CN203925753U - Secondary flow control structure for gas turbine - Google Patents
Secondary flow control structure for gas turbine Download PDFInfo
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- CN203925753U CN203925753U CN201420226832.0U CN201420226832U CN203925753U CN 203925753 U CN203925753 U CN 203925753U CN 201420226832 U CN201420226832 U CN 201420226832U CN 203925753 U CN203925753 U CN 203925753U
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- gas turbine
- secondary flow
- pore
- flow control
- control structure
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Abstract
The utility model relates to the secondary flow control structure for gas turbine.The rotor that described gas turbine comprises casing, is fixed to the stator of described casing and is arranged on axial the place ahead of described stator.The described secondary flow control structure for gas turbine comprises a plurality of introducing pores, a plurality of pore and a plurality of bleed air line of drawing, described a plurality of introducing pore is arranged at position corresponding with the axial length first half section of described rotor on described casing, described a plurality of pore of drawing is arranged at position corresponding with the axial rearward direction of described stator on described casing, and described a plurality of bleed air line are for being communicated to described a plurality of introducing pore by described a plurality of pores of drawing respectively.Secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: can be when possessing simple structure and less parts, and realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Description
Technical field
The utility model relates to the secondary flow control structure for gas turbine.
Background technique
Rotor blade and the gap between casing in aero-gas turbine are very big to the performance impact of motor.For example, Britain's Rolls-Royce (RR) company shows the research of modern gas turbine engines: the every increase length of blade 1% of tip clearance (that is, the gap between rotor blade and casing), and efficiency approximately reduces by 1.5%, and oil consumption rate approximately increases by 2%.Oil consumption rate not only affects engine performance, also affects its life cycle cost simultaneously.According to company of General Electric (U.S.A.) (GE) on CF5-50 engine analysis: tip clearance accounts for 67% of blade profile and clearance seal total losses to the impact of oil consumption rate.
In order to control tip clearance, existing active gap control means are mainly divided into thermal type gap control, mechanical type gap control, pressure type gap control and new gap control.Thermal type gap control is mainly the air cooling turbine casing of drawing from fan outlet or gas compressor grade, by controlling the parameter adjustment turbine casing temperature distribution such as amount of air entrainment or bleed temperature, to reach the radial thermal expansion displacement of controlling turbine casing turbine outer ring assembly, reduce motor cruising condition tip clearance, shortcoming is that response is slow, and expend gas and can cause thrust loss large, the layout of air entraining pipe also can increase weight and the complexity of system in addition.Mechanical type gap control jointly realizes tip clearance by connection set and driving mechanism (water pressing, electromechanical formula, electromagnetic type etc.) and changes, and response is fast, and control accuracy is higher.Pressure type gap control is mainly utilized the pressure of engine interior or outside supply and to controlling the adjusting of valve, is caused load offset, causes turbine outer ring radially to produce displacement, realizes tip clearance and controls.The feedback of new gap control general using gap sensor, controls to optimum value by tip clearance fast, as utilizes shape memory metal etc.Yet all there is the defects such as complex structure, parts be various in above-mentioned existing active gap control means.
Therefore, hope can have a kind of secondary flow control structure for gas turbine, and it can be when possessing simple structure and less parts, realizes the effective control to tip clearance, and then effectively reduces or hinder blade tip secondary flow.
Model utility content
An object of the present utility model is, above-mentioned defect for gas turbine in prior art, a kind of secondary flow control structure for gas turbine is provided, it can be when possessing simple structure and less parts, the effective control of realization to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Above object of the present utility model realizes by a kind of secondary flow control structure for gas turbine, described gas turbine comprises casing, be fixed to the stator of described casing, and the rotor that is arranged on axial the place ahead of described stator, the described secondary flow control structure for gas turbine comprises a plurality of introducing pores, a plurality of pore and a plurality of bleed air line of drawing, described a plurality of introducing pore is arranged at position corresponding with the axial length first half section of described rotor on described casing, described a plurality of pore of drawing is arranged at position corresponding with the axial rearward direction of described stator on described casing, described a plurality of bleed air line is for being communicated to described a plurality of introducing pore by described a plurality of pores of drawing respectively.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: can be when possessing simple structure and less parts, the effective control of realization to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, described a plurality of introducing pore be arranged on described casing with the axial length of described rotor first three/mono-section of corresponding position.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: balancing rotor blade two sides pressure reduction better, thereby realize better the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, the bleed direction of described a plurality of introducing pores becomes 0-20 degree with respect to the diametric(al) of described casing.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: can make bleed direction and rotor direction liquidate, thereby realize better the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, described a plurality of bleed air line is positioned at outside described casing.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by suitable bleed air line set-up mode, realize suitable bleed, thereby realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, within described a plurality of bleed air line is positioned at described casing.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by suitable bleed air line set-up mode, realize suitable bleed, thereby realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow; In addition, be positioned at other parts that bleed air line within casing can not disturbed aero-gas turbine.
Preferably, described a plurality of introducing pore, described a plurality of quantity of drawing pore and described a plurality of bleed air line equate.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by suitable introducing pore, draw pore and bleed air line quantity arranges, realize suitable bleed, thereby realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, described a plurality of introducing pores are divided into N group and introduce pore, wherein, the rotor blade quantity that N is described rotor, every group introduce pore be arranged on described casing with two adjacent rotor blades between corresponding position, position.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by introducing pore, be arranged on suitable corresponding position, the position with between two adjacent rotor blades, thereby realize better the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, every group of introducing pore comprises that three are introduced pores.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by suitable introducing pore quantity, arrange, realize suitable bleed, thereby realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, described aero-gas turbine comprises multistage, and every grade comprises stator and rotor, and described a plurality of pores of drawing are arranged at position corresponding with the axial rearward direction of stator at the same level on described casing.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by the suitable pore position of drawing, arrange, realize suitable bleed, thereby realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Preferably, described aero-gas turbine comprises multistage, and every grade comprises stator and rotor, and described a plurality of pores of drawing are arranged at position corresponding with the axial rearward direction of rear class stator on described casing.
According to technique scheme, secondary flow control structure for gas turbine of the present utility model can play following useful technique effect: by the suitable pore position of drawing, arrange, realize suitable bleed, thereby realize the effective control to tip clearance, and then effectively reduce or hinder blade tip secondary flow.
Accompanying drawing explanation
Fig. 1 is rotor and the casing cross-sectional view of aero-gas turbine gas compressor.
Fig. 2 is the casing cross-sectional view of the aero-gas turbine gas compressor of the utility model one preferred embodiment.
Fig. 3 is that the introducing pore of the secondary flow control structure for gas turbine of the utility model one preferred embodiment is with respect to the schematic diagram of rotor axial position.
Fig. 4 is the schematic diagram of the first bleed mode of the secondary flow control structure for gas turbine of the utility model one preferred embodiment.
Fig. 5 is the schematic diagram of the second bleed mode of the secondary flow control structure for gas turbine of the utility model one preferred embodiment.
Reference numerals list:
1, rotor;
2, stator;
3, casing;
4, draw pore;
5, introduce pore;
6, bleed air line;
11, rotor blade;
111, pressure side;
112, suction surface;
S, secondary flow;
V, swirl flow;
L, bleed direction;
A, axial.
Embodiment
Below in conjunction with specific embodiments and the drawings, the utility model is described in further detail; set forth in the following description more details so that fully understand the utility model; but the utility model obviously can be implemented with the multiple alternate manner that is different from this description; those skilled in the art can do similar popularization, deduction according to practical situations without prejudice to the utility model intension in the situation that, therefore should be with the content constraints of this specific embodiment protection domain of the present utility model.
Fig. 1 is rotor and the casing cross-sectional view of aero-gas turbine gas compressor.
Rotor 1 is rotated counterclockwise, and locates the secondary flow S from the pressure side 111 of rotor blade 11 to the suction surface 112 of rotor blade 11 in the gap of rotor blade 11 and casing 3 (that is, tip clearance).The object of controlling tip clearance is in order to reduce or to hinder this secondary flow S, and then weakens the swirl flow V between rotor blade 11, reduces loss.
Fig. 2 is the casing cross-sectional view of the aero-gas turbine gas compressor of the utility model one preferred embodiment.Fig. 3 is that the introducing pore of the secondary flow control structure for gas turbine of the utility model one preferred embodiment is with respect to the schematic diagram of rotor axial position.Fig. 4 is the schematic diagram of the first bleed mode of the secondary flow control structure for gas turbine of the utility model one preferred embodiment.Fig. 5 is the schematic diagram of the second bleed mode of the secondary flow control structure for gas turbine of the utility model one preferred embodiment.
As Figure 1-5, the aero-gas turbine gas compressor of the utility model one preferred embodiment comprises casing 3, be fixed to the stator 2 of casing, and the rotor 1 that is arranged on axial the place ahead of stator 2, wherein, secondary flow control structure for gas turbine comprises a plurality of introducing pores 5, a plurality of pore 4 and a plurality of bleed air line 6 of drawing, a plurality of introducing pores 5 be arranged on casing 3 with rotor 1 (specifically, rotor blade 11) the corresponding position of axial length first half section, a plurality of pores 4 of drawing are arranged at position corresponding with the axial rearward direction of stator 2 on casing 3, a plurality of bleed air line 6 are for being communicated to a plurality of introducing pores 5 by a plurality of pores 4 of drawing respectively.
Like this, air is guided to and introduces pore 5 from stator 2 axial rearward direction, for hindering secondary flow S, and then can reduce between leaf and leak, reduce leakage vortex between leaf, wipe the impact of staying vortex pair runner flow field, improve cascade flow field simultaneously, improve compressor efficiency.
It should be noted that in Fig. 3-5 and all indicated axial A, this axial A is consistent with the airintake direction of aero-gas turbine gas compressor.Therefore, " axially the place ahead " mentioned above also can be regarded as " airintake direction upstream ", and " axial rearward direction " mentioned above also can be regarded as " airintake direction downstream ".
Comparatively speaking, the axial length of rotor blade 11 first three/mono-section be the larger region of pressure reduction, rotor blade two sides.Therefore, preferably, as shown in Figure 3, a plurality of introducing pores 5 be arranged on casing 3 with the axial length of rotor 1 (specifically, rotor blade 11) first three/mono-section of corresponding position.
Preferably, as shown in Figure 2, the bleed direction L of a plurality of introducing pores 5 becomes 0-20 degree with respect to the diametric(al) of casing 3.More preferably, the bleed direction L of a plurality of introducing pores 5 becomes 10-20 degree with respect to the diametric(al) of casing 3.
As shown in Figure 2, bleed direction L and compressor rotor 1 sense of rotation liquidate.The diameter of introducing pore 5 can be 0.5-3mm.Concrete introducing hole diameter and angle can be determined by concrete gas compressor runnability.
In the first bleed mode, as shown in Figure 4, a plurality of bleed air line 6 are positioned at outside casing 3.For example, a plurality of bleed air line 6 are by being welded to connect to casing 3.
In the second bleed mode, as shown in Figure 5, within a plurality of bleed air line 6 are positioned at casing 3.That is to say, within casing 3, imbed a plurality of bleed air line 6.
Preferably, a plurality of introducing pores 5, a plurality of quantity of drawing pore 4 and a plurality of bleed air line 6 equate.Certainly, those skilled in the art are appreciated that on the basis of the utility model disclosure also can adopt a bleed main pipe rail, and this bleed main pipe rail is branched into a plurality of bleed bye-passes, for respectively a plurality of pores 4 of drawing being communicated to a plurality of introducing pores 5.Within above-mentioned modification falls into protection domain of the present utility model equally.
Preferably, as shown in Figure 3, a plurality of introducing pores 5 are divided into N group and introduce pore, and wherein, N is rotor blade 11 quantity of rotor 1, every group introduce pore 5 be arranged on casing 3 with two adjacent rotor blades 11 between corresponding position, position.
Preferably, as shown in Figure 3, introduce pore 5 for every group and comprise that three are introduced pore 5,5,5.Certainly, those skilled in the art are appreciated that on the basis of the utility model disclosure every group is introduced the introducing pore that pore also can adopt other quantity.Within above-mentioned modification falls into protection domain of the present utility model equally.
Preferably, as shown in Figure 3, introducing pore 5,5,5 for three every group can be uniformly distributed vertically.
Aero-gas turbine gas compressor can comprise multistage, and every grade comprises stator 2 and rotor 1.For brevity, Fig. 4 and Fig. 5 only show the wherein one-level of aero-gas turbine gas compressor, have omitted other level.A plurality of pores 5 of drawing can be arranged at position corresponding with the axial rearward direction of stator 2 at the same level on casing 3 (as shown in Figure 4 and Figure 5).A plurality of pores 5 of drawing also can be arranged on casing 3 the position (not shown) for example, with the axial rearward direction of rear class (, rear one-level or afterwards what) stator corresponding.
Although above embodiment is mainly for by secondary flow control structure of the present utility model, the gas compressor for gas turbine describes, but those skilled in the art are appreciated that on the basis of the utility model disclosure, secondary flow control structure of the present utility model equally can be for other parts in gas turbine, such as turbine etc.
Above embodiment of the present utility model is described, but it will be understood to those of skill in the art that, above-mentioned embodiment does not form restriction of the present utility model, those skilled in the art can carry out multiple modification on the basis of above disclosure, and do not exceed scope of the present utility model.
Claims (10)
1. the secondary flow control structure for gas turbine, described gas turbine comprises casing, be fixed to the stator of described casing, and the rotor that is arranged on axial the place ahead of described stator, it is characterized in that, the described secondary flow control structure for gas turbine comprises a plurality of introducing pores, a plurality of pore and a plurality of bleed air line of drawing, described a plurality of introducing pore is arranged at position corresponding with the axial length first half section of described rotor on described casing, described a plurality of pore of drawing is arranged at position corresponding with the axial rearward direction of described stator on described casing, described a plurality of bleed air line is for being communicated to described a plurality of introducing pore by described a plurality of pores of drawing respectively.
2. the secondary flow control structure for gas turbine as claimed in claim 1, is characterized in that, described a plurality of introducing pores be arranged on described casing with the axial length of described rotor first three/mono-section of corresponding position.
3. the secondary flow control structure for gas turbine as claimed in claim 1, is characterized in that, the bleed direction of described a plurality of introducing pores becomes 0-20 degree with respect to the diametric(al) of described casing.
4. the secondary flow control structure for gas turbine as claimed in claim 1, is characterized in that, described a plurality of bleed air line are positioned at outside described casing.
5. the secondary flow control structure for gas turbine as claimed in claim 1, is characterized in that, within described a plurality of bleed air line are positioned at described casing.
6. the secondary flow control structure for gas turbine as claimed in claim 1, is characterized in that, described a plurality of introducing pores, described a plurality of quantity of drawing pore and described a plurality of bleed air line equate.
7. the secondary flow control structure for gas turbine as claimed in claim 1, it is characterized in that, described a plurality of introducing pore is divided into N group and introduces pore, wherein, N is the rotor blade quantity of described rotor, every group introduce pore be arranged on described casing with two adjacent rotor blades between corresponding position, position.
8. the secondary flow control structure for gas turbine as claimed in claim 7, is characterized in that, introduces pore for every group and comprises that three are introduced pore.
9. the secondary flow control structure for gas turbine as claimed in claim 1, it is characterized in that, described aero-gas turbine comprises multistage, and every grade comprises stator and rotor, and described a plurality of pores of drawing are arranged at position corresponding with the axial rearward direction of stator at the same level on described casing.
10. the secondary flow control structure for gas turbine as claimed in claim 1, it is characterized in that, described aero-gas turbine comprises multistage, and every grade comprises stator and rotor, and described a plurality of pores of drawing are arranged at position corresponding with the axial rearward direction of rear class stator on described casing.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201420226832.0U CN203925753U (en) | 2014-05-05 | 2014-05-05 | Secondary flow control structure for gas turbine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201420226832.0U CN203925753U (en) | 2014-05-05 | 2014-05-05 | Secondary flow control structure for gas turbine |
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| Publication Number | Publication Date |
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| CN203925753U true CN203925753U (en) | 2014-11-05 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN201420226832.0U Expired - Lifetime CN203925753U (en) | 2014-05-05 | 2014-05-05 | Secondary flow control structure for gas turbine |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104675755A (en) * | 2015-01-14 | 2015-06-03 | 西北工业大学 | Circumferential staggered self-circulating casing treating method for axial-flow compressor |
-
2014
- 2014-05-05 CN CN201420226832.0U patent/CN203925753U/en not_active Expired - Lifetime
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104675755A (en) * | 2015-01-14 | 2015-06-03 | 西北工业大学 | Circumferential staggered self-circulating casing treating method for axial-flow compressor |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| CP03 | Change of name, title or address | ||
| CP03 | Change of name, title or address |
Address after: 200241 Minhang District Lianhua Road, Shanghai, No. 3998 Patentee after: AECC COMMERCIAL AIRCRAFT ENGINE Co.,Ltd. Address before: 201108 Shanghai city Minhang District Lotus Road No. 3998 Patentee before: AVIC Commercial Aircraft Engine Co.,Ltd. |
|
| CX01 | Expiry of patent term | ||
| CX01 | Expiry of patent term |
Granted publication date: 20141105 |