CN203214108U - Turbine blade front edge cooling structure - Google Patents
Turbine blade front edge cooling structure Download PDFInfo
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- CN203214108U CN203214108U CN 201320100569 CN201320100569U CN203214108U CN 203214108 U CN203214108 U CN 203214108U CN 201320100569 CN201320100569 CN 201320100569 CN 201320100569 U CN201320100569 U CN 201320100569U CN 203214108 U CN203214108 U CN 203214108U
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- conical
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Abstract
A turbine blade front edge cooling structure comprises a tenon, a margin plate, a blade body, a blade front edge, a blade tail edge, and a blade interior cooling cavity. The turbine blade front edge cooling structure is characterized in that a cylindrical cooling hole and a conical cooling hole are further included, the blade body is axial-flow type blade molded surface, a blade front edge air membrane cooling air membrane hole structure is formed by the cylindrical cooling hole and the conical cooling hole in a combination mode. The axis of the cylindrical cooling hole and the axis of the conical cooling hole are located on the same line, an included angle alpha formed by the axis of an air membrane hole and the outer wall face of a blade is smaller than 60 degrees, the taper ratio of the conical air membrane hole, namely the value of the outlet area of the air membrane hole divided by the cylindrical cooling hole is larger than 1.3, the diameter of the cylindrical cooling hole is larger than 0.25mm, and the effective length of the cylindrical cooling hole is larger than 0.3mm. The turbine blade front edge cooling structure has the advantages that the conical air membrane hole is adopted at the front edge of a turbine blade, so that the cooling effect of a blade front edge air membrane is improved, and the high-temperature turbine blade front edge cooling is improved.
Description
Technical field
The utility model relates to gas-turbine unit turbine cooled blade structure field, particularly a kind of turbine blade leading edge cooling structure.
Background technique
Because advanced high-temperature fuel gas turbogenerator fuel gas temperature is higher, turbine blade need adopt advanced efficient cooling technology, particularly turbine blade leading edge, because impression is the combustion gas stagnation temperature, it is become experience the highest zone of fuel gas temperature in the turbine blade.For turbine blade intensity life-span and functional reliability are met the demands, must efficiently cool off blade inlet edge.
The high-temperature fuel gas turbogenerator turbine blade type of cooling mainly is that the blade inner chamber adopts efficient forced convection cooling, namely impacts cooling, and the air film cooling is adopted on blade combustion gas side runner surface.The cooling of the forced convection of blade inner chamber makes the cooling air in the blade inner chamber can high efficiency absorption combustion gas conduct heat to blade, reduces the blade wall surface temperature; The air film cooling is adopted on blade combustion gas side runner surface, is to form one deck cooling air film at blade outer surface, and high-temperature fuel gas and blade wall " isolation " are opened, and significantly reduces combustion gas to the heat transfer of blade, reduces the blade wall surface temperature.
The high-temperature fuel gas turbine blade must adopt the air film cooling can make the blade wall surface temperature be cooled to spendable temperature, and fuel gas temperature is more high, and the effect of air film cooling is more important.At present, the film cooling holes in high-temperature fuel gas turbogenerator turbine blade leading edge zone all adopts cylindrical structural.Easy to process, the low cost of manufacture of cylindrical air film hole, but the cooling air more easily breaks away from the blade inlet edge surface, the turbine blade leading edge that weakened air film cooling effect when cylindrical air film hole enters the combustion gas runner; Because turbine blade leading edge curvature is bigger, air film covers difficulty, and cylindrical air film hole air film overlay area is less, and therefore, when the turbine blade leading edge adopted cylindrical air film hole, the air film cooling effect was lower.Because high-temperature fuel gas turbogenerator fuel gas temperature height, turbine blade leading edge need adopt more effective air film cooling, to improve blade air film cooling effect, reduce the blade wall surface temperature.
The model utility content
The purpose of this utility model is in order to improve the air film cooling effect that the cylindrical air film hole of turbine blade leading edge exists, and the spy provides a kind of turbine blade leading edge cooling structure.
The utility model provides a kind of turbine blade leading edge cooling structure, it is characterized in that: described turbine blade leading edge cooling structure comprises tenon 1, listrium 2, blade 3, blade inlet edge 4, blade trailing edge 5, blade interior cooling chamber 6, cylindrical cooling hole 7 and conical cooling hole 8;
Blade is the axial flow airfoil surface, and high-temperature fuel gas stream enters blade grid passage from blade inlet edge 4, flows out blade grid passages around leaf basin and the mobile back of blade back of airfoil surface from blade trailing edge 5, and turbine blade is wrapped in the high-temperature fuel gas stream.The cooling air that draws from the motor gas compressor enters blade interior cooling chamber 6 from cooling air inlet hole, blade tenon 1 bottom, then enters the combustion gas runner by the blade inlet edge air film hole.
By cylindrical cooling hole 7 and the conical cooling hole 8 compound blade inlet edge air film cooling air membrane pore structures of forming, the axis in cylindrical cooling hole 7 and conical cooling hole 8 is located along the same line, air film hole axis and blade outer wall angle α＜60 °, the taper ratio of conical air film hole, it is the value 1.3 of air film hole discharge area/cylindrical cooling hole area, cylindrical cooling bore dia is greater than 0.25mm, and the effective length in cylindrical cooling hole is greater than 0.3mm.The cooling air enters cylindrical cooling hole 7 by blade interior cooling chamber 6, then enters the combustion gas runner by taper shape cooling hole 8, and the intake and exhaust stream schematic representation of turbine blade leading edge cooling air as shown in Figure 1.
Cylindrical cooling hole 7 in the air film hole and conical cooling hole 8 are used for control by the cooling air delivery of air film hole; Flow velocity when air film hole is discharged for reducing the cooling air in cylindrical cooling hole 7 and conical cooling hole 8 increases the cooling air film in the area coverage of blade inlet edge, improves the air film cooling effect.
Advantage of the present utility model:
Turbine blade leading edge cooling structure described in the utility model by adopting conical air film hole in the turbine blade leading edge, improves the validity of blade inlet edge air film cooling, improves the cooling of high-temperature turbine blade inlet edge.
Description of drawings
Below in conjunction with drawings and the embodiments the utility model is described in further detail:
Fig. 1 is turbine blade leading edge cold air intake and exhaust stream schematic representation.
Embodiment
Embodiment 1
Present embodiment provides a kind of turbine blade leading edge cooling structure, it is characterized in that: described turbine blade leading edge cooling structure comprises tenon 1, listrium 2, blade 3, blade inlet edge 4, blade trailing edge 5, blade interior cooling chamber 6, cylindrical cooling hole 7 and conical cooling hole 8;
Blade is the axial flow airfoil surface, and high-temperature fuel gas stream enters blade grid passage from blade inlet edge 4, flows out blade grid passages around leaf basin and the mobile back of blade back of airfoil surface from blade trailing edge 5, and turbine blade is wrapped in the high-temperature fuel gas stream.The cooling air that draws from the motor gas compressor enters blade interior cooling chamber 6 from cooling air inlet hole, blade tenon 1 bottom, then enters the combustion gas runner by the blade inlet edge air film hole.
By cylindrical cooling hole 7 and the conical cooling hole 8 compound blade inlet edge air film cooling air membrane pore structures of forming, the axis in cylindrical cooling hole 7 and conical cooling hole 8 is located along the same line, air film hole axis and blade outer wall angle α＜60 °, the taper ratio of conical air film hole, it is the value 1.3 of air film hole discharge area/cylindrical cooling hole area, cylindrical cooling bore dia is greater than 0.25mm, and the effective length in cylindrical cooling hole is greater than 0.3mm.The cooling air enters cylindrical cooling hole 7 by blade interior cooling chamber 6, then enters the combustion gas runner by taper shape cooling hole 8, and the intake and exhaust stream schematic representation of turbine blade leading edge cooling air as shown in Figure 1.
Cylindrical cooling hole 7 in the air film hole and conical cooling hole 8 are used for control by the cooling air delivery of air film hole; Flow velocity when air film hole is discharged for reducing the cooling air in cylindrical cooling hole 7 and conical cooling hole 8 increases the cooling air film in the area coverage of blade inlet edge, improves the air film cooling effect.
Claims (1)
1. turbine blade leading edge cooling structure, comprise tenon (1), listrium (2), blade (3), blade inlet edge (4), blade trailing edge (5), blade interior cooling chamber (6), it is characterized in that: described turbine blade leading edge cooling structure also comprises, cylindrical cooling hole (7) and conical cooling hole (8);
Blade (3) is the axial flow airfoil surface, by the compound blade inlet edge air film cooling air membrane pore structure of forming in cylindrical cooling hole (7) and conical cooling hole (8), the axis in cylindrical cooling hole (7) and conical cooling hole (8) is located along the same line, air film hole axis and blade outer wall angle α＜60 °, the taper ratio of conical air film hole, be that the value of air film hole discharge area/cylindrical cooling hole area is greater than 1.3, cylindrical cooling bore dia is greater than 0.25mm, and the effective length in cylindrical cooling hole is greater than 0.3mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 201320100569 CN203214108U (en) | 2013-03-05 | 2013-03-05 | Turbine blade front edge cooling structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN 201320100569 CN203214108U (en) | 2013-03-05 | 2013-03-05 | Turbine blade front edge cooling structure |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN203214108U true CN203214108U (en) | 2013-09-25 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN 201320100569 Expired - Fee Related CN203214108U (en) | 2013-03-05 | 2013-03-05 | Turbine blade front edge cooling structure |
Country Status (1)
| Country | Link |
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| CN (1) | CN203214108U (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103806953A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Novel air film holes with inclination angles |
| CN104832218A (en) * | 2015-04-20 | 2015-08-12 | 西北工业大学 | Staggered opposite jetting air film hole row structure used for turbine blade leading edge air film cooling |
| CN117489418A (en) * | 2023-12-28 | 2024-02-02 | 成都中科翼能科技有限公司 | Turbine guide vane and cold air guide piece of front cold air cavity thereof |
-
2013
- 2013-03-05 CN CN 201320100569 patent/CN203214108U/en not_active Expired - Fee Related
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103806953A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Novel air film holes with inclination angles |
| CN104832218A (en) * | 2015-04-20 | 2015-08-12 | 西北工业大学 | Staggered opposite jetting air film hole row structure used for turbine blade leading edge air film cooling |
| CN117489418A (en) * | 2023-12-28 | 2024-02-02 | 成都中科翼能科技有限公司 | Turbine guide vane and cold air guide piece of front cold air cavity thereof |
| CN117489418B (en) * | 2023-12-28 | 2024-03-15 | 成都中科翼能科技有限公司 | Turbine guide vane and cold air guide piece of front cold air cavity thereof |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| CF01 | Termination of patent right due to non-payment of annual fee | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20130925 |