CN102828781B - Fuel gas turbine cooling blade - Google Patents
Fuel gas turbine cooling blade Download PDFInfo
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- CN102828781B CN102828781B CN201110163286.1A CN201110163286A CN102828781B CN 102828781 B CN102828781 B CN 102828781B CN 201110163286 A CN201110163286 A CN 201110163286A CN 102828781 B CN102828781 B CN 102828781B
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Abstract
The invention discloses a fuel gas turbine cooling blade, which comprises a blade main body, a cooling cavity and a plurality of pin fins, wherein the cooling cavity is formed by the wall part of the blade main body and is used for cooling fluid flowing therethrough, and the pin fins are fixed in the cooling cavity. The fuel gas turbine cooling blade is characterized in that the pin fins are provided with holes for the cooling fluid flowing therethrough. Through additionally arranging the pin fins with the holes in bodies inside the turbine blade, the flow area of the cooling fluid is increased and the flow resistance is reduced, moreover, through designing the holes as inclined holes to generate a jet flow, heat exchange in the interior of a blade is enhanced and the blade can remain a lower temperature level, so that the service life of the blade is prolonged.
Description
Technical field
The present invention relates to gas turbine engine, particularly relate to the technique for cooling blades field of gas turbine.
Background technique
Along with developing by leaps and bounds of aero engine technology, before aero-engine compressor pressure ratio and turbine, inlet temperature all significantly improves, and this must cause the heat load suffered by turbine blade to increase, and makes it bear very serious thermal stress.For addressing this problem, except development new material and new process, one of a crucial factor adopts advanced efficient strengthening cooling technology to turbine blade.The cooling technology of turbine blade mainly carries out from two aspects: one is the disturbance strengthening turbine blade internal cooling air, increases the heat exchange area of turbine blade inside; Two is adopt gaseous film control, effectively to intercept the convection heat exchange of high-temperature fuel gas to turbine blade at blade surface.In the past, in order at inner forced heat exchanging, most of turbine blade adopts and installs coarse fin additional in turbine blade inside or install the structure of column type turbulence columns additional.But this structure exists following deficiency: when (1) cooling-air flows through this structure, flow resistance loss is comparatively large, cooling-air under identical inlet pressure is made to be not enough to meet the demand to turbine blade cooling; (2) heat transfer effect is not good, because larger flow resistance loss makes cooling-air less, thus limits heat transfer effect.
Carried out large quantifier elimination to turbine blade inside forced heat exchanging technology both at home and abroad, its technological core and the measure taked are all round the heat exchange area increasing blade interior, and the disturbance of the cooling-air of strengthening blade interior is carried out.Such as someone proposes and adopts triangular fin and V-arrangement to be interrupted the structure of fin in blade interior.Also someone proposes and adopts cylindrical turbulence columns to strengthen the structural type of heat exchange in turbine blade inside.Somebody proposes the structural type adopting the heat exchange of oval turbulence columns strengthening blade interior.The structural type of these blade interior all enhances the disturbance of cooling-air in blade interior above, and traditional straight rib is compared, and heat exchange strengthens all to some extent, but also reflects that these structures have larger fluid resistance losses simultaneously.
Summary of the invention
For this reason, the present invention is intended to the turbulence columns by installing additional in turbine blade inside through remodeling, the circulation area that the perforate on shaft of this turbulence columns adds cooling fluid decreases flow resistance, and produce jet by the hole of flow-disturbing shaft is set to slope hole, the heat exchange of strengthening blade interior, make blade remain on lower temperature levels, improve the working life of blade.
At this, disclose a kind of gas combustion turbine cooling blade, comprise blade body, cooling chamber and multiple turbulence columns, wherein, cooling chamber is formed in order to supply cooling fluid from wherein flowing through by the wall portion of blade body, turbulence columns is fixed in cooling chamber, it is characterized in that: turbulence columns has the hole flowed through for cooling fluid.
Wherein, above described holes can be set to the parallel channels that its center line parallels with the straight line at the flow direction place of the cooling fluid before non-access aperture, the aperture of this parallel channels can be constant on the fore-and-aft direction of center line, also can be change such as form first big after small contraction passage or first little rear large expanding passage, design this parallel channels in any case, the flowing space all can be made to increase, and the flow velocity of fluid reduces, and flow resistance loss reduces.Moreover, turbulence columns is processed parallel channels also makes total heat exchange area increase thus enhance heat exchange, further, the cooling fluid of the parallel channels in turbulence columns also can form jet with the cylinder of other adjacent turbulence columns, thus can strengthen heat exchange.
Preferably, above described holes can also be set to the straight line ramp way at an angle at the flow direction place of the cooling fluid before its center line and non-access aperture.The size at the angle of inclination of ramp way should be determined according to the flow velocity of the height of the spacing of turbulence columns streamwise, turbulence columns and cooling fluid.When cooling fluid flows through the space being mounted with this kind of structure turbulence columns, part cooling fluid flows through the ramp way in turbulence columns, this part cooling fluid is flowed out by ramp way and forms impact jet flow at the wall of blade body, thus strengthens the heat exchange between cooling fluid and wall.Meanwhile, due to the increase of the flowing space, flow velocity when making cooling fluid flow through turbulence columns decreases, and the pressure loss reduces.
More preferably, the scope at angle of inclination is between 30 °-60 °.
More preferably, angle of inclination is substantially 45 °.When so arranging, enhanced heat transfer effect is best, and flow resistance loss is not maximum.
More preferably, the wall portion of blade body has on the pressure side wall portion and suction side wall portion, hole comprises the first through hole and the second through hole, first through hole is used for cooling fluid and flows through and forms jet with the surface of on the pressure side wall portion thereafter, and the second through hole flows through for cooling fluid and forms jet with the surface of suction side wall portion thereafter.After arranging like this, cooling fluid can respectively on the pressure side wall and suction side wall formation impact jet flow after passing through the first through hole and the second through hole.Therefore, the heat exchange between the on the pressure side wall portion of cooling fluid and blade and the surface of suction side wall portion is strengthened.Meanwhile, due to the increase of the flowing space, flow velocity when making cooling fluid flow through turbulence columns decreases, and the pressure loss reduces.
Selectively, turbulence columns can be put for in-line arrangement or fork arrangement, and preferably, turbulence columns is put for pitching arrangement.Like this, when cooled gas flows through cooling chamber, cooled gas and then will form turbulent flow before turbulence columns thus the disturbance of strengthening cooling-air increase the heat exchange area of blade interior.
Parallel channels in above-mentioned turbulence columns or ramp way can be set to multiple along the flow direction of fluid according to actual needs in turbulence columns.
In a word, adopt in turbine blade inside the turbulence columns it offering parallel channels or ramp way, enhance heat exchange on the one hand, reduce flow losses on the other hand.Like this, effectively reduce temperature levels and the temperature gradient of turbine blade, turbine blade is significantly improved working life.
Accompanying drawing explanation
In order to explain the present invention, its illustrative embodiments will be described with reference to the drawings hereinafter,
In accompanying drawing:
Fig. 1 is the schematic diagram of the arrangement of turbulence columns;
Fig. 2 is the sectional view along the line A-A in Fig. 1;
Fig. 3 is the internal structure schematic diagram of the turbine blade of the arrangement of the turbulence columns had in Fig. 1;
Fig. 4 is the sectional view along the line B-B in Fig. 3;
Fig. 5 is a kind of sectional view with the turbulence columns of two slope holes;
Similar features in different figure is indicated by similar reference character.
Embodiment
Fig. 3 shows the internal structure schematic diagram of the turbine blade of the arrangement of the turbulence columns had in Fig. 1, and Fig. 4 shows the sectional view along the line B-B in Fig. 3.As can be seen from Fig. 3 and Fig. 4, blade 10 is divided into left and right two-part by fragmenting plate 4 by blade 10 inside, i.e. left cooling chamber and right cooling chamber.Turbulence columns 1 with the arrangement of Fig. 1 in the cooling chamber of left and right two.That is, fork row arrangement.Particularly, as shown in Figure 2, this turbulence columns 1 also has center hole 2, and this center hole is inclined hole, and the center line of this inclined hole 2 and blade inwall angularly α.More specifically, as shown in Figure 4, the upper end of turbulence columns 1 and the blade inwall of suction side fix, and the lower end of turbulence columns 1 is connected with blade inwall on the pressure side, and several turbulence columns 1 is arranged in the cooling chamber of left and right from the leading edge of blade respectively to trailing edge.Those skilled in the art should be appreciated that the fork row arrangement of turbulence columns is one preferably mode, at this, and also can by turbulence columns with in-line arrangement arrangement.
As shown in Figure 5, those skilled in the art should be appreciated that, corresponding to the blade inwall on the pressure side of blade 10 and the blade inwall of suction side, turbulence columns 1 can have the first inclined hole 21 and the second inclined hole 22, the angle of inclination of the first inclined hole 21 is α, flow through for cooling fluid and form jet with the surface of on the pressure side wall portion thereafter, the angle of inclination of the second through hole 22 is θ, flows through form jet with the surface of suction side wall portion thereafter for described cooling fluid.In addition, those skilled in the art also should be appreciated that, in a kind of scheme, above-mentioned inclined hole also can be substituted by parallel hole, namely the center line in hole parallels with the straight line at the flow direction place of the cooling fluid before non-access aperture, at this moment, parallel channels be arranged so that the flowing space increases, the flow velocity of fluid reduces, and flow resistance loss reduces.Moreover, turbulence columns is processed parallel channels and makes total heat exchange area increase thus enhance heat exchange, and the fluid flowing through parallel channels can form jet with the cylinder of other adjacent turbulence columns thus also strengthen heat exchange.Also should be appreciated that the inclined hole that can arrange some in each turbulence columns according to actual needs, in other words, inclined hole number is unrestricted.Moreover the inclined hole number in each turbulence columns can be not identical.
The row of turbulence columns 1 is determined according to blade 10 size, and namely blade 10 is larger, the row of turbulence columns 1 the more, otherwise then row is fewer.Tilt angle alpha heat exchanging and the flow losses of the inclined hole 2 of turbulence columns 1 have larger impact, and along with the reduction of angle of inclination a, the flow losses of fluid also reduce gradually, and meanwhile, the increase of heat exchange area also makes the heat exchange amount of fluid in circular hole improve.In addition, the reduction of the angle of inclination a of inclined hole 2 makes the center line of inclined hole 2 and the intersection point of wall closer to next row turbulence columns, and in other words, the downward row's turbulence columns in region, stationary point of impact jet flow moves.If those skilled in the art should understand impact stationary point, region distance next row turbulence columns is too small, then fully can not obtain the invigoration effect of impact jet flow heat exchanging, therefore, the angle of inclination a of inclined hole 2 will determine according to the flowing to spacing of turbulence columns 1.Preferably, this tilt angle alpha is between 30 °-60 °, and more preferably, when tilt angle alpha is 45 °, enhanced heat transfer effect is best, and flow resistance loss is not now maximum.Therefore from the angle optimized, be preferably scheme when the angle of inclination of inclined hole 2 is substantially 45 °.
When blade 10 works, cooling fluid and cooling-air enter into left cooling chamber and right cooling chamber from the inlet hole 7 of blade 10 first end, due in left cooling chamber and right cooling chamber respectively the row of pitching have a lot of turbulence columns 1, therefore a part of cooled gas forms turbulent flow after flowing through turbulence columns 1, part cooled gas forms jet to blade wall after flowing through the inclined hole 2 in turbulence columns 1, finally flow out from the exhaust port 8 of blade 10 second end after the cooled gas heat exchange of left cooling chamber, after the cooled gas heat exchange of right cooling chamber, major part is discharged from the exhaust seam 9 of blade trailing edge, sub-fraction flows out from the exhaust port 8 of the second end of blade.Like this, compared to the blade with turbulence columns of routine, on the one hand, the jet acting on blade wall can make the enhanced heat transfer effect of wall strengthen, on the other hand, the flowing space in the cooling chamber of left and right is increased inclined hole in turbulence columns thus the flow velocity making to flow through the cooled gas of turbulence columns decreases, and the pressure loss reduces simultaneously.
Technology contents of the present invention and technical characterstic have disclosed as above; but be appreciated that; under creative ideas of the present invention; those skilled in the art can make various changes said structure and improve; comprise the combination of disclosure or claimed technical characteristics separately here, comprise other combination of these features significantly.These distortion and/or combination all fall in technical field involved in the present invention, and fall into the protection domain of the claims in the present invention.It should be noted that by convention, in claim, use discrete component to be intended to comprise one or more such element.In addition, any reference mark in claims should be configured to limit the scope of the invention.
Claims (8)
1. a gas combustion turbine cooling blade, comprise blade body, cooling chamber and multiple turbulence columns, wherein, described cooling chamber is formed in order to supply cooling fluid from wherein flowing through by the wall portion of described blade body, described turbulence columns is fixed in described cooling chamber, it is characterized in that: described turbulence columns has the hole flowed through for described cooling fluid, the wall portion of described blade body has on the pressure side wall portion and suction side wall portion, described hole comprises the first through hole and the second through hole, described first through hole flows through for described cooling fluid and forms jet with the surface of described on the pressure side wall portion thereafter, described second through hole flows through for described cooling fluid and forms jet with the surface of described suction side wall portion thereafter.
2. gas combustion turbine cooling blade according to claim 1, is characterized in that, the center line in described hole with do not enter described hole before the straight line at flow direction place of described cooling fluid parallel.
3. gas combustion turbine cooling blade according to claim 1, is characterized in that, the straight line at the flow direction place of the center line in described hole and the described cooling fluid before not entering described hole has an angle of inclination.
4. gas combustion turbine cooling blade according to claim 3, is characterized in that, the scope at described angle of inclination is between 30-60 °.
5. gas combustion turbine cooling blade according to claim 3, is characterized in that, described angle of inclination is 45 °.
6. the gas combustion turbine cooling blade according to any one of claim 1-5, is characterized in that, described turbulence columns is that in-line arrangement or fork arrangement are put.
7. the gas combustion turbine cooling blade according to any one of claim 1-5, is characterized in that, described cooled blade also comprises fragmenting plate, and it is for being divided into two-part by described cooling chamber.
8. gas combustion turbine cooling blade according to claim 6, is characterized in that, described cooled blade also comprises fragmenting plate, and it is for being divided into two-part by described cooling chamber.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201110163286.1A CN102828781B (en) | 2011-06-16 | 2011-06-16 | Fuel gas turbine cooling blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201110163286.1A CN102828781B (en) | 2011-06-16 | 2011-06-16 | Fuel gas turbine cooling blade |
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| Publication Number | Publication Date |
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| CN102828781A CN102828781A (en) | 2012-12-19 |
| CN102828781B true CN102828781B (en) | 2015-06-10 |
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| CN201110163286.1A Active CN102828781B (en) | 2011-06-16 | 2011-06-16 | Fuel gas turbine cooling blade |
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| CN106089316A (en) * | 2016-06-02 | 2016-11-09 | 西北工业大学 | A kind of stator part projective table type periodic pressure wave generation device |
| CN105927288A (en) * | 2016-06-02 | 2016-09-07 | 西北工业大学 | Rotor disc boss type periodic pressure wave generating device |
| CN107191230B (en) * | 2017-07-04 | 2019-05-14 | 西安理工大学 | A kind of blade cooling microchannel structure |
| CN108386234B (en) * | 2018-02-23 | 2021-03-16 | 西安交通大学 | Internal cooling structure of combustion engine blade with column row fins as basic cooling unit |
| CN109931114A (en) * | 2019-03-15 | 2019-06-25 | 南京航空航天大学 | A kind of novel impinging cooling turbulence structure |
| CN109944645A (en) * | 2019-03-25 | 2019-06-28 | 南京航空航天大学 | A kind of flow-disturbing rod structure for turbo blade enhanced heat exchange |
| KR102180395B1 (en) * | 2019-06-10 | 2020-11-18 | 두산중공업 주식회사 | Airfoil and gas turbine comprising it |
| CN113090335A (en) * | 2021-05-14 | 2021-07-09 | 中国航发湖南动力机械研究所 | Impact air-entraining film double-wall cooling structure for turbine rotor blade |
| CN114215609B (en) * | 2021-12-30 | 2023-07-04 | 华中科技大学 | Blade internal cooling channel capable of enhancing cooling and application thereof |
Family Cites Families (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS62271902A (en) * | 1986-01-20 | 1987-11-26 | Hitachi Ltd | Cooled blade for gas turbine |
| CN1318735C (en) * | 2005-12-26 | 2007-05-30 | 北京航空航天大学 | Pulsing impact cooling blade for gas turbine engine |
| US7645122B1 (en) * | 2006-12-01 | 2010-01-12 | Florida Turbine Technologies, Inc. | Turbine rotor blade with a nested parallel serpentine flow cooling circuit |
| US7901183B1 (en) * | 2008-01-22 | 2011-03-08 | Florida Turbine Technologies, Inc. | Turbine blade with dual aft flowing triple pass serpentines |
| US8517684B2 (en) * | 2008-03-14 | 2013-08-27 | Florida Turbine Technologies, Inc. | Turbine blade with multiple impingement cooled passages |
| CN201218110Y (en) * | 2008-07-03 | 2009-04-08 | 西北工业大学 | Gas turbine cooling blade |
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Address after: 200241 Minhang District Lianhua Road, Shanghai, No. 3998 Patentee after: China Hangfa commercial aviation engine limited liability company Address before: 201109 Shanghai city Minhang District Hongmei Road No. 5696 Room 101 Patentee before: AVIC Commercial Aircraft Engine Co.,Ltd. |