CN103184889A - Turbine rotor blade, rotating assembly and turbine engine - Google Patents
Turbine rotor blade, rotating assembly and turbine engine Download PDFInfo
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- CN103184889A CN103184889A CN2011104487056A CN201110448705A CN103184889A CN 103184889 A CN103184889 A CN 103184889A CN 2011104487056 A CN2011104487056 A CN 2011104487056A CN 201110448705 A CN201110448705 A CN 201110448705A CN 103184889 A CN103184889 A CN 103184889A
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Abstract
The invention relates to a turbine rotor blade, which comprises a wing-shaped part and a blade crown part, wherein the blade crown part is formed on the free end of the wing-shaped part through generally, circumferentially and axially extending along the center of a turbine rotor, and is provided with a blade crown front edge and a blade crown back edge, wherein the blade crown front edge and the blade crown back edge are respectively in smoothly transited arc shapes. The blade crown front edge and the blade crown back edge are designed into arc shapes, the processibility of the blade crown part can be improved, and meanwhile, the weight distribution of the blade crown part can be more uniform through the arc-shaped blade crown front edge and the arc-shaped blade crown back edge, so the stress concentration effect is reduced.
Description
Technical field
The present invention relates to a kind of turbine rotor blade, Runner assembly and turbogenerator.
Background technique
In a lot of fields, combustion gas turbine is used for driving generator or work mechanism, especially aboard.At this, utilize the energy of fuel to make turbine shaft rotate.For this reason, fuel burns in a firing chamber, is also imported wherein by the air after the gas compressor compression.In the firing chamber, form the High Temperature High Pressure working medium by a turbine unit expansion acting that is connected on the back, firing chamber by fuel combustion.
Turbine rotor blade is a vitals in the turbine unit, and it is used for the mechanical work that becomes turbine rotor from the thermal power transfer of firing chamber.The performance of turbine can be by being strengthened turbine bucket tip sealing, thereby prevent that combustion gas from escaping into gap between vane tip and the casing from air-flow.Usually the encapsulating method with the gap between vane tip and the turbine casing is to be provided with integral shroud at vane tip.Integral shroud not only can strengthen turbine performance can also be used as vibration damper, particularly especially true for the long turbine blade of radial length.The turbine blade that typically has integral shroud has airfoil section and integral shroud part.Generally during fabrication, airfoil section and integral shroud part are integrally formed forms as casting.Wherein, the airfoil-shaped portion pressure side and the suction surface that have leading edge, trailing edge and extend to trailing edge from leading edge; Integral shroud part is extended perpendicular to the leading edge of airfoil-shaped portion and trailing edge and at the top of airfoil-shaped portion substantially, it has integral shroud leading edge and integral shroud trailing edge, form interlocking structure between the integral shroud trailing edge of the integral shroud part of the turbine rotor blade that this integral shroud leading edge and integral shroud trailing edge and are adjacent and the integral shroud leading edge of another adjacent turbine rotor blade, typically, this integral shroud leading edge and integral shroud trailing edge are formed with zigzag fashion, specifically, this zigzag fashion is generally Z-shaped.By the circumferential integral shroud group that this interlocking links together, vibration can be lowered and turbine efficiency is improved.Generally partly also be provided with envelope gas comb tooth away from an example of turbine shaft at integral shroud, it is perpendicular to the turbine axis, and the purpose of envelope gas comb tooth is further to make the top leakage on every side of turbine rotor blade to reduce to minimum.
The Chinese invention patent application of publication number CN101372895A discloses a kind of edge contour of turbine bucket integral shroud, and the profile of this integral shroud is Z-shaped.Yet clearly, this integral shroud difficulty of processing with Z-shaped profile is bigger, and in addition, thereby also the inhomogeneous partial structurtes cantilever amount that causes is big and caused stress to concentrate for mass distribution, therefore the working life of having reduced turbine rotor blade.Therefore, wish to improve the profile of turbine blade tip shroud, with the reduction creep, thereby optimize the wheel blade life-span, guarantee that simultaneously wheel blade has acceptable performance and manufacturability.
Summary of the invention
An object of the present invention is to provide a kind of turbine rotor blade, it has improved integral shroud structure, by this, the mass distribution of integral shroud part is tended to balance more with rationally.To achieve these goals, the invention discloses a kind of turbine rotor blade, comprise airfoil section and integral shroud part.The integral shroud part circumferentially and is axially extended along the turbine rotor center substantially and is formed on the free end of airfoil section, and integral shroud partly has integral shroud leading edge and integral shroud trailing edge.Wherein, integral shroud leading edge and integral shroud trailing edge all are the arc that seamlessly transits.
More preferably, the curve that forms arc is consistent with the mean camber line of the aerofoil profile closed curve of airfoil section substantially.
More preferably, integral shroud leading edge and integral shroud trailing edge are basic symmetries with respect to the mean camber line of the aerofoil profile closed curve of airfoil section.
More preferably, make progress in week, project to the integral shroud leading edge from what the mean camber line of the aerofoil profile closed curve of airfoil section radially was incident upon the integral shroud part, and project to integral shroud trailing edge, the radial thickness attenuation gradually of integral shroud part from what the mean camber line of the aerofoil profile closed curve of airfoil section radially was incident upon the integral shroud part.
Optionally, make progress in week, from the free-ended connection area of integral shroud part and airfoil section to the integral shroud leading edge and from the free-ended connection area of integral shroud part and airfoil section to the integral shroud trailing edge, the radial thickness attenuation gradually of integral shroud part.
More preferably, integral shroud leading edge and integral shroud trailing edge have all been set up wear resistant coating.
More preferably, integral shroud partly has along circumferential and/or axially extended stiffening rib.
The invention also discloses a kind of Runner assembly, be used for gas turbine engine, it comprises: rolling disc, blade body and integral shroud part.Wherein, blade body comprises: blade inlet edge, trailing edge, be fixed on the blade root on the such rotation dish, with the blade root opposed free ends, and be formed on pressure side and suction surface between blade inlet edge and the trailing edge.The integral shroud part is formed on the free end perpendicular to blade body substantially, and it has integral shroud leading edge and integral shroud trailing edge.Wherein, the integral shroud leading edge is the arc that seamlessly transits with the integral shroud trailing edge and adjacent integral shroud part is in the same place by adjacent arc is interlocked with one another.
More preferably, pressure side and suction surface have formed the aerofoil profile closed curve jointly on perpendicular to the cross section of blade inlet edge and trailing edge, and the curve that forms arc is basic consistent with the mean camber line of aerofoil profile closed curve.
More preferably, integral shroud leading edge and integral shroud trailing edge are basic symmetries with respect to the mean camber line of the aerofoil profile closed curve of blade body.
More preferably, the radial thickness of integral shroud part from the centre of integral shroud part to the direction of integral shroud leading edge and integral shroud trailing edge on attenuation gradually.
The invention also discloses a kind of turbogenerator, it has above-mentioned Runner assembly.
Integral shroud leading edge and integral shroud trailing edge are designed to arc, the workability of integral shroud part is improved, thereby the integral shroud leading edge of arc and integral shroud trailing edge make the mass distribution of integral shroud part more evenly reduce stress and concentrate effect simultaneously, in addition, attenuation and set up the requirement of strength that strengthening rib can not only satisfy the integral shroud part simultaneously and simultaneously the quality of integral shroud part is reduced gradually on the radial thickness of integral shroud part being designed to from the centre of integral shroud part to the direction of integral shroud leading edge and integral shroud trailing edge.Therefore, the present invention can reduce the load level of turbine rotor blade and increase its working life.
Description of drawings
Fig. 1 is three turbine rotor blades disposed adjacent schematic representation together in twos;
Fig. 2 is the schematic representation of the integral shroud part of a turbine rotor blade among Fig. 1;
Fig. 3 is the aerofoil profile closed curve of airfoil section of a turbine rotor blade among Fig. 1 and the schematic representation of mean camber line, wherein shows respectively by two kinds of formed mean camber lines of different modes with figure (b) to scheme (a).
Embodiment
As depicted in figs. 1 and 2, wherein, Fig. 1 shows three turbine rotor blades disposed adjacent schematic representation together in twos, and Fig. 2 is the schematic representation of the integral shroud part of a turbine rotor blade among Fig. 1.Three turbine rotor blades represent with reference character 100,200 and 300 that respectively the turbine rotor blade shown in the figure is of the present invention preferred embodiment a kind of, below, be that example describes with one of them turbine rotor blade 100.Turbine rotor blade 100 comprises airfoil section 102 and integral shroud part 104.Wherein, this airfoil section 102 has blade inlet edge 102a, trailing edge 102b, pressure side 102c and suction surface 102d; Integral shroud part 104 circumferentially and is axially extended along the turbine rotor center substantially and is formed on the free end of airfoil section 102, and it has integral shroud leading edge 106 and integral shroud trailing edge 108.At this, airfoil section 102 and integral shroud part 104 are integrally formed.Integral shroud leading edge 106 and integral shroud trailing edge 108 all are the arc that seamlessly transits.Because integral shroud part 104 needs and the turbine rotor blade 200 of adjacency, 300 integral shroud part 204,304 at the turbine rotor blade 100 that makes progress in week mesh together, so, the arc of integral shroud leading edge 106 needs and the arc of integral shroud trailing edge 108 is the interlocking shape, like this, be assembled on the rolling disc of turbine rotor when all turbine rotor blades that week is made progress after, thereby just all being engaged with each other, these integral shrouds parts are interlocked and airfoil section forms the fuel gas flow passage jointly.As shown in Figure 2, the shape of integral shroud leading edge 106 and integral shroud trailing edge 108 is identical, and namely integral shroud leading edge 106 and integral shroud trailing edge 108 respectively can be with all identical integral shroud part 204,304 integral shroud trailing edge and integral shroud leading edge be engaged with each other and circumferentially be interlocked with its shape.It will be understood by those skilled in the art that airfoil section and integral shroud part also can make respectively and can be to be fixed together such as Placements such as welding.
Again in conjunction with Fig. 3, wherein, with (a) and (b) two aerofoil profile closed curve L1 and mean camber line L2 that illustrate the airfoil section 104 of turbine rotor blade 100, aerofoil profile closed curve L1 shown in broken lines represents with solid line among two figure at (a) of Fig. 3 with (b) in Fig. 2 among Fig. 3.The airfoil section that so-called aerofoil profile closed curve refers to turbine rotor blade is perpendicular to the profilogram on the cross section of blade inlet edge and trailing edge.Mean camber line is limited in this aerofoil profile closed curve, and it can refer to the continous curve that the common line in the center of circle of a plurality of inscribed circles of aerofoil profile closed curve constitutes.Shown in figure (a) among Fig. 3, wherein show aerofoil profile closed curve L1, be L2 and string of a musical instrument L3 by the mean camber line that is joined together to form such as center of circle A such as inscribed circle L4 and L5.Mean camber line also can refer to perpendicular many straight lines of the string of a musical instrument of aerofoil profile closed curve respectively with the common continous curve that constitutes in center of the formed line segment of two intersection points of the curve intersection up and down of aerofoil profile closed curve.As shown in figure (b) among Fig. 3, wherein show aerofoil profile closed curve L1 and by such as the mean camber line L2 that is joined together to form perpendicular to the center B of the line segment L4 ' of string of a musical instrument L3 and L5 ' etc. respectively.At this, the arc curve that forms integral shroud leading edge 106 and integral shroud trailing edge 108 is consistent with the mean camber line L2 of the aerofoil profile closed curve L1 of this airfoil section 102 substantially.Therefore, the arc of the arc of integral shroud trailing edge 108 and integral shroud leading edge 106 is essentially identical.And integral shroud leading edge 106 and integral shroud trailing edge 108 more preferably are basic symmetrical with respect to the mean camber line L2 of the aerofoil profile closed curve L1 of blade body.Like this, integral shroud leading edge 106 is recess-like, integral shroud trailing edge 108 is convex shapes, thereby when turbine rotor blade 100 when upwards assembling with adjacent turbine rotor blade 200,300 in week, as shown in Figure 1, the integral shroud trailing edge of the integral shroud part 204 of adjacent turbine rotor blade 200 and the integral shroud leading edge 106 of turbine rotor blade 100 are engaged with each other, and the integral shroud leading edge of the integral shroud part 304 of adjacent turbine rotor blade 300 and the integral shroud trailing edge 108 of turbine rotor blade 100 are engaged with each other.The structure of above-mentioned integral shroud part not only is conducive to improve manufacturing, thereby and the arc that seamlessly transits of above-mentioned integral shroud part owing to have especially the curved profile that has with the mean camber line basically identical of aerofoil profile closed curve of integral shroud leading edge and integral shroud trailing edge make the cantilever amount comparatively even, therefore, the design of integral shroud part of the present invention can increase the working life of turbine rotor blade.
Again referring to Fig. 1 and Fig. 2, integral shroud part 104 also has a pair of envelope gas comb tooth 110 and several stiffening ribs 112, wherein, envelope gas comb tooth 110 also can be called as obturage comb tooth or sealing guide rail etc., it is convexly set in the upper surface of integral shroud part 104, integrally formed with integral shroud part 104, it is used for reducing the air loss that radial clearance causes.Owing to envelope gas comb tooth is set is being known in the art in order to improve turbine efficiency, so do not repeat them here.Have in several stiffening ribs 112 what be that parallel (can be described as circumferential stiffening rib) have with envelope gas comb tooth 110 is vertical (can be described as shaft orientation reinforcing rib) with envelope gas comb tooth 110, particularly, in the border that above-mentioned a pair of envelope gas comb tooth 110 forms, has a pair of shaft orientation reinforcing rib, thereby it is fixed together with envelope gas comb tooth 110 and has constituted the shape that is similar to ladder, in addition, also have a pair of circumferential stiffening rib that lays respectively at envelope gas comb tooth 110 both sides outside above-mentioned a pair of envelope gas comb tooth 110, this rear and front end to circumferential stiffening rib is connected respectively on the integral shroud leading edge 106 and integral shroud trailing edge 108 of integral shroud part 104.Be appreciated that, stiffening rib 112 should be lower than the height that above-mentioned envelope gas comb tooth 110 is given prominence to from integral shroud part 104 from integral shroud part 104 outstanding height, thereby can guarantee stiffening rib 112 not with the stator casing on the sealing configuration of obturaging bump mill, and, owing to adopted the stiffening rib form of cross network, guaranteed that not only the intensity of integral shroud part has reduced the weight of integral shroud part simultaneously.Those skilled in the art should be appreciated that the number of envelope gas comb tooth and stiffening rib is not limited at this, and stiffening rib can be non-parallel or non-perpendicular with envelope gas comb tooth also,, is not limited to circumferential stiffening rib and/or shaft orientation reinforcing rib that is.
For the mass distribution that makes integral shroud part 104 more reasonable, in a kind of more excellent mode of execution of the present invention, make progress in week, project to integral shroud leading edge 106 from what the mean camber line L2 of the aerofoil profile closed curve L1 of airfoil section 102 radially was incident upon integral shroud part 104, and project to integral shroud trailing edge 108, the attenuation gradually of the radial thickness of integral shroud part 104 from what the mean camber line L2 of the aerofoil profile closed curve L1 of airfoil section 102 radially was incident upon integral shroud part 104.In another embodiment of the invention, make progress in week, from the free-ended connection area of integral shroud part 104 and airfoil section 102 to integral shroud leading edge 106, and from the free-ended connection area of integral shroud part 104 and airfoil section 102 to integral shroud trailing edge 108, the attenuation gradually of the radial thickness of integral shroud part 104.Above the mode of executions divided of two kinds of zones only be exemplary, those skilled in the art should be appreciated that as long as the radial thickness of integral shroud part 104 is attenuation gradually from the centre of integral shroud part 104 to the direction of integral shroud leading edge 106 and integral shroud trailing edge 108.At this, " centre " can represent a medium line, and for example the mean camber line L2 of the aerofoil profile closed curve of airfoil section radially projects the projection line on the integral shroud part 104, also can represent a zone, for example above-mentioned connection area.Therefore, can guarantee that the center of gravity of integral shroud part 104 is in the centre of integral shroud part 104, simultaneously, guarantee when turbine rotor blade 100 rotates at a high speed, do not avoid this centrifugal force that the connection area of integral shroud part 104 and airfoil section 102 is produced destruction thereby the integral shroud leading edge 106 of its integral shroud part 104 and integral shroud trailing edge 108 can not produce too big centrifugal force.
For the vibration that the reduces adjacent integral shroud part wearing and tearing to the Surface of action of integral shroud leading edge and integral shroud trailing edge, can be on the Surface of action of integral shroud leading edge and integral shroud trailing edge wear-resistant coating.
The invention also discloses a kind of Runner assembly, be used for gas turbine engine, it comprises: rolling disc, blade body and integral shroud part.Wherein, this blade body comprises: blade inlet edge, trailing edge, be fixed on the blade root on the rolling disc, with the blade root opposed free ends, and be formed on pressure side and suction surface between blade inlet edge and the trailing edge.This integral shroud partly has all features of foregoing integral shroud part, particularly, it is formed on the free end perpendicular to blade body substantially, it has integral shroud leading edge and integral shroud trailing edge, wherein, the integral shroud leading edge is the arc that seamlessly transits with the integral shroud trailing edge and adjacent integral shroud part is in the same place by adjacent arc is interlocked with one another.
The invention also discloses a kind of turbogenerator, it has a kind of Runner assembly, and this Runner assembly comprises: rolling disc, blade body and integral shroud part.Wherein, this blade body comprises: blade inlet edge, trailing edge, be fixed on the blade root on the rolling disc, with the blade root opposed free ends, and be formed on pressure side and suction surface between blade inlet edge and the trailing edge.This integral shroud partly has all features of foregoing integral shroud part, particularly, it is formed on the free end perpendicular to blade body substantially, it has integral shroud leading edge and integral shroud trailing edge, wherein, the integral shroud leading edge is the arc that seamlessly transits with the integral shroud trailing edge and adjacent integral shroud part is in the same place by adjacent arc is interlocked with one another.。
Though foregoing description is to the present invention's detailed explanation of contrasting, these are just illustrative to the present invention, rather than limitation of the present invention, and any innovation and creation that do not exceed in the connotation of the present invention all fall within the scope of protection of the present invention.
Claims (12)
1. turbine rotor blade, comprise airfoil section and integral shroud part, described integral shroud part circumferentially and is axially extended along the turbine rotor center substantially and is formed on the free end of described airfoil section, described integral shroud partly has integral shroud leading edge and integral shroud trailing edge, wherein, described integral shroud leading edge and described integral shroud trailing edge all are the arc that seamlessly transits.
2. turbine rotor blade according to claim 1 is characterized in that, the curve that forms described arc is substantially consistent with the mean camber line of the aerofoil profile closed curve of described airfoil section.
3. turbine rotor blade according to claim 2 is characterized in that, described integral shroud leading edge and described integral shroud trailing edge are basic symmetries with respect to the mean camber line of the aerofoil profile closed curve of described airfoil section.
4. turbine rotor blade according to claim 3, it is characterized in that, make progress in week, radially be incident upon the described integral shroud leading edge that projects to of described integral shroud part from the mean camber line of the aerofoil profile closed curve of described airfoil section, and radially be incident upon the described integral shroud trailing edge that projects to of described integral shroud part, the radial thickness attenuation gradually of described integral shroud part from the mean camber line of the aerofoil profile closed curve of described airfoil section.
5. turbine rotor blade according to claim 3, it is characterized in that, make progress in week, from the free-ended connection area of described integral shroud part and described airfoil section to described integral shroud leading edge and from the free-ended connection area of described integral shroud part and described airfoil section to described integral shroud trailing edge, the radial thickness attenuation gradually of described integral shroud part.
6. turbine rotor blade according to claim 1 is characterized in that, described integral shroud leading edge and described integral shroud trailing edge have all been set up wear resistant coating.
7. turbine rotor blade according to claim 1 is characterized in that, described integral shroud partly has along circumferential and/or axially extended stiffening rib.
8. a Runner assembly is used for gas turbine engine, and it comprises:
Rolling disc;
Blade body, it comprises: blade inlet edge, trailing edge, be fixed on the blade root on the described rolling disc, with described blade root opposed free ends, and be formed on pressure side and suction surface between described blade inlet edge and the described trailing edge;
The integral shroud part, it is formed on the described free end perpendicular to described blade body substantially, it has integral shroud leading edge and integral shroud trailing edge, and wherein, described integral shroud leading edge is the arc that seamlessly transits with described integral shroud trailing edge and adjacent integral shroud part is in the same place by adjacent arc is interlocked with one another.
9. described Runner assembly according to Claim 8, it is characterized in that, described pressure side and described suction surface have formed the aerofoil profile closed curve jointly on perpendicular to the cross section of described blade inlet edge and described trailing edge, the curve that forms described arc is basic consistent with the mean camber line of described aerofoil profile closed curve.
10. according to the described Runner assembly of claim 9, it is characterized in that described integral shroud leading edge and described integral shroud trailing edge are basic symmetries with respect to the mean camber line of the aerofoil profile closed curve of described blade body.
11. according to the described Runner assembly of claim 10, it is characterized in that, the radial thickness of described integral shroud part from the centre of integral shroud part to the direction of described integral shroud leading edge and described integral shroud trailing edge on attenuation gradually.
12. a turbogenerator, it has each described Runner assembly among the claim 8-11.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201110448705.6A CN103184889B (en) | 2011-12-28 | A kind of turbine rotor blade, runner assembly and turbogenerator |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201110448705.6A CN103184889B (en) | 2011-12-28 | A kind of turbine rotor blade, runner assembly and turbogenerator |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN103184889A true CN103184889A (en) | 2013-07-03 |
| CN103184889B CN103184889B (en) | 2016-12-14 |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106870017A (en) * | 2015-09-23 | 2017-06-20 | 通用电气公司 | For the nozzle and nozzle assembly of gas-turbine unit |
| CN109630207A (en) * | 2018-12-10 | 2019-04-16 | 中国航发四川燃气涡轮研究院 | A kind of hollow turbine rotor blade with integral shroud reinforcing rib |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US1423466A (en) * | 1920-10-02 | 1922-07-18 | Westinghouse Electric & Mfg Co | Interlocking blade shroud |
| US3606574A (en) * | 1969-10-23 | 1971-09-20 | Gen Electric | Cooled shrouded turbine blade |
| US6030178A (en) * | 1998-09-14 | 2000-02-29 | General Electric Co. | Axial entry dovetail segment for securing a closure bucket to a turbine wheel and methods of installation |
| US20010019695A1 (en) * | 1999-11-01 | 2001-09-06 | Correia Victor H.S | Stationary flowpath components for gas turbine engines |
| CN1500969A (en) * | 2002-11-12 | 2004-06-02 | 通用电气公司 | Method and apparatus for reducing flow across compressor airfoil tips |
| CN201650377U (en) * | 2010-05-19 | 2010-11-24 | 中国航空动力机械研究所 | Blade with crest |
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US1423466A (en) * | 1920-10-02 | 1922-07-18 | Westinghouse Electric & Mfg Co | Interlocking blade shroud |
| US3606574A (en) * | 1969-10-23 | 1971-09-20 | Gen Electric | Cooled shrouded turbine blade |
| US6030178A (en) * | 1998-09-14 | 2000-02-29 | General Electric Co. | Axial entry dovetail segment for securing a closure bucket to a turbine wheel and methods of installation |
| US20010019695A1 (en) * | 1999-11-01 | 2001-09-06 | Correia Victor H.S | Stationary flowpath components for gas turbine engines |
| CN1500969A (en) * | 2002-11-12 | 2004-06-02 | 通用电气公司 | Method and apparatus for reducing flow across compressor airfoil tips |
| CN201650377U (en) * | 2010-05-19 | 2010-11-24 | 中国航空动力机械研究所 | Blade with crest |
Cited By (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN106870017A (en) * | 2015-09-23 | 2017-06-20 | 通用电气公司 | For the nozzle and nozzle assembly of gas-turbine unit |
| CN106870017B (en) * | 2015-09-23 | 2021-01-29 | 通用电气公司 | Nozzle and nozzle assembly for a gas turbine engine |
| CN109630207A (en) * | 2018-12-10 | 2019-04-16 | 中国航发四川燃气涡轮研究院 | A kind of hollow turbine rotor blade with integral shroud reinforcing rib |
| CN109630207B (en) * | 2018-12-10 | 2021-07-09 | 中国航发四川燃气涡轮研究院 | Hollow turbine rotor blade with blade shroud reinforcing ribs |
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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. |
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