CN101126324A - Combined cooling structure for turbine blade middle-part inclined impact aerating film - Google Patents
Combined cooling structure for turbine blade middle-part inclined impact aerating film Download PDFInfo
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- CN101126324A CN101126324A CNA2007101187655A CN200710118765A CN101126324A CN 101126324 A CN101126324 A CN 101126324A CN A2007101187655 A CNA2007101187655 A CN A2007101187655A CN 200710118765 A CN200710118765 A CN 200710118765A CN 101126324 A CN101126324 A CN 101126324A
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Abstract
The invention discloses an inclined impact and air film combination cooling structure used at the middle of the turbine blades of an aeroengine. The cooling structure is that: air film holes with the diameter of 1.0 mm-1.5 mm are arranged on the outer surface of the blade, the amount of the air film holes at each extension direction is 10-20, inclined impact holes are arranged inside the blades in the downstream area of the air film holes, the amount of the inclined impact holes equals to that of the air film holes, the angle of the impact holes is 30-45 degrees. The larger impact torque is generated through the inclined impact holes and a highly cooling area with the large area is formed inside the blades; at the same times, the air film holes outside the blades form an air film protection area to jointly realize the blade-cooling purpose. The results of the model test and the three-dimensional numerical simulation indicate that the cooling effect of the blades can reach 0.7 at least; at the same time, the pneumatic loss can be reduced distinctly and the flow resistance is distinctly lower than the ordinary turbine blades owing to the characteristic of cooling structure.
Description
Technical field
The present invention relates to a kind of combined cooling structure of inclined impact aerating film, this cooling structure is mainly used in the middle part of aero engine turbine blades, can produce the cooling effect more than 0.7, satisfies the requirement of aero engine turbine blades cooling.
Background technique
The main performance index of motor are thrust weight ratios, along with people improve constantly the performance requirement of motor, the requirement of thrust weight ratio are also improved constantly.And the most effective means that thrust weight ratio adopted that improve motor improve the preceding fuel gas temperature of turbine exactly.The turbine inlet temperature of the thrust weight ratio 10 one-level aeroengines of China's beforehand research is about 1850K～1950K.And the various materials that use at present can only just can be kept its higher intensity index about 1300 ℃ under non-refrigerated situation.Turbine rotor under hot environment can be safe and reliable work, depend primarily on the temperature levels and the temperature distribution of each heating part (turbine blade, the turbine disk, axle etc.) in the rotor.In addition, because turbine blade (working blade) is in the middle of the very high centrifugal field work under the high rotating speed (changeing more than the scooter 15000rpm).In bad working environment like this, guarantee the work that blade is normal, reliable, long-term, just must effectively cool off turbine blade, guarantee that the blade self-temperature is under operating temperature, high again creep rupture strength and anti-corrosion capacity, the least possible use cooled gas when guaranteeing reliably working.Therefore, invent efficiently that cooling structure is very important, also be very important.Mostly designed conventional turbine blade is the fin at the internal placement different shape of blade at present, be used for increasing inner disturbance, improve the heat exchange effect, and usually arrange the air film hole that some diameters are less at the outer surface of blade, forming full air film covers, She Ji turbine blade in this way, its cooling effect is generally about 0.5, raising along with fuel gas temperature before the turbine, such cooling effect is significantly not enough concerning the blade cooling, so the invention of turbine cooling structure efficiently is very important and urgent.
Summary of the invention
The impact opening that the objective of the invention is to tilt combines with air film hole, and a kind of cooling structure that is applicable to aero engine turbine blades is provided.This cooling structure is: the outer surface cloth at blade is equipped with the air film hole that diameter is 1.0mm～1.5mm, the exhibition of every exhaust fenestra to number be 10～15, downstream area at the blade interior air film hole is furnished with the inclined impact hole identical with air film hole quantity, the impact opening angle is 30 degree～45 degree, the direction of the center line deflection air film hole of impact opening, impact opening is arranged in the downstream of corresponding air film hole, with the distance of air film hole be 8～10 times of impact opening diameter, and impact opening and air film hole are staggered, the span distance of impact opening and air film hole is 7～9 times of impact opening diameter, the diameter of air film hole is 2～2.5 times of impact opening diameter, and the exhibition of air film hole is 4～5 times of the air film hole diameter to spacing.。
The advantage of the combined cooling structure of inclined impact aerating film of the present invention is:
(1) adopted the impact opening structure that tilts, formed bigger impact distance in blade interior by the impact opening that tilts, and produce large-area high cooled region at the blade internal surface.
(2) exhibition of air film hole and impact opening to, flow to be and be staggered, avoided the mutually mutual interference of impact opening with air film hole, help improving the cooling effect of blade.
(3) center line of impact opening tilts to the air film hole direction, helps reducing flow resistance, has increased the flowing velocity of air-flow on internal surface simultaneously, has improved the heat exchange effect.
(4) the air film hole diameter is bigger, can effectively reduce flow resistance, has increased the air film coverage area of outer surface simultaneously, and is also very favourable to improving heat exchange.
Description of drawings
Fig. 1 is a blade back structural drawing of the present invention.
Fig. 2 is a leaf basin structural drawing of the present invention.
Fig. 3 is the distribution schematic diagram of impact opening and air film hole.
Fig. 4 is the impact opening partial enlarged drawing.
Among the figure: 1. air film hole 2. impact openings
Embodiment
The present invention is described in further detail below in conjunction with accompanying drawing.
As shown in Figure 1, a kind of combined cooling structure that is applied to the inclined impact aerating film at aero engine turbine blades middle part, it comprises: air film hole and impact opening, it is characterized in that: the outer surface cloth at blade is equipped with the air film hole that diameter is 1.0mm～1.5mm, the exhibition of every exhaust fenestra to number be 10～15, downstream area at the blade interior air film hole is furnished with the inclined impact hole identical with air film hole quantity, the impact opening angle is 30 degree～45 degree, the direction of the center line deflection air film hole of impact opening, impact opening is arranged in the downstream of corresponding air film hole, with the distance of air film hole be 8～10 times of impact opening diameter, and impact opening and air film hole are staggered, the span distance of impact opening and air film hole is 7～9 times of impact opening diameter, the diameter of air film hole is 2～2.5 times of impact opening diameter, and the exhibition of air film hole is 4～5 times of the air film hole diameter to spacing.
Fig. 1 is that cooling structure of the present invention is applied to the blade integral structural drawing on the blade back.Fig. 2 is that cooling structure of the present invention is applied to the blade integral structural drawing on the leaf basin.1 is air film hole among the figure, 2 is impact opening, the air film hole of leaf basin and blade back is identical with the arrangement mode of impact opening, cooled gas enters into blade from the blade center cavity, at this moment the pressure at blade center cavity place can be higher than the pressure of blade outer surface, under the driving of certain pressure difference, cooled gas will flow to the area of low pressure from the high-pressure area.In the designed cooling structure of the present invention, cooled gas can penetrate by impact opening under the driving of pressure reduction, and arrives the blade internal surface with certain velocity shock, thereby form large-area impact cooled region at the blade internal surface, the cooling effect of this cooled region is very high.Cooled gas flows along the blade internal surface then, then discharges from air film hole, and forms the air film covering at blade outer surface, is used for the combustion gas and the blade of heat are kept apart, and allows blade form the protective film of one deck cold air.
Fig. 3 is the planimetric map of the designed a kind of cooling structure of parameter area according to the present invention, from figure, can clearly find out the position relation of impact opening and air film hole, impact opening is arranged in the scope of 7～10 times of impact opening diameters in downstream of corresponding air film hole, and between two air film holes, impact opening and air film hole are staggered, the diameter of air film hole is 2～2.5 times of impact opening diameter, and the exhibition of air film hole is 4～5 times of air film hole diameter to spacing.Fig. 4 is the enlarged view of impact opening of the present invention, and the impact opening angle is 30 degree～45 degree, and the direction of center line deflection air film hole.
In the present invention, impact opening diameter 0.6～0.8mm, air film hole diameter 1.0～1.5mm.Adding man-hour can be with air film hole and blade outer surface while casting, impact opening adopts the technology of laser boring, and processing of leaves technology is simple like this, can reduce processing cost greatly, and directly the blade of casting has good intensity, can adapt to higher thermal stress and centrifugal stress.
The present invention has not only improved whole heat exchange effect from the thermal conduction study angle, and the overall thermal stress distribution is even, and flow resistance is also well below common turbine blade, and whole channel inner pressure loss is well below the turbine cooling blade of routine.
Use the turbine blade of above-mentioned novel cooling structure design, through simplified model experiment and its heat-exchange performance of three-dimensional numerical value simulation test and flow resistance, the average cooling effect of integral blade can reach more than 0.7, and flow resistance is starkly lower than conventional turbine cooling blade, its pitot loss is significantly less than common interior cold blade, and the blending loss that while air film jet brings also is less than conventional turbine blade.
Claims (1)
1. combined cooling structure that is used for the inclined impact aerating film at aero engine turbine blades middle part, it comprises: air film hole and impact opening, it is characterized in that: the outer surface cloth of blade is equipped with the air film hole that diameter is 1.0mm～1.5mm (1), the exhibition of every exhaust fenestra to number be 10～15, downstream area at the blade interior air film hole is furnished with the inclined impact hole (2) identical with air film hole quantity, the impact opening angle is 30 degree～45 degree, the direction of the center line deflection air film hole of impact opening, impact opening is arranged in the downstream of corresponding air film hole, with the distance of air film hole be 8～10 times of impact opening diameter, and impact opening and air film hole are staggered, the span distance of impact opening and air film hole is 7～9 times of impact opening diameter, the diameter of air film hole is 2～2.5 times of impact opening diameter, and the exhibition of air film hole is 4～5 times of the air film hole diameter to spacing.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2007101187655A CN101126324A (en) | 2007-07-13 | 2007-07-13 | Combined cooling structure for turbine blade middle-part inclined impact aerating film |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2007101187655A CN101126324A (en) | 2007-07-13 | 2007-07-13 | Combined cooling structure for turbine blade middle-part inclined impact aerating film |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN101126324A true CN101126324A (en) | 2008-02-20 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CNA2007101187655A Pending CN101126324A (en) | 2007-07-13 | 2007-07-13 | Combined cooling structure for turbine blade middle-part inclined impact aerating film |
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Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102022139A (en) * | 2010-12-10 | 2011-04-20 | 南京航空航天大学 | Internal cooling device and method thereof for ground gas turbine blade |
| CN103277145A (en) * | 2013-06-09 | 2013-09-04 | 哈尔滨工业大学 | Cooling blade of gas turbine |
| CN103806951A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Turbine blade combining cooling seam gas films with turbulence columns |
| CN105114128A (en) * | 2015-09-07 | 2015-12-02 | 沈阳航空航天大学 | Turbine blade structure under air cooling and heat pipe heat transfer combined action |
| CN113374536A (en) * | 2021-06-09 | 2021-09-10 | 中国航发湖南动力机械研究所 | Gas turbine guide vane |
| CN114282323A (en) * | 2021-12-27 | 2022-04-05 | 北京航空航天大学 | Flow distribution prediction method for turbine blade laminate cooling structure |
-
2007
- 2007-07-13 CN CNA2007101187655A patent/CN101126324A/en active Pending
Cited By (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102022139A (en) * | 2010-12-10 | 2011-04-20 | 南京航空航天大学 | Internal cooling device and method thereof for ground gas turbine blade |
| CN103277145A (en) * | 2013-06-09 | 2013-09-04 | 哈尔滨工业大学 | Cooling blade of gas turbine |
| CN103806951A (en) * | 2014-01-20 | 2014-05-21 | 北京航空航天大学 | Turbine blade combining cooling seam gas films with turbulence columns |
| CN105114128A (en) * | 2015-09-07 | 2015-12-02 | 沈阳航空航天大学 | Turbine blade structure under air cooling and heat pipe heat transfer combined action |
| CN113374536A (en) * | 2021-06-09 | 2021-09-10 | 中国航发湖南动力机械研究所 | Gas turbine guide vane |
| CN114282323A (en) * | 2021-12-27 | 2022-04-05 | 北京航空航天大学 | Flow distribution prediction method for turbine blade laminate cooling structure |
| CN114282323B (en) * | 2021-12-27 | 2024-05-14 | 北京航空航天大学 | Flow distribution prediction method for turbine blade laminate cooling structure |
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Open date: 20080220 |