CN109736898A - A kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle - Google Patents
A kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle Download PDFInfo
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- CN109736898A CN109736898A CN201910026614.XA CN201910026614A CN109736898A CN 109736898 A CN109736898 A CN 109736898A CN 201910026614 A CN201910026614 A CN 201910026614A CN 109736898 A CN109736898 A CN 109736898A
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Abstract
The purpose of the present invention is to provide a kind of blade inlet edge gaseous film control pore structures of staggeredly compound angle, including first group of cylinder film cooling holes and second group of cylinder film cooling holes, the import of first group of cylinder film cooling holes and second group of cylinder film cooling holes is fixed on blade inner wall, outlet is fixed in blade outside wall surface, the location arrangements cold air import of leaf top, blade cold air ante-chamber is located at below cold air import, cooling gas enters the import that first group of cylinder film cooling holes and second group of cylinder film cooling holes are flowed into through blade cold air ante-chamber from cold air import, first group of cylinder film cooling holes and second group of cylinder film cooling holes are staggered.The present invention is preferably attached on blade inlet edge wall surface by increasing the cooling gas that compound angle appropriate flows out leading edge air film hole, the interlaced and opposite compound angle air film hole of two column advantageously forms anti-kidney shape whirlpool, prevents high-temperature fuel gas volume to be drawn on wall surface and is blown off with cold air.
Description
Technical field
The present invention relates to a kind of gas-turbine units, the specifically blade of gas-turbine unit.
Background technique
With the development of gas turbine technology, the thermal efficiency of cycle of gas turbine is continuously improved, and incident is higher
Compressor pressure ratio and turbine before fuel gas temperature.The raising speed of metal heat resisting temperature is far less than temperature before gas turbine turbine
Height-regulating speed, gas turbine cooling technology is at solving the problems, such as this key.
Gaseous film control is the most important external type of cooling in gas turbine cooling technology.Gaseous film control structure uses one
Or the cooling working medium jet stream of multiply, it is sprayed via the film cooling holes of specific direction, forms cold air film in blade surface, pass through cold air
With combustion gas blending come reduce blade impression fuel gas temperature.
Using since gaseous film control since the combustion gas turbine in 1972, with research deeply and the development of technology, air film
The type of cooling plays increasingly important role in the cooling of combustion gas turbine high-temperature component.The level-one rotor and stator blade of modern turbine
The gaseous film control structure of surface, shroud, leaf top and end wall universal placement of discrete well format.
Region near the front edge area of blade, especially stagnation line is the highest region of thermic load on blade, is most to need
Carry out the place of cooling protection.Meanwhile the flowing of this partial region again it is extremely complex, by mainstream stagnation, suppress force gradient,
The influence of many factors such as the interaction between vane curvature, jet hole row.
Summary of the invention
Before a kind of staggeredly blade of compound angle less the purpose of the present invention is to provide cold air flow and less hole number of rows
Edge gaseous film control pore structure.
The object of the present invention is achieved like this:
A kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle of the present invention, including blade, blade include in blade
Wall surface, blade outside wall surface, Ye Ding, blade root, it is characterized in that: further including that first group of cylinder film cooling holes and second group are cylindrical
The import of film cooling holes, first group of cylinder film cooling holes and second group of cylinder film cooling holes is fixed on blade inner wall
On face, outlet is fixed in blade outside wall surface, and leaf top location arrangements cold air import, blade cold air ante-chamber is located under cold air import
Side, cooling gas enters from cold air import flows into first group of cylinder film cooling holes and second group of cylinder through blade cold air ante-chamber
The import of shape film cooling holes, first group of cylinder film cooling holes and second group of cylinder film cooling holes are staggered.
The present invention may also include:
1, first group of cylinder film cooling holes and second group of cylinder film cooling holes and blade radial angle be 15 °~
90°。
2, the compound angle of first group of cylinder film cooling holes and second group of cylinder film cooling holes is contrary, compound
Angle is 10 °~45 °.
3, phase is distinguished in the aperture and pitch-row of first group of cylinder film cooling holes and second group of cylinder film cooling holes
Deng aperture d=0.6~1.2mm, pitch-row S=1~3d.
Present invention has an advantage that the cooling air that the present invention flows out leading edge air film hole by increasing compound angle appropriate
Body is preferably attached on blade inlet edge wall surface, and the interlaced and opposite compound angle air film hole of two column advantageously forms anti-kidney shape
Whirlpool prevents high-temperature fuel gas volume to be drawn on wall surface and is blown off with cold air, and the angles of the opposite center line of two column hole compound angles are less than or wait
Marginal pore outflow cooling gas well covers leading edge more uniform simultaneously to pressure face and suction surface point before can guaranteeing in 90 °
Stream.Two column holes replace common multiple row pore structure that can substantially reduce the cooling consumed air conditioning quantity of leading edge.
Detailed description of the invention
Fig. 1 is blade inner cavity main view；
Fig. 2 is blade inner cavity top view；
Fig. 3 is blade inner cavity right view；
Fig. 4 is blade schematic diagram；
Fig. 5 is the embodiment blade inner cavity main view on simplified model；
Fig. 6 is the embodiment blade inner cavity left view on simplified model；
Fig. 7 is the embodiment blade inner cavity top view on simplified model；
Fig. 8 a is that the embodiment leading edge on simplified model is insulated one of gas film cooling efficiency distribution map, and Fig. 8 b is simplified model
On embodiment leading edge insulation gas film cooling efficiency distribution map two；
Fig. 9 is marginal pore streamline distribution figure before embodiment 2.
Specific embodiment
It illustrates with reference to the accompanying drawing and the present invention is described in more detail:
In conjunction with Fig. 1-9, usually requires to arrange multiple rows of air film hole the present invention be directed to blade inlet edge part and consume larger stream
What the case where measuring cold air was designed.The present invention is a kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle, the air film hole knot
Structure includes two groups of staggered cylindrical film cooling holes 4, and two groups of film cooling holes imports 9 are separately fixed at blade inner wall
On 5, outlet 10 is fixed in blade outside wall surface 6, and cooling gas flows into air film hole by the preceding cold air cavity 3 of blade interior.Two groups cold
But it is opposite with main flow direction angle while hole 4 is identical as blade radial angle.
Embodiment 1 is certain type turbine blade of gas turbine, including wall surface 5,6 inside and outside blade and contains film cooling holes 4
Ante-chamber cooling structure.In conjunction with Fig. 1 to Fig. 4, in embodiment 1, cooling gas uses 7 air inlet of leaf top, flows into from blade cold air ante-chamber 3
Film cooling holes 4.Leading edge film cooling holes import 9 is located on blade inner wall 5, and outlet is located at blade outside wall surface 6, leading edge air film
It is i.e. of the present invention to flow to angle with the formation of mainstream flow direction for the center line and the linear radially angle of blade inlet edge of cooling hole
Compound angle.Two column film cooling holes 4 are in being staggered, their radial angle is identical, and compound angle size is identical, contrary.
The diameter and pitch-row of two column film cooling holes 4 are equal, respectively d=0.8mm, S=d, film cooling holes 4 and blade radial angle
It is 45 °, compound angle is 15 °.
Blade inlet edge region thermic load and bigger, the film cooling holes 4 that the present invention arranges staggeredly compound angle by two of drying
Design can effectively improve the docile degree and coverage area of Gas Cooling, before reducing while improving gas film cooling efficiency
4 number of rows of edge cooling hole and cold air flow needed for greatly reducing leading edge gaseous film control.
Example 2 is blade simplified model, and such as Fig. 5 to Fig. 7, cooling gas uses 11 air inlets of top, from simplified model cold air
Ante-chamber flows into film cooling holes.Leading edge film cooling holes import is located on simplified model inner wall, and outlet 15 is located at simplified model
Outside wall surface, the center line and costa of 16 film cooling holes of leading edge form radial angle, form flow direction folder with mainstream flow direction
Angle, that is, compound angle of the present invention.First group of film cooling holes, 13, second groups of film cooling holes 14 are in be staggered, their diameter
Identical to angle, compound angle size is identical, contrary.The diameter and pitch-row of two column film cooling holes are equal, respectively d=
0.8mm, S=d, film cooling holes and blade radial angle are 45 °, and compound angle is 15 °.If Fig. 8 (a), (b) are respectively example 2
Marginal pore scheme simplified model front edge area air film is cold before three rows of simplified model front edge area gas film cooling efficiency distribution map and tradition
But the distribution map of efficiency, Fig. 9 are marginal pore streamline distribution figure before example 2.It can be seen from the figure that only two rows of leading edges in example 2
In the case where hole, leading edge insulation gas film cooling efficiency is very high and distribution is more uniform.Stream needed for 2 leading edge film cooling holes of embodiment
Amount is about 38% of flow needed for three rounds design leading edge film cooling holes, greatly reduces the consumption of cold air.
Claims (5)
1. a kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle, including blade, blade include blade inner wall, blade
Outside wall surface, Ye Ding, blade root, it is characterized in that: further including first group of cylinder film cooling holes and second group of cylinder gaseous film control
The import in hole, first group of cylinder film cooling holes and second group of cylinder film cooling holes is fixed on blade inner wall, out
Mouth is fixed in blade outside wall surface, the location arrangements cold air import of leaf top, and blade cold air ante-chamber is located at below cold air import, cooling air
Body enters cold through blade cold air ante-chamber first group of cylinder film cooling holes of inflow and second group of cylinder air film from cold air import
But the import in hole, first group of cylinder film cooling holes and second group of cylinder film cooling holes are staggered.
2. the blade inlet edge gaseous film control pore structure of a kind of staggeredly compound angle according to claim 1, it is characterized in that: first
The cylindrical film cooling holes of group and second group of cylinder film cooling holes and blade radial angle are 15 °~90 °.
3. a kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle according to claim 1 or 2, it is characterized in that:
The compound angle of first group of cylinder film cooling holes and second group of cylinder film cooling holes is contrary, compound angle be 10 °~
45°。
4. a kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle according to claim 1 or 2, it is characterized in that:
Equal, aperture d=is distinguished in the aperture and pitch-row of first group of cylinder film cooling holes and second group of cylinder film cooling holes
0.6~1.2mm, pitch-row S=1~3d.
5. the blade inlet edge gaseous film control pore structure of a kind of staggeredly compound angle according to claim 3, it is characterized in that: first
The aperture and pitch-row difference of the cylindrical film cooling holes of group and second group of cylinder film cooling holes are equal, and aperture d=0.6~
1.2mm, pitch-row S=1~3d.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910026614.XA CN109736898A (en) | 2019-01-11 | 2019-01-11 | A kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201910026614.XA CN109736898A (en) | 2019-01-11 | 2019-01-11 | A kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle |
Publications (1)
| Publication Number | Publication Date |
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| CN109736898A true CN109736898A (en) | 2019-05-10 |
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| CN201910026614.XA Pending CN109736898A (en) | 2019-01-11 | 2019-01-11 | A kind of blade inlet edge gaseous film control pore structure of staggeredly compound angle |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113279818A (en) * | 2021-06-24 | 2021-08-20 | 中国科学院工程热物理研究所 | Contraction type double-jet air film hole |
| CN116950723A (en) * | 2023-09-19 | 2023-10-27 | 中国航发四川燃气涡轮研究院 | Low-stress double-wall turbine guide vane cooling structure and design method thereof |
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| CN105673089A (en) * | 2016-03-31 | 2016-06-15 | 中国船舶重工集团公司第七�三研究所 | Crown-free air film cooling rotor blade for turbine of gas turbine |
| US20180223672A1 (en) * | 2017-02-07 | 2018-08-09 | General Electric Company | Investment casting core |
| CN108425705A (en) * | 2018-01-23 | 2018-08-21 | 中国科学院工程热物理研究所 | A kind of cooling of double wall and gaseous film control combined type turbine blade structure |
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2019
- 2019-01-11 CN CN201910026614.XA patent/CN109736898A/en active Pending
Patent Citations (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5486093A (en) * | 1993-09-08 | 1996-01-23 | United Technologies Corporation | Leading edge cooling of turbine airfoils |
| GB9604652D0 (en) * | 1996-03-05 | 1996-05-01 | Rolls Royce Plc | fluid cooled wall |
| US5779437A (en) * | 1996-10-31 | 1998-07-14 | Pratt & Whitney Canada Inc. | Cooling passages for airfoil leading edge |
| CN1724849A (en) * | 2004-07-05 | 2006-01-25 | 西门子公司 | Turbine blade |
| CN1818349A (en) * | 2005-01-28 | 2006-08-16 | 通用电气公司 | Rectangular diffusion hole for film cooled turbine airfoil |
| US7540712B1 (en) * | 2006-09-15 | 2009-06-02 | Florida Turbine Technologies, Inc. | Turbine airfoil with showerhead cooling holes |
| US8317473B1 (en) * | 2009-09-23 | 2012-11-27 | Florida Turbine Technologies, Inc. | Turbine blade with leading edge edge cooling |
| CN102425459A (en) * | 2011-11-21 | 2012-04-25 | 西安交通大学 | Heavy-type combustion engine high-temperature turbine double-medium cooling blade |
| US8858176B1 (en) * | 2011-12-13 | 2014-10-14 | Florida Turbine Technologies, Inc. | Turbine airfoil with leading edge cooling |
| CN103806952A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Turbine blade with leading-edge concaved cavity |
| 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 |
| CN105673089A (en) * | 2016-03-31 | 2016-06-15 | 中国船舶重工集团公司第七�三研究所 | Crown-free air film cooling rotor blade for turbine of gas turbine |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113279818A (en) * | 2021-06-24 | 2021-08-20 | 中国科学院工程热物理研究所 | Contraction type double-jet air film hole |
| CN116950723A (en) * | 2023-09-19 | 2023-10-27 | 中国航发四川燃气涡轮研究院 | Low-stress double-wall turbine guide vane cooling structure and design method thereof |
| CN116950723B (en) * | 2023-09-19 | 2024-01-09 | 中国航发四川燃气涡轮研究院 | Low-stress double-wall turbine guide vane cooling structure and design method thereof |
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