily contemplated by those skilled in the art within the scope of the present application should be included in the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (7)
1. A turbine blade pressure side cooling structure, comprising:
cooling units (1), wherein a plurality of cooling units (1) are arranged on the blade pressure surface, the cooling units (1) are divided into a plurality of sections by partition ribs (2) in the height direction, the cooling units (1) are provided with an impact area, a cross flow area and an outlet section, wherein,
the impact area is provided with an impact hole (3) and an impact cavity (4) communicated with the impact hole (3);
the cross flow area is provided with turbulence columns (5), turbulence channels (6) are arranged between the turbulence columns (5), and the turbulence channels (6) are communicated with the impact cavity (4);
the outlet section is provided with a throttling column, the throttling column comprises a first throttling column and a second throttling column, the upstream sections of two adjacent wall surfaces of the first throttling column and the second throttling column are in concentric circles, the downstream sections are in expanded forms, an annular throttling channel upstream section (71) is formed between the two concentric wall surface upstream sections, an expanding throttling channel downstream section (72) is formed between the two wall surface downstream sections in the expanded forms, the throttling channel upstream section (71) and the throttling channel downstream section (72) jointly form a throttling channel (7), the throttling channel (7) is communicated with the turbulence channel (6), and an air film hole (8) communicated with the throttling channel (7) is formed in the outlet section;
the cooling air flow inside the blade sequentially passes through the impact hole (3), the impact cavity (4), the turbulence channel (6) and the throttling channel (7) and then flows to the outside of the blade through the air film hole (8).
2. Turbine blade pressure side cooling structure according to claim 1, characterized in that the cooling units (1) are arranged in 4 rows in sequence along the blade leading edge to the trailing edge.
3. Turbine blade pressure side cooling structure according to claim 1, characterized in that the cooling unit (1) is divided into three sections in height direction by two of the barrier ribs (2).
4. Turbine blade pressure side cooling structure according to claim 1, wherein the turbulator posts (5) are diamond-shaped, and a plurality of the turbulator posts (5) are staggered.
5. Turbine blade pressure side cooling structure according to claim 1, characterized in that the film holes (8) are of an expanding type.
6. Turbine blade pressure side cooling structure according to claim 5, characterized in that the film holes (8) are inclined holes.
7. Turbine blade pressure side cooling structure according to claim 6, characterized in that the angle of inclination of the film holes (8) is 45 degrees.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111342660.4A CN114109514B (en) | 2021-11-12 | 2021-11-12 | Turbine blade pressure surface cooling structure |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111342660.4A CN114109514B (en) | 2021-11-12 | 2021-11-12 | Turbine blade pressure surface cooling structure |
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| CN114109514A CN114109514A (en) | 2022-03-01 |
| CN114109514B true CN114109514B (en) | 2023-11-28 |
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Families Citing this family (2)
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| CN114575931B (en) * | 2022-03-16 | 2024-06-07 | 中国航发沈阳发动机研究所 | Turbine blade cooling structure with high temperature bearing capacity |
| CN114991880A (en) * | 2022-08-01 | 2022-09-02 | 中国航发沈阳发动机研究所 | Double-wall rotor blade of high-pressure turbine of aircraft engine |
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