CN103806953A - Novel air film holes with inclination angles - Google Patents
Novel air film holes with inclination angles Download PDFInfo
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- CN103806953A CN103806953A CN201410026121.3A CN201410026121A CN103806953A CN 103806953 A CN103806953 A CN 103806953A CN 201410026121 A CN201410026121 A CN 201410026121A CN 103806953 A CN103806953 A CN 103806953A
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- air film
- film hole
- film holes
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- blade
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Abstract
The invention discloses a cooling structure applicable to aero-engine turbine blades and provides novel air film holes with inclination angles. The novel air film holes comprise front air film holes located in the front portion of a blade basin, middle air film holes located in the middle portion of the blade basin and tip air film holes located in the tips of the blades. The novel air film holes are characterized in that included angles ranging from 20 degrees to 50 degrees are formed between the front air film cooling holes and the normal lines of the surfaces of the blades in the flowing direction. The included angles ranging from minus 45 degrees to 45 degrees are formed between the middle air film holes and an S1 surface. The included angles ranging from zero degree to 60 degrees is formed between the tip air film holes and the S1 surface. Angular jet flow generated from the novel air film holes enables the surfaces of the blades to be covered with cooling air more reasonably, flow resistance is reduced, and good cooling effect and high flowing efficiency are achieved.
Description
Technical field
The present invention relates to one and be applicable to aero engine turbine blades cooling structure, this structure can meet the requirement of High Performance Aeroengine of new generation pair and high-temperature component cooling performance.
Background technique
Aviation industry has obtained swift and violent development in recent years, and improving as the performance of the aeroengine of power source is the most important thing.Analyzing the thermodynamic cycle process of aero-turbine can find, improving turbine inlet temperature is the technical measures that improves aero-engine performance.According to calculating, 55 ℃ of the every raisings of turbine inlet temperature, under the constant condition of size of engine, motor power approximately can improve 10%.The nineties in 20th century, the turbine inlet temperature of thrust weight ratio 10 one-level turbogenerators was 1800～1950K, before the turbine of the E3 motor of the U.S., maximum temperature has reached 2012K, " advanced core engine ordnance engine " plan that " combination property turbogenerator technology " plan that U.S. Department of Defense carries out and Britain carry out is all decided to be thrust weight ratio 20 target of the year two thousand twenty, realizes turbine inlet temperature and reaches 2273～2473K.Which kind of in the face of the series of problems that high temperature like this brings, must need to consider to take technology to guarantee the work that turbine blade and the turbine disk can be safe and reliable under hot environment.Practice shows, topmost means have two kinds: development of new material on the one hand, relies on the heat resistance that improves material self to guarantee hot-end component intensity index at high temperature; Adopt on the other hand advanced cooling technology to carry out hot-end component effectively cooling, make hot-end component still keep lower operating temperature.
Air film cooling technology is as one of Major Cooling Methods of high-temperature unit of aircraft engine, aspect effective reduction turbine blade surface temperature and thermal stress, playing an important role, it is carried out to lasting and deep research is the important channel of improving aero-engine performance.In the past few decades, near mainly concentrating on irregularly-shaped hole and outlet about the research work of hole shape, the air film hole of turbulence structure is set, these structures, in improving Film Cooling, have also been brought the flow resistance can not be ignored, and part irregularly-shaped hole mechanical processing difficulty is large, realizability is poor.
The present invention proposes a kind of Novel air fenestra with angle, by air film hole is increased to inclination angle, obtains better cooling effect.Because the jet with angle of this structure generation can make the covering of blade surface cold air more reasonable, reduce flow resistance simultaneously, on cooling and flow efficiency, obtained good effect simultaneously.
Summary of the invention
The object of this invention is to provide a kind of air film cooling structure with angle for aero engine turbine blades.It comprises: be positioned at leaf basin front portion anterior air film hole (1), be positioned at leaf basin middle part middle part air film hole (2), be positioned at the tip air film hole (3) of blade tip.It is characterized in that: be 1.0mm～1.5mm at anterior anterior air film hole (1) diameter of arranging of leaf basin, the exhibition of every exhaust fenestra is 10～20 to number, with S1 face angle (α 1) be-45 °～45 °, flowing to blade face normal angle (β 1) edge is 20 °～50 °, and exhibition is 5～10D to spacing; Middle part air film hole (2) diameter of arranging at leaf basin middle part is 1.0mm～2.0mm, the exhibition of every exhaust fenestra is 20～30 to number, with S1 face angle (α 2) be-45 °～45 °, to being 20 °～50 °, exhibition is 1～2D to spacing with blade face normal angle (β 2) adverse current; Leaf basin tip arrange tip air film hole (3) diameter be 1.0mm～1.5mm, the exhibition of every exhaust fenestra is 10～20 to number, with S1 face angle (α 3) be 0 °～60 °, with blade face normal angle (β 3) along flow to be 0 °～60 °.Anterior air film hole (1) is 1/8～1/4 of width of blade (W) to the distance (P1) of blade inlet edge, middle part air film hole (2) is 2/5～3/5 of leaf wide (W) to the distance (P2) of blade inlet edge, and tip air film hole (3) is 1/15～1/20 of chord length (L) to the distance (P3) of blade tip.
The advantage of air film cooling structure of the present invention is: (1) connection and reasonable arrangement, and reliable in structure, simple in processing; (2) form the cold air jet with angle, make to cover more even for the cold air of blade surface; (3) reverse cold air jet and the cold air jet to blade tip can effectively reduce flow losses.
Accompanying drawing explanation
The layout schematic diagram of Fig. 1 air film hole
The angle schematic diagram of the anterior air film hole of Fig. 2
The angle schematic diagram of Fig. 3 middle part air film hole
The angle schematic diagram of Fig. 4 tip air film hole
In figure: 1. anterior air film hole 2. middle part air film hole 3. tip air film hole 4. outer gas streams
5.S1 stream interface 6. blade surface normals
Embodiment
Below in conjunction with accompanying drawing, the present invention is described in further detail.
Referring to shown in Fig. 1 to Fig. 4, the present invention is a kind of air film cooling structure with angle for aero engine turbine blades.Wherein air film hole structure with angle is the major character that is different from other conventional air film hole.
Shown in Fig. 2,3,4; cooled gas is respectively from anterior air film hole (1) with angle, middle part air film hole (2), tip air film hole (3) ejection; form peripheric jet flow; picture downstream bending under the pressure of outer gas stream (4) and rubbing action; stick near wall; the cold air film that formation temperature is lower, by the outer gas stream of high temperature (4) and blade surface isolation, thereby has avoided high-temperature gas to inject region and downstream area surface to the ablation protection of leaf table.The cooling-air that the present invention's air film hole with angle blows out and outer gas stream (4) jet at an angle, make the more even of cooled gas covering, thereby make cooling effectiveness higher.
The present invention has passed through the checking of Three-dimensional simulation and experiment, and under specific blowing ratio, its cooling effectiveness can improve 80%～150% than the cooling effectiveness in conventional film hole, and cold air covers wider more even.
Claims (1)
1. the air film cooling structure with angle for aero engine turbine blades, it comprises: be positioned at leaf basin front portion anterior air film hole (1), be positioned at leaf basin middle part middle part air film hole (2), be positioned at the tip air film hole (3) of leaf basin tip, it is characterized in that: be 1.0mm～1.5mm at anterior anterior air film hole (1) diameter of arranging of leaf basin, the exhibition of every exhaust fenestra is 10～20 to number, with S1 face angle (α 1) be-45 °～45 °, flowing to blade face normal angle (β 1) edge is 20 °～50 °, and exhibition is 5～10D to spacing; Middle part air film hole (2) diameter of arranging at leaf basin middle part is 1.0mm～2.0mm, the exhibition of every exhaust fenestra is 20～30 to number, with S1 face angle (α 2) be-45 °～45 °, to being 20 °～50 °, exhibition is 1～2D to spacing with blade face normal angle (β 2) adverse current; Leaf basin tip arrange tip air film hole (3) diameter be 1.0mm～1.5mm, the exhibition of every exhaust fenestra is 10～20 to number, with S1 face angle (α 3) be 0 °～60 °, with blade face normal angle (β 3) along flow to be 0 °～60 °.Anterior air film hole (1) is 1/8～1/4 of width of blade (W) to the distance (P1) of blade inlet edge, middle part air film hole (2) is 2/5～3/5 of leaf wide (W) to the distance (P2) of blade inlet edge, and tip air film hole (3) is 1/15～1/20 of chord length (L) to the distance (P3) of blade tip.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410026121.3A CN103806953A (en) | 2014-01-20 | 2014-01-20 | Novel air film holes with inclination angles |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410026121.3A CN103806953A (en) | 2014-01-20 | 2014-01-20 | Novel air film holes with inclination angles |
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| Publication Number | Publication Date |
|---|---|
| CN103806953A true CN103806953A (en) | 2014-05-21 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201410026121.3A Pending CN103806953A (en) | 2014-01-20 | 2014-01-20 | Novel air film holes with inclination angles |
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Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20060210399A1 (en) * | 2003-11-21 | 2006-09-21 | Tsuyoshi Kitamura | Turbine cooling vane of gas turbine engine |
| CN202417612U (en) * | 2011-12-27 | 2012-09-05 | 中航商用航空发动机有限责任公司 | Turbine guide blade |
| CN102678189A (en) * | 2011-12-13 | 2012-09-19 | 河南科技大学 | Turbine cooling blade with blade tip leakage prevention structure |
| CN103277145A (en) * | 2013-06-09 | 2013-09-04 | 哈尔滨工业大学 | Cooling blade of gas turbine |
| CN203214108U (en) * | 2013-03-05 | 2013-09-25 | 中国航空工业集团公司沈阳发动机设计研究所 | Turbine blade front edge cooling structure |
-
2014
- 2014-01-20 CN CN201410026121.3A patent/CN103806953A/en active Pending
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20060210399A1 (en) * | 2003-11-21 | 2006-09-21 | Tsuyoshi Kitamura | Turbine cooling vane of gas turbine engine |
| CN102678189A (en) * | 2011-12-13 | 2012-09-19 | 河南科技大学 | Turbine cooling blade with blade tip leakage prevention structure |
| CN202417612U (en) * | 2011-12-27 | 2012-09-05 | 中航商用航空发动机有限责任公司 | Turbine guide blade |
| CN203214108U (en) * | 2013-03-05 | 2013-09-25 | 中国航空工业集团公司沈阳发动机设计研究所 | Turbine blade front edge cooling structure |
| CN103277145A (en) * | 2013-06-09 | 2013-09-04 | 哈尔滨工业大学 | Cooling blade of gas turbine |
Non-Patent Citations (1)
| Title |
|---|
| 林宇震 等: "《不同偏角多斜孔壁气膜冷却绝热温比研究》", 《推进技术》, vol. 19, no. 5, 31 October 1998 (1998-10-31), pages 43 - 46 * |
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Application publication date: 20140521 |