CN104533541A - Gas turbine engine integral turbine guide device with heat deformation compensating structure - Google Patents
Gas turbine engine integral turbine guide device with heat deformation compensating structure Download PDFInfo
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- CN104533541A CN104533541A CN201410666630.2A CN201410666630A CN104533541A CN 104533541 A CN104533541 A CN 104533541A CN 201410666630 A CN201410666630 A CN 201410666630A CN 104533541 A CN104533541 A CN 104533541A
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- inner ring
- circumferential cut
- guide device
- gas turbine
- turbine engine
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- 239000000463 material Substances 0.000 claims description 3
- 229910000851 Alloy steel Inorganic materials 0.000 claims description 2
- 239000010935 stainless steel Substances 0.000 claims description 2
- 229910001220 stainless steel Inorganic materials 0.000 claims description 2
- 238000010438 heat treatment Methods 0.000 abstract 2
- 230000008642 heat stress Effects 0.000 abstract 1
- 230000002035 prolonged effect Effects 0.000 abstract 1
- 101100008048 Caenorhabditis elegans cut-4 gene Proteins 0.000 description 17
- 239000007789 gas Substances 0.000 description 17
- 230000008646 thermal stress Effects 0.000 description 12
- 239000002737 fuel gas Substances 0.000 description 8
- 238000002485 combustion reaction Methods 0.000 description 7
- 238000000034 method Methods 0.000 description 5
- 238000010586 diagram Methods 0.000 description 4
- 239000000567 combustion gas Substances 0.000 description 3
- 230000001351 cycling effect Effects 0.000 description 3
- 230000008569 process Effects 0.000 description 3
- 230000000717 retained effect Effects 0.000 description 3
- 230000009471 action Effects 0.000 description 2
- 230000003292 diminished effect Effects 0.000 description 2
- 238000005516 engineering process Methods 0.000 description 2
- 238000012423 maintenance Methods 0.000 description 2
- 238000002679 ablation Methods 0.000 description 1
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000015556 catabolic process Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 238000005336 cracking Methods 0.000 description 1
- 238000006731 degradation reaction Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000010304 firing Methods 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 239000002184 metal Substances 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a gas turbine engine integral turbine guide device with a heat deformation compensating structure. The heat deformation compensating structure is characterized in that a plurality of circumferential cuts are cut in an inner ring and/or outer ring of the guide device, and the direction of the circumferential cuts is basically the same as the blade type chord length direction of a blade root section. The heat deformation compensation structure is creatively designed, and compared with a segmented turbine guide device, the guide device with the heat deformation compensation structure can effectively reduce circumferential leakage of the inner ring of high pressure gas, and performance of an engine is ensured, and meanwhile, the number of parts is reduced; compared with a conventional whole-ring type guide device, the guide device with the heat deformation compensation structure can effectively reduce heat stress generated by discordance of inner ring heat deformation and outer ring heat deformation, and therefore the service life of the guide device is prolonged, so that the reliability of the guide device is improved; and the technical problem that inner ring heating deformation and outer ring heating deformation of the turbine guide device are not matched, and the circumferential direction of a wing plate is sealed tightly is well solved.
Description
Technical field
The present invention relates to turbine technology field, more specifically, relate to a kind of guider with the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure or the high-pressure turbine of ground combustion machine.
Background technique
The high-pressure turbine guider of gas turbine engine and ground combustion machine is one of major part of composition air-flow path, is made up of the guide vane of some and inner and outer rings.Inner and outer rings forms an annular pass, and blade is circumferentially distributed between inner and outer rings.High-temperature high-pressure fuel gas from firing chamber flows through from the space between the inner and outer rings and blade of guider.According to the difference of motor advanced level, fuel gas temperature and pressure are also different, and it is even higher that its pressure can reach 10 ~ 20 pressure, and its temperature levels can reach 1200 DEG C ~ 1800 DEG C.Due to the limiting temperature that the temperature of high-temperature fuel gas can bear considerably beyond material, outside the blast tube of the therefore inner and outer rings of guider and blade interior usually need logical cooled gas to reduce the temperature of the work of metal.Guider, under the acting in conjunction of high-temperature fuel gas and cooled gas, can produce very large thermal stress because guider internal and external temperature gradient is large.Meanwhile, because high-temperature fuel gas thermo parameters method is uneven, the temperature at usual blade place is higher than inner and outer rings, can guider inner and outer rings also be therefore different from the thermal distortion degree of blade.Like this under the effect of high-temperature fuel gas and thermal stress, faults such as guider very easily cracks, ablation and cause guider to be scrapped.
Current gas turbine engine or the high-pressure turbine of ground combustion machine, in order to ensure structural life-time and the reliability of high pressure guider, reduce operation and maintenance cost, common way is that domain guider is divided into some sections, is mainly: the guider of (1) segmental structure can be out of shape thus reduce thermal stress by releasing heat; (2) segmented guide of damage can be changed during engine maintenance as required, and do not need to change whole guider.But after domain guider is divided into some sections, because fuel gas temperature level is high, between guider inner and outer rings, can gap be there is, thus cause the leakage of high-pressure gas to cause the lower degradation problem of engine performance.In addition, high-temperature fuel gas leaks and the temperature of the casing of guider outside and the supporting structure of guider inside also can be caused to rise, and causes the working life of structural member and reliability to reduce.
Have no at present and solve the temperature distortion of nozzle ring inner and outer rings and not mate and listrium circumference is obturaged the correlative study of problem and technology report.
Summary of the invention
The technical problem to be solved in the present invention is to overcome in the high-pressure turbine guider of existing gas-turbine unit turbine integrated type guider or ground combustion machine, the problem that outer shroud temperature distortion is not mated and listrium circumference is obturaged, thering is provided one to have both can reduce in guider, outer shroud thermal stress can ensure again the gas-turbine unit turbine integrated type guider of thermal distortion compensation structure that guider circumference is obturaged or the high-pressure turbine guider of ground combustion machine, after this guider has described thermal distortion compensation structure, the listrium thermal stress of guider can be compensated well, improve guide structure life and reliability.
Above-mentioned purpose of the present invention is achieved by the following technical programs:
There is provided a kind of and there is the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure or the high-pressure turbine guider of ground combustion machine, comprise the blade of some, outer shroud and inner ring, inner ring and outer shroud form an annular blast tube, blade is circumferentially distributed between inner ring and outer shroud, and the surface of contact of blade and inner ring and outer shroud is blade root; Described thermal distortion compensation structure has cut several circumferential cut in described inner ring and/or on outer shroud, and circumferential cut direction is basic consistent with blade profile chord length direction, described blade root cross section;
Further, described circumferential cut is not described inner ring or outer shroud are cut completely, but and retains certain thickness between described blast tube.
Further, the basic uniform thickness in direction is cut on the thickness edge that described circumferential cut and described blast tube retain.
Preferably, described circumferential cut is arranged in described inner ring.Further, described circumferential cut is evenly cut between adjacent two blades of the inner ring of guider, circumferentially uniform, and direction is basic consistent with blade profile chord length direction, blade root cross section, and position is roughly between adjacent leaf grating.
Further, the value of the thickness between described circumferential cut from described blast tube is different according to the different values of motor.This thickness can value be between 0.5 ~ 1mm.
Further, the present invention is provided with tinsel in described circumferential cut.The present invention retains certain thickness between circumferential cut and described blast tube, between described circumferential cut, install tinsel, and tinsel makes circumferential cut circumferential deformation be inhibited and radially free deformation, can prevent caliper portion from cracking.
Further preferably, the material of described tinsel is different according to operating temperature level, selects common stainless steel or high-temperature alloy steel.
Further preferably, described tinsel thickness is vertically suitable with described circumferential cut thickness vertically.In the present invention, the direction perpendicular to inner ring or outer shroud is axially.
The present invention cuts several circumferential cut in inner ring described in guider and/or on outer shroud, and circumferential cut direction is basic consistent with blade profile chord length direction, described blade root cross section; This structure, compared with the domain formula guider of routine, effectively can reduce the inharmonious and thermal stress that produces of inside and outside ring thermal distortion, thus improve the life and reliability of guider; Compared with existing segmented nozzle ring, effectively can reduce the inner circumference of high-pressure gas to leakage, be conducive to the performance ensureing motor, decrease number of spare parts simultaneously.Do not get through between circumferential cut of the present invention and blast tube, but and retain certain thickness between described blast tube, effectively can prevent gas leakage.
On this basis, the present invention is provided with tinsel between described circumferential cut, the present invention combines research summary discovery further through creatively analyzing, and because the temperature levels of guider inner ring (or outer shroud) is lower than blade blade, therefore the thermal distortion of blade is greater than inner ring and outer shroud.And inner ring (or outer shroud) has circumferentially cut circumferential cut, therefore along with the lifting of engine behavior, temperature has constantly increased, and circumferential cut has the trend diminished.If circumferential cut is not cut inner ring (or outer shroud) completely, the circumferential cut of such inner ring (or outer shroud) can produce distortion gradually, this distortion can be decomposed into radially with the distortion of circumferential both direction.Along with motor experience parking---running---work cycle of parking, cold circulation that guider also experiences from cold---heat---, the caliper portion that circumferential cut and inner ring blast tube place retain circumferentially with under the repeated action of radial alternating force will produce certain fatigue, and then crack, thus inner ring is caused to obturage inefficacy.Therefore the present invention is on the guide structure basis of improving, in circumferential cut, tinsel is installed further, like this in the cold cycling process of guider, the distortion circumferentially of circumferential cut position is limited by the constraint of tinsel, and simultaneously distortion radially continues to retain.Ensure so well to discharge a part of thermal distortion, reduce thermal stress, the caliper portion retained can be avoided again to engender fatigue crack, and the combustion gas realizing to longer-term inner ring is obturaged simultaneously.
For convenience of the assembling of tinsel, between tinsel and circumferential cut circumference wall, remain with gap.Further, the gap between described tinsel and described circumferential cut circumference wall is 0.1 ~ 0.2mm.
The present invention has following beneficial effect:
The present invention creatively devises thermal distortion compensation structure at the high-pressure turbine guider of gas turbine engine integrated type nozzle ring or ground combustion machine, there is the guider of thermal distortion compensation structure of the present invention compared with segmented nozzle ring, the present invention effectively can reduce the inner circumference of high-pressure gas to leakage, be conducive to the performance ensureing motor, decrease number of spare parts simultaneously; Compared with the domain formula guider of routine, the present invention effectively can reduce the inharmonious and thermal stress that produces of inside and outside ring thermal distortion, thus improves the life and reliability of guider.Technical solution of the present invention is simple, solves the technical problem that the temperature distortion of nozzle ring inner and outer rings is not mated and listrium circumference is obturaged well simultaneously.
Accompanying drawing explanation
The Figure of description forming a application's part is used to provide a further understanding of the present invention, and schematic description and description of the present invention, for explaining the present invention, does not form inappropriate limitation of the present invention.In the accompanying drawings:
Fig. 1 is preferred embodiment of the present invention part-structure schematic diagram.
Fig. 2 is the partial schematic diagram in Fig. 1 shown in A.
Fig. 3 is preferred embodiment of the present invention circumferential cut temperature distortion partial schematic diagram (not installing tinsel).
Fig. 4 is preferred embodiment of the present invention circumferential cut temperature distortion partial schematic diagram (being provided with tinsel).
Embodiment
Below in conjunction with accompanying drawing, embodiments of the invention are described in detail, but the multitude of different ways that the present invention can be defined by the claims and cover is implemented.
As depicted in figs. 1 and 2, the invention provides a kind of gas turbine engine integrated type nozzle ring with thermal distortion compensation structure, comprise the blade 1 of some, outer shroud 2, inner ring 3, circumferential cut 4 and tinsel 5, inner ring 3 and outer shroud 2 form an annular blast tube, and blade 1 is circumferentially distributed between inner ring 3 and outer shroud 2; Described inner ring 3 is evenly cut several circumferential cut 4, and described circumferential cut 4 is between adjacent described blade 1; Circumferential cut 4 direction is basic consistent with blade profile chord length direction, described blade root (not indicating in figure) cross section; Described circumferential cut 4 is not described inner ring 3 cut completely, but and retain certain thickness between described inner ring 3 blast tube, different according to motor of thickness between described circumferential cut 4 from described inner ring 3 blast tube, this thickness can be got between 0.5 ~ 1mm, and the thickness that described circumferential cut 4 and inner ring 3 blast tube retain is along cutting the basic uniform thickness in direction.
The present invention cuts several circumferential cut 4 in inner ring described in guider 3 and/or on outer shroud 2, and circumferential cut 4 direction is basic consistent with blade profile chord length direction, described blade root cross section; Described circumferential cut 4 is not described inner ring 3 or outer shroud 2 are cut completely, but and retain certain thickness between described blast tube, this structure is compared with the domain formula guider of routine, inside and outside ring thermal distortion effectively can be reduced inharmonious and the thermal stress of generation, thus improve the life and reliability of guider; Compared with existing segmented nozzle ring, effectively can reduce the inner circumference of high-pressure gas to leakage, be conducive to the performance ensureing motor, decrease number of spare parts simultaneously.Do not get through between circumferential cut of the present invention and blast tube, effectively can prevent gas leakage.
Shown in accompanying drawing 4, between described circumferential cut 4, described tinsel 5 is installed.Described tinsel 5 is arranged in circumferential cut 4, and tinsel 5 makes circumferential cut 4 circumferential deformation be inhibited and radially free deformation.The present invention through creatively analyze combine further research sum up find, due to guider inner ring 3(or outer shroud 2) temperature levels lower than blade blade, therefore the thermal distortion of blade is greater than inner ring and outer shroud.And inner ring (or outer shroud) has circumferentially cut circumferential cut, therefore along with the lifting of engine behavior, temperature has constantly increased, and circumferential cut has the trend diminished.If circumferential cut is not cut inner ring (or outer shroud) completely, the circumferential cut of such inner ring (or outer shroud) can produce distortion gradually, this distortion can be decomposed into radially with the distortion of circumferential both direction.Along with motor experience parking---running---work cycle of parking, cold circulation that guider also experiences from cold---heat---, the caliper portion that circumferential cut and inner ring blast tube place retain circumferentially with under the repeated action of radial alternating force will produce certain fatigue, and then crack, thus inner ring is caused to obturage inefficacy.Therefore the present invention is on the guide structure basis of improving, in circumferential cut, tinsel is installed further, like this in the cold cycling process of guider, the distortion circumferentially of circumferential cut position is limited by the constraint of tinsel, and simultaneously distortion radially continues to retain.Ensure so well to discharge a part of thermal distortion, reduce thermal stress, the caliper portion retained can be avoided again to engender fatigue crack, and the combustion gas realizing to longer-term inner ring is obturaged simultaneously.
Described tinsel 5 is suitable with the thickness of circumferential cut 4.Described tinsel 5 is 0.1 ~ 0.2mm with the circumferential wall gap of described circumferential cut 4.Like this in the cold cycling process of guider, the distortion circumferentially of circumferential cut 4 position is limited by the constraint of tinsel 5, and simultaneously distortion radially continues to retain, and in accompanying drawing 4, dotted line represents and added tinsel 5, the circumferential cut 4 after temperature distortion.Can either discharge a part of thermal distortion like this, reduce thermal stress, the caliper portion retained can be avoided again to occur fatigue crack, and the combustion gas realizing inner ring 3 is obturaged simultaneously.
The gas turbine engine integrated type nozzle ring application implementation on the engine with thermal distortion compensation structure (band tinsel and not with tinsel) after the present invention is improved, experimental result finds, there is not the technical problem that the temperature distortion of nozzle ring inner and outer rings is not mated and listrium circumference is obturaged, compared with segmented nozzle ring, the present invention effectively can reduce the inner circumference of high-pressure gas to leakage, be conducive to the performance ensureing motor, decrease number of spare parts simultaneously; Compared with the domain formula guider of routine, the present invention effectively can reduce the inharmonious and thermal stress that produces of inside and outside ring thermal distortion, thus improves the life and reliability of guider.Especially with the guider of tinsel, obtain excellent application and realize result.
The foregoing is only the preferred embodiments of the present invention, be not limited to the present invention, for a person skilled in the art, the present invention can have various modifications and variations.Within the spirit and principles in the present invention all, any amendment done, equivalent replacement, improvement etc., all should be included in of the present invention comprising within scope.
Claims (9)
1. one kind has the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure, comprise the blade of some, outer shroud and inner ring, inner ring and outer shroud form an annular blast tube, blade is circumferentially distributed between inner ring and outer shroud, and the surface of contact of blade and inner ring and outer shroud is blade root; It is characterized in that, in described inner ring and/or on outer shroud, cut several circumferential cut, circumferential cut direction is basic consistent with blade profile chord length direction, described blade root cross section.
2. there is the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 1, it is characterized in that, described circumferential cut is not described inner ring or outer shroud are cut completely, but and retains certain thickness between described blast tube.
3. have the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 2, it is characterized in that, the thickness that described circumferential cut and described blast tube retain is along cutting the basic uniform thickness in direction.
4. have the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 1, it is characterized in that, described circumferential cut is arranged in described inner ring; Described circumferential cut is between adjacent two blades of the inner ring of guider.
5. have the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 2, it is characterized in that, the value retaining certain thickness between described circumferential cut and described blast tube is 0.5 ~ 1mm.
6. there is the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 2, it is characterized in that, in described circumferential cut, tinsel is installed.
7. have the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 6, it is characterized in that, the material of described tinsel is different according to operating temperature level, selects common stainless steel or high-temperature alloy steel.
8. have the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 6, it is characterized in that, the thickness vertically of described tinsel is suitable with described circumferential cut thickness vertically.
9. have the gas turbine engine integrated type nozzle ring of thermal distortion compensation structure according to claim 6, it is characterized in that, the gap between described tinsel and the circumferential wall of described circumferential cut is 0.1 ~ 0.2mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410666630.2A CN104533541B (en) | 2014-11-20 | 2014-11-20 | A kind of gas-turbine unit monoblock type nozzle ring with thermal distortion compensation structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201410666630.2A CN104533541B (en) | 2014-11-20 | 2014-11-20 | A kind of gas-turbine unit monoblock type nozzle ring with thermal distortion compensation structure |
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| Publication Number | Publication Date |
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| CN104533541A true CN104533541A (en) | 2015-04-22 |
| CN104533541B CN104533541B (en) | 2016-06-08 |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109404051A (en) * | 2018-12-29 | 2019-03-01 | 中国科学院工程热物理研究所 | A kind of the floating positioning and torsion pass structure of nozzle ring |
| CN112177688A (en) * | 2020-09-28 | 2021-01-05 | 宁国市华成金研科技有限公司 | Engine precision casting guider and machining method thereof |
| CN114109521A (en) * | 2021-11-26 | 2022-03-01 | 中国科学院工程热物理研究所 | Gas turbine guider connecting structure for reducing thermal stress |
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| CN101059081A (en) * | 2006-03-27 | 2007-10-24 | 阿尔斯托姆科技有限公司 | Turbine blade and diaphragm construction |
| CN101338688A (en) * | 2008-08-15 | 2009-01-07 | 中国航空动力机械研究所 | Gas-turbine unit turbine guider link construction |
| CN101652534A (en) * | 2007-06-22 | 2010-02-17 | 三菱重工业株式会社 | Stator blade ring and axial flow compressor using the same |
| US20100124492A1 (en) * | 2008-11-17 | 2010-05-20 | Honeywell International Inc. | Turbine nozzles and methods of manufacturing the same |
| CN102140936A (en) * | 2010-01-29 | 2011-08-03 | 通用电气公司 | Retaining ring for a turbine nozzle with improved thermal isolation |
-
2014
- 2014-11-20 CN CN201410666630.2A patent/CN104533541B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101059081A (en) * | 2006-03-27 | 2007-10-24 | 阿尔斯托姆科技有限公司 | Turbine blade and diaphragm construction |
| CN101652534A (en) * | 2007-06-22 | 2010-02-17 | 三菱重工业株式会社 | Stator blade ring and axial flow compressor using the same |
| CN101338688A (en) * | 2008-08-15 | 2009-01-07 | 中国航空动力机械研究所 | Gas-turbine unit turbine guider link construction |
| US20100124492A1 (en) * | 2008-11-17 | 2010-05-20 | Honeywell International Inc. | Turbine nozzles and methods of manufacturing the same |
| CN102140936A (en) * | 2010-01-29 | 2011-08-03 | 通用电气公司 | Retaining ring for a turbine nozzle with improved thermal isolation |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN109404051A (en) * | 2018-12-29 | 2019-03-01 | 中国科学院工程热物理研究所 | A kind of the floating positioning and torsion pass structure of nozzle ring |
| CN112177688A (en) * | 2020-09-28 | 2021-01-05 | 宁国市华成金研科技有限公司 | Engine precision casting guider and machining method thereof |
| CN114109521A (en) * | 2021-11-26 | 2022-03-01 | 中国科学院工程热物理研究所 | Gas turbine guider connecting structure for reducing thermal stress |
| CN114109521B (en) * | 2021-11-26 | 2024-04-12 | 中国科学院工程热物理研究所 | Gas turbine guide connection structure for reducing thermal stress |
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| CN104533541B (en) | 2016-06-08 |
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