CN101280692A - Turbine blade middle part microchannel inner-cooling structure - Google Patents
Turbine blade middle part microchannel inner-cooling structure Download PDFInfo
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- CN101280692A CN101280692A CNA2008101142023A CN200810114202A CN101280692A CN 101280692 A CN101280692 A CN 101280692A CN A2008101142023 A CNA2008101142023 A CN A2008101142023A CN 200810114202 A CN200810114202 A CN 200810114202A CN 101280692 A CN101280692 A CN 101280692A
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Abstract
The invention relates to a cooling structure applied in micro-channels at the middle part of the turbine blade of an air engine, which consists essentially of micro-channels and partition walls and is characterized in that: the blade body of the turbine blade is divided into a plurality of partition walls; the micro-channels are densely distributed between every two partition walls; after flowing through the micro-channels, a cooling gas enters the central cavity of the blade and is discharged from gas film holes on the partition walls; the cooling requirement of the blade is met by the cooling of the micro-channels and the gas films of an outer surface. The interval of channels is 0.3-0.8mm; the thickness of walls is 0.2-0.5mm; the thickness of the partition walls is 1.0- 1.5mm; the diameter of the gas film holes is 0.8- 1.2mm. Judging from model tests and the results of three-dimensional numerical simulation, the cooling effect of the blade can reach more than 0.7; meanwhile, due to the characteristics of the cooling structure, processing is convenient, thereby greatly reducing processing cost; meanwhile the cooling structure is of extremely good strength.
Description
Technical field
The present invention relates to the inner cooling structure of a kind of partition rib and micro passage combined blade, 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.
Existing cooling technology commonly used has the following disadvantages:
(1) cooling effectiveness does not always increase with the increase of turbine blade cooling air volume, has arrived certain flow, and cooling effectiveness will be difficult to that lifting is arranged again.
(2) along with the improving constantly of fuel gas temperature before the turbine, particularly fuel gas temperature reaches 1300 ℃ even higher before turbine, and existing cooling technology commonly used will be difficult to satisfy the effectively needs of cooling, cause the turbine blade thermal fatigue to be damaged easily.
Summary of the invention
The objective of the invention is to partition rib and micro passage combination, form a kind of should be in the micro passage cooling structure at aero engine turbine blades middle part, this cooling result is: at the blade thickness of turbine blade is that the partition of 1～1.5mm is divided into a plurality of zones, zone between per two partitions is densely covered with small passage, micro passage spacing l is 0.3～0.8mm, wall thickness h is 0.2～0.5mm, the ratio of wall thickness and channel pitch
In 0.5～1 scope, and if under the situation of the permission of processing technology,
Value the smaller the better, the height of micro passage is 1.5～2.5mm, all have air film hole on every row's partition, air film hole diameter 0.8～1.2mm on the partition, the number of every exhaust fenestra is 15～20, cut off with the micro passage and locate to continue up to the blade tip place by blade root, the material that is adopted is identical with the material of turbine blade.
The invention has the advantages that: adopt micro-channel to replace the cooling structure at original blade middle part, strengthen the cooling effect of blade greatly, reduced the blade surface temperature, improved the cooling effect of blade integral, improve the intensity of blade, increased the ability of blade opposing thermal stress and centrifugal stress.
Description of drawings
Fig. 1 leaf cross-section figure of the present invention
The small conduit of Fig. 2 is at the schematic representation of blade applications
The small conduit of Fig. 3 is used the calorifics schematic diagram
Among the figure: 1, cut off 4 in blade inlet edge 2, blade outer wall 3, the leaf, micro passage 5, blade middle part 6, blade center cavity 7, leading edge impact opening 8, vane tip air-flow path 9, blade middle part 10, middle part air-flow path 11, afterbody micro passage 12, blade trailing edge 13, afterbody split seam
Embodiment
The present invention is in order to be applied to turbine blade with techniques for microchannel cooling, to cut off rib combines with the micro passage, form a kind of micro passage cooling structure that is adapted to the aero engine turbine blades middle part, this cooling result is: at the blade thickness of turbine blade is that the partition of 1～1.5mm is divided into a plurality of zones, zone between per two partitions is densely covered with small passage, micro passage spacing l is 0.3～0.8mm, and wall thickness h is 0.2～0.5mm, the ratio of wall thickness and channel pitch
In 0.5～1 scope, and if under the situation of the permission of processing technology,
Value the smaller the better, the height of micro passage is 1.5～2.5mm, all have air film hole on every row's partition, air film hole diameter 0.8～1.2mm on the partition, the number of every exhaust fenestra is 15～20, cut off with the micro passage and locate to continue up to the blade tip place by blade root, the material that is adopted is identical with the material of turbine blade.
Cooled gas enters in the blade micro passage from root of blade, and at this moment the pressure in the blade micro passage can be higher than the pressure of blade outer surface, and 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 is under the driving of pressure reduction; can pass through the micro passage; and enter the center cavity 6 of up/down perforation via vane tip passage 8 turnover; and discharge with the air film hole of certain speed from the blade; form air film at blade outer surface and cover, be used for the combustion gas and the blade of heat are kept apart, allow blade form the protective film of one deck cold air.
Cooled gas enters in the micro passage, blade middle part from root of blade can produce strong heat exchange effect, owing to cut off the existence of rib 3, has strengthened the conduction effect of wall 2 with micro passage 4 simultaneously, and it is more even that the temperature field is distributed.
Among the present invention, the micro passage spacing is 0.3～0.8mm, and wall thickness is 0.2～0.5mm, and the height of micro passage is 1.5～2.5mm, adds the processing technology that can adopt the line cutting man-hour, and welded together with the blade external form.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 the intensity of blade also can be higher than common turbine blade, and whole channel inner pressure loss also is lower than conventional turbine cooling blade.
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, blade has good intensity, can resist higher thermal stress and centrifugal stress.
Claims (1)
1, a kind of micro passage cooling structure that is applied to the aero engine turbine blades middle part, it is mainly formed by cutting off with the micro passage, it is characterized in that: at the blade thickness of turbine blade is that the partition (3) of 1～1.5mm is divided into a plurality of zones, zone between per two partitions (3) is densely covered with small passage (4), micro passage spacing l is 0.3～0.8mm, wall thickness h is 0.2～0.5mm, the ratio of wall thickness and channel pitch
In 0.5～1 scope, the height of micro passage is 1.5～2.5mm, all have air film hole on every row's partition (3), cut off (3) and go up air film hole diameter 0.8～1.2mm, the number of every exhaust fenestra is 15～20, cut off (3) and continue up to the blade tip place with micro passage (4) by the blade root place, the material that is adopted is identical with the material of turbine blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2008101142023A CN101280692A (en) | 2008-06-02 | 2008-06-02 | Turbine blade middle part microchannel inner-cooling structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2008101142023A CN101280692A (en) | 2008-06-02 | 2008-06-02 | Turbine blade middle part microchannel inner-cooling structure |
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| Publication Number | Publication Date |
|---|---|
| CN101280692A true CN101280692A (en) | 2008-10-08 |
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| Application Number | Title | Priority Date | Filing Date |
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| CNA2008101142023A Pending CN101280692A (en) | 2008-06-02 | 2008-06-02 | Turbine blade middle part microchannel inner-cooling structure |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101832181A (en) * | 2010-03-25 | 2010-09-15 | 北京航空航天大学 | Novel film cooling hole with anti-whorl hole branch structure |
| CN102562176A (en) * | 2010-12-22 | 2012-07-11 | 通用电气公司 | Cooling channel systems for high-temperature components covered by coatings, and related processes |
| CN103831574A (en) * | 2013-12-10 | 2014-06-04 | 贵州黎阳航空动力有限公司 | Process for repairing hollow turbine blade body blocking cover |
-
2008
- 2008-06-02 CN CNA2008101142023A patent/CN101280692A/en active Pending
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101832181A (en) * | 2010-03-25 | 2010-09-15 | 北京航空航天大学 | Novel film cooling hole with anti-whorl hole branch structure |
| CN101832181B (en) * | 2010-03-25 | 2014-01-29 | 北京航空航天大学 | Novel film cooling hole with anti-whorl hole branch structure |
| CN102562176A (en) * | 2010-12-22 | 2012-07-11 | 通用电气公司 | Cooling channel systems for high-temperature components covered by coatings, and related processes |
| CN103831574A (en) * | 2013-12-10 | 2014-06-04 | 贵州黎阳航空动力有限公司 | Process for repairing hollow turbine blade body blocking cover |
| CN103831574B (en) * | 2013-12-10 | 2016-03-23 | 贵州黎阳航空动力有限公司 | Hollow turbine vane blade blanking cover renovation technique |
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Open date: 20081008 |