CN114109514A - Turbine blade pressure surface cooling structure - Google Patents
Turbine blade pressure surface cooling structure Download PDFInfo
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- CN114109514A CN114109514A CN202111342660.4A CN202111342660A CN114109514A CN 114109514 A CN114109514 A CN 114109514A CN 202111342660 A CN202111342660 A CN 202111342660A CN 114109514 A CN114109514 A CN 114109514A
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- 238000001816 cooling Methods 0.000 title claims abstract description 71
- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 17
- 238000005192 partition Methods 0.000 claims description 6
- QNRATNLHPGXHMA-XZHTYLCXSA-N (r)-(6-ethoxyquinolin-4-yl)-[(2s,4s,5r)-5-ethyl-1-azabicyclo[2.2.2]octan-2-yl]methanol;hydrochloride Chemical compound Cl.C([C@H]([C@H](C1)CC)C2)CN1[C@@H]2[C@H](O)C1=CC=NC2=CC=C(OCC)C=C21 QNRATNLHPGXHMA-XZHTYLCXSA-N 0.000 claims 1
- 238000000034 method Methods 0.000 abstract 1
- 239000007789 gas Substances 0.000 description 8
- 238000004378 air conditioning Methods 0.000 description 6
- 239000002131 composite material Substances 0.000 description 3
- 238000012546 transfer Methods 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 238000013461 design Methods 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 230000015572 biosynthetic process Effects 0.000 description 1
- 238000005266 casting Methods 0.000 description 1
- 150000001875 compounds Chemical class 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000009826 distribution Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 230000003116 impacting effect Effects 0.000 description 1
- 230000014759 maintenance of location Effects 0.000 description 1
- 238000004519 manufacturing process Methods 0.000 description 1
- 239000011159 matrix material Substances 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 230000001737 promoting effect Effects 0.000 description 1
- 239000002356 single layer Substances 0.000 description 1
- 238000006467 substitution reaction Methods 0.000 description 1
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
The application belongs to the field of aeroengine blades, and particularly relates to a turbine blade pressure surface cooling structure. The method comprises the following steps: the cooling unit arranges a plurality ofly at the blade pressure face, and the cooling unit separates into the multistage through separating the 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 turbulence columns, turbulence channels are arranged between the turbulence columns, and the turbulence 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 walls of the first throttling column and the second throttling column are in the form of concentric circles, the downstream section is in an expansion form, an annular throttling channel upstream section is formed between two wall upstream sections in the form of concentric circles, an expansion-type throttling channel downstream section is formed between two wall downstream sections in the expansion form, the throttling channel upstream section and the throttling channel downstream section jointly form a throttling channel, the throttling channel is communicated with the turbulent flow channel, and a gas film hole is further formed in the outlet section.
Description
Technical Field
The application belongs to the field of aeroengine blades, and particularly relates to a turbine blade pressure surface cooling structure.
Background
With the development of aviation technology, the performance of an aero-engine is continuously improved, the temperature in front of a turbine is continuously increased, the temperature in front of the turbine is increased from 1700K level of a third-generation engine to 2000K level, and the heat load of turbine blades is greatly increased; meanwhile, the amount of cold air used by the turbine blades is continuously reduced in pursuit of improvement of engine efficiency. The composite cooling structure commonly adopted in the third and fourth generation machines cannot meet the cooling requirement of the turbine blade at the inlet temperature of more than 2000K.
What present the most compound cooling structure adopted is the cooling method of impact convection current + air film, and the hole formation of strikeing on the air conditioning through the pipe is to the impingement cooling of base member internal face, and then, flows through the air film hole on the blade, forms the air film cooling to the base member outer wall face. The composite cooling structure has the following disadvantages:
a) the comprehensive cooling capacity is basically limited by the consumption of cold gas and the processing capacity;
b) the utilization rate of the cold air is low, the consumption of the cold air is higher, the cold air is directly discharged from the air film hole after impacting the wall surface, the retention time of the cold air in the blade is short, and the temperature increase is limited;
c) the single-layer wall structure has thicker wall thickness and large heat conduction and heat resistance, and is not beneficial to cooling the blade;
d) circular air film holes are generally adopted, and the air film cooling efficiency is low.
Accordingly, a technical solution is desired to overcome or at least alleviate at least one of the above-mentioned drawbacks of the prior art.
Disclosure of Invention
The object of the present application is to provide a turbine blade pressure side cooling structure to solve at least one problem 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 a pressure surface of the blade, the cooling units being divided into a plurality of sections in a height direction by partition ribs, 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 turbulence columns, turbulence channels are arranged among the turbulence columns, and the turbulence 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 a concentric circle form, the downstream section is in an expansion form, an annular throttling channel upstream section is formed between the two wall surface upstream sections in the concentric circle form, an expansion type throttling channel downstream section is formed between the two wall surface downstream sections in the expansion form, the throttling channel upstream section and the throttling channel downstream section jointly form a throttling channel, the throttling channel is communicated with the turbulent flow channel, and the outlet section is also provided with a gas film hole communicated with the throttling channel;
the cooling air flow in the blade flows to the outside of the blade from 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 present application, the cooling units are arranged in 4 rows in sequence from 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 partition ribs.
In at least one embodiment of this application, the vortex post is the rhombus, a plurality of the vortex post staggered distribution.
In at least one embodiment of the present application, the gas film hole is flared.
In at least one embodiment of the present application, the film hole is an oblique hole.
In at least one embodiment of the present application, the inclination angle of the film holes is 45 degrees.
The invention has at least the following beneficial technical effects:
the cooling structure for the pressure surface of the turbine blade can greatly improve the comprehensive cooling capacity of the blade; the utilization rate of the cold air is improved, and the amount of the cold air is reduced; the heat conduction resistance of the matrix is reduced, and the heat transfer efficiency is improved; the air film cooling efficiency is obviously improved.
Drawings
FIG. 1 is a schematic view of a turbine blade pressure side cooling configuration arrangement according to an embodiment of the present application;
FIG. 2 is a schematic view of a cooling unit according to an embodiment of the present application;
FIG. 3 is a cross-sectional view of a cooling unit according to an embodiment of the present application.
Wherein:
1-a cooling unit; 2-ribs; 3-an impact hole; 4-an impingement chamber; 5-a turbulence column; 6-a turbulent flow channel; 7-a throttling channel; 71-the upstream section of the throttling channel; 72-a downstream section of the throttling channel; 8-air film hole.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the 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 a subset of the embodiments in the present application and not all embodiments in the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application and should not be construed as limiting the present application. 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 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 is to be understood that the terms "center", "longitudinal", "lateral", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are used merely for convenience in describing the present application and for simplifying the description, and do not indicate or imply that the referenced device or element must have a particular orientation, be constructed in a particular orientation, and be operated, and therefore should not be construed as limiting the scope of the present application.
The present application is described in further detail below with reference to fig. 1 to 3.
The application provides a turbine blade pressure side cooling structure, includes: a plurality of cooling units 1.
Specifically, as shown in fig. 1, a plurality of cooling units 1 are arranged on the pressure surface of the blade, the cooling units 1 are divided into a plurality of sections in the height direction by partition ribs 2, and the cooling units 1 are provided with an impingement area, a cross flow area, 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 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 walls of the first throttling column and the second throttling column are in the form of concentric circles, the downstream sections are in the form of expansion, an annular throttling channel upstream section 71 is formed between the two wall upstream sections in the form of concentric circles, an expansion-type throttling channel downstream section 72 is formed between the two wall downstream sections in the form of expansion, 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 turbulent flow channel 6, and the outlet section is further provided with a gas film hole 8 communicated with the throttling channel 7.
According to the cooling structure for the pressure surface of the turbine blade, as shown in fig. 3, cooling airflow inside the blade flows to the outside of the blade through the air film hole 8 after sequentially passing through the impact hole 3, the impact cavity 4, the turbulent flow channel 6 and the throttling channel 7.
According to the turbine blade pressure surface cooling structure, a cooling unit 1 consists of an impact area, a transverse flow area and an outlet section, and cold air enters an impact cavity 4 in the cooling unit 1 through impact holes 3 to form impact cooling on the inner wall; then, cold air enters the turbulence channels 6 between the turbulence columns 5, so that the heat exchange area is increased, the turbulence is enhanced, and the heat exchange is enhanced; finally, the cold air passes through a throttling channel 7 of the outlet section and then is discharged out of the blades through an air film hole 8, and air film cooling on the outer wall surface is formed.
In the preferred embodiment of the present application, the cooling units 1 are arranged in 4 rows in order from the leading edge to the trailing edge of the blade. The number of units can be adjusted according to the cooling requirement in practical engineering application. In the height direction, the cooling unit 1 is divided into three sections by 2 partition ribs 2, so that the risk of deformation during casting due to overlong cores is reduced, and the manufacturing difficulty is reduced.
In the preferred embodiment of the present application, the turbulence columns 5 are diamond-shaped, and a plurality of turbulence columns 5 are distributed in a staggered manner. In this embodiment, the air film hole 8 is an expanding inclined hole, and the inclination angle is 45 degrees.
The utility model provides a turbine blade pressure side cooling structure, whole cooling unit 1 is in the export section throttle, throttle passage 7 is including throttle passage upper reaches section 71 and throttle passage downstream section 72, throttle passage upper reaches section 71 is the annular passage that the concentric circular wall face of two throttle post upper reaches sections formed, the expanding type passageway that the expanding form wall face of two throttle post upper reaches sections formed, after through throttle passage 7, the velocity of flow can show the reduction, arrange into the main entrance through air film hole 8 at last, form seam formula air film cooling at the surface of blade, effectively increase air film cooling efficiency. Meanwhile, the design can effectively reduce the outflow speed of the gas film on the back side, improve the covering effect of the gas film, avoid the great pneumatic loss caused by the mixing of the outflow and the gas of the high-speed gas film and reduce the efficiency of the turbine.
The utility model provides a turbine blade pressure surface cooling structure, the design that has adopted impact + crossflow formula + air film formula has strengthened the heat transfer intensity of air conditioning, has prolonged the dwell time of air conditioning in the blade, is showing and is improving the air conditioning utilization ratio. Different from a composite cooling structure, the whole cooling channel is arranged in the blade base body, and the cooling channel is not required to be constructed through a guide pipe, 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 utility model provides a turbine blade pressure surface cooling structure synthesizes cold-efficiently high, and it can the reinforce to bear the temperature, and the air conditioning utilization ratio is high, and the air conditioning quantity is few, and air film cooling efficiency is showing and is promoting.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of 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:
a plurality of cooling units (1), the cooling units (1) being arranged on a blade pressure surface, the cooling units (1) being divided into a plurality of sections in a height direction by partition ribs (2), the cooling units (1) being provided with an impingement zone, a cross flow zone and an outlet section, wherein,
the impact area is provided with impact holes (3) and an impact cavity (4) communicated with the impact holes (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 a concentric circle form, the downstream section is in an expansion form, an annular throttling channel upstream section (71) is formed between the two wall surface upstream sections in the concentric circle form, an expansion type throttling channel downstream section (72) is formed between the two wall surface downstream sections in the expansion form, 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 turbulent flow channel (6), and the outlet section is further provided with a gas film hole (8) communicated with the throttling channel (7);
the cooling airflow in the blade flows to the outside of the blade from the air film hole (8) after sequentially passing through the impact hole (3), the impact cavity (4), the turbulent flow channel (6) and the throttling channel (7).
2. The turbine blade pressure surface cooling structure according to claim 1, wherein the cooling units (1) are arranged in 4 rows in order from the leading edge to the trailing edge of the blade.
3. The turbine blade pressure surface cooling structure according to claim 1, wherein the cooling unit (1) is divided into three sections in a height direction by two of the partition ribs (2).
4. The turbine blade pressure surface cooling structure according to claim 1, wherein the turbulence columns (5) are rhombic, and a plurality of the turbulence columns (5) are distributed in a staggered manner.
5. The turbine blade pressure side cooling structure according to claim 1, wherein the film hole (8) is of an expanding type.
6. The turbine blade pressure side cooling structure according to claim 5, wherein the film hole (8) is an inclined hole.
7. The turbine blade pressure side cooling structure according to claim 6, wherein the inclination angle of the film hole (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 |
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| CN202111342660.4A CN114109514B (en) | 2021-11-12 | 2021-11-12 | Turbine blade pressure surface cooling structure |
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| CN114109514A true CN114109514A (en) | 2022-03-01 |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114575931A (en) * | 2022-03-16 | 2022-06-03 | 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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| CN114109514B (en) | 2023-11-28 |
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