CN1497131A - Method and device for preventing damaging blade of gas turbine engine - Google Patents
Method and device for preventing damaging blade of gas turbine engine Download PDFInfo
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- CN1497131A CN1497131A CNA2003101024132A CN200310102413A CN1497131A CN 1497131 A CN1497131 A CN 1497131A CN A2003101024132 A CNA2003101024132 A CN A2003101024132A CN 200310102413 A CN200310102413 A CN 200310102413A CN 1497131 A CN1497131 A CN 1497131A
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Abstract
A method of fabricating a rotor assembly for a gas turbine engine is presented. The method includes forming a blade (30) including an airfoil extending from an integral dovetail (52) used to mount the blade within the rotor assembly, and extending a projection (94) from at least a portion (86) of the blade, such that the stresses induced within at least a portion of the blade are facilitated to be maintained below a predetermined failure threshold for the blade to facilitate preventing failure of the blade.
Description
Invention field
The present invention relates generally to the blade of some gas turbine engines, more particularly, relates to the method and apparatus that some blades that help preventing gas turbine engine damage.
Background technique
At least some known gas turbine engines comprise core-engine, and this motor flow arrangement in order has fan component and high pressure compressor, and this high pressure compressor compresses into the air-flow in the motor.Burner is lighted by turbine nozzle assembly and is directed to fuel-air mixture in low-pressure turbine and the high-pressure turbine, wherein each in these low-pressure turbines and the high-pressure turbine comprises some rotor blades, the rotating energy in the air-flow that these rotor blades absorb to come out in burner.
Intrasystem damage parts can greatly be damaged this system and/or this intrasystem other part, and also may need system to quit work when changing or repairing the part of this damage.In particular, when this part was the fan blade of turbofan formula gas turbine engine, leaf abscission can cause that the blade that is positioned at the blade downstream part that is disconnected suffers damage.More particularly, according to the infringement order of severity to downstream blade, other blade that is positioned at the downstream part of trailing blade of the blade that disconnected or damage also may be damaged.The damage of trailing blade may cause trails blade fault, therefore may need turbofan formula gas turbine engine to quit work, and/or damages other fan blade and/or other part in the turbofan formula gas turbine engine.
For example, at least some known turbofan formula gas turbine engines comprise the fan matrix, and this fan matrix has the plurality of fans blade, and these fan blade are extended radially outwardly.May cause on the blade that the blade bump that disconnects is trailed that trailing blade shroud swings around rotating tangent axis with fan.Trail blade and swing towards this front portion of trailing blade around this tangent axis at first, so this is trailed blade and can move apart its dish-shaped groove radially outwardly.It is reverse owing to the rotation of fan around the motion of this tangent axis to trail blade, thereby causes that trailing blade swings backward towards this rear end of trailing blade.The swing of blade can produce pressure stress and tensile stress in blade.Trail the tensile stress in the blade and the size of pressure stress and may surpass the damaging thresholding of blade material, thereby cause that this trails blade and damage.
Summary of the invention
According to one aspect of the present invention, provide a kind of manufacturing to be used for the method for the fan component of gas turbine engine.This method comprises: form blade, wherein this blade comprises a wing, this aerofoil extends from the all-in-one-piece dovetail, this dovetail is used for a blade installation in rotor assembly, and protuberance is extended from least a portion of blade, make the stress that at least a portion of blade, is produced help remaining under the predetermined damaging thresholding of this blade, thereby help preventing that this blade from damaging.
According to another aspect of the present invention, a kind of blade of gas turbine engine is provided, this blade comprises: aerofoil; One dovetail, it and described aerofoil form one; And a protuberance, its at least one from aerofoil and dovetail stretches out.This protuberance is shaped as and helps the motion of limit blade at least in part, thereby helps preventing that this blade from damaging.
According to another aspect of the present invention, provide a kind of fan component that is used for gas turbine engine.This fan component comprises fan hub and at least one fan blade, and this fan blade extends radially outwardly from fan hub.This fan blade comprises: a dovetail; One aerofoil, it stretches out from dovetail; And protuberance, it stretches out from dovetail, thus the stress that is produced at least one in dovetail and aerofoil of handle remains under the predetermined damaging thresholding of fan blade.
The accompanying drawing summary
Fig. 1 is the schematic representation of exemplary turbofan formula gas turbine engine;
Fig. 2 is the perspective view of a part of exemplary fan blade, and wherein this fan blade can be included in the turbofan formula gas turbine engine shown in Figure 1;
Fig. 3 is the sectional elevation along line 3-3 a part of fan component that intercepted, shown in Figure 1 of Fig. 2; And
Fig. 4 is the sectional elevation along line 4-4 a part of fan component that intercepted, shown in Figure 3 of Fig. 3.
Detailed description of the present invention
As used herein the same, term " inefficacy " and " fault " can comprise any damage or damage a part at least partly away from other situation that runs well, for example, any damage or damage a part at least partly and can comprise fracture fully, local fracture, the change of this part shape and the change of performance of this part away from other situation that runs well, but be not limited to these.Above these examples just hope as exemplary, therefore never want to limit the definition and/or the implication of term " inefficacy " and " fault ".In addition, although describe the present invention in conjunction with turbofan formula gas turbine engine here, and more specifically the fan blade in turbofan formula gas turbine engine is used, and should be understood that the present invention can be applied in any part.Therefore, enforcement of the present invention is not limited to fan blade or other part of turbofan formula gas turbine engine.
Fig. 1 is the schematic representation of turbofan formula gas turbine engine 10, and wherein motor 10 comprises fan component 12, high pressure compressor 14 and burner 16.Motor 10 also comprises high-pressure turbine 18, low-pressure turbine 20 and booster rocket 22.Fan component 12 comprises fan hub 24, and this fan hub 24 has some dish-shaped grooves (not illustrating) in Fig. 1, and these dish-shaped grooves are positioned at fan hub and separate along circumferencial direction around fan hub 24.Fan component 12 also comprises a ventilating fan blade 30, and these blades 30 extend to the aerofoil top 32 of fan blade radially outwardly from dish-shaped groove and fan hub 24.Fan component 12 is around spin axis 40 rotations.Motor 10 has air inlet side 42 and exhaust side 44.In one embodiment, motor 10 is GE-90 motors, and this motor can have been bought from the General Electric Aircraft Engines company of Cincinnati state Ohio commercial.
In when work, air stream passing through fan the assembly 12 and air that compressed supplied in the high pressure compressor 14.The air that fully compressed is transported in the burner 16, and this air mixes with fuel and lights there.Combustion gas are drawn from burner 16 and are used for driving turbine 18 and 20, and turbine 20 drive fan assemblies 12.
Fig. 2 is the perspective view of a part of example fan blade 30, and wherein fan blade 30 can be used with fan component 12 (not illustrating in Fig. 1).Each blade 30 comprises the aerofoil 50 and the all-in-one-piece dovetail 52 of hollow, and wherein dovetail 52 is used in known manner aerofoil 50 being installed on the fan hub 24.Each aerofoil 50 comprises that first sidewall 54 and second of making profile makes the sidewall 56 of profile.The first side wall 54 is convex and suction side that limit aerofoil 50, second sidewall 56 be spill and limit aerofoil 50 on the pressure side.Sidewall 54 and 56 is joined together in the leading edge 58 of aerofoil 50 and trailing edge 60 places that separate vertically.More particularly, the trailing edge 60 of aerofoil separates along tangential leading edge 58 at downstream part and aerofoil.First and second sidewalls 54 and 56 correspondingly extend to the top 32 of aerofoil (being illustrated among Fig. 1) from root of blade 62 along the longitudinal or radially outwardly, and wherein root 62 is arranged in the contiguous dovetail 52.Length that fan blade 30 is extended 64 for from front end 66 to the rear end 68.Dovetail 52 comprises the contact surface 70 of first pressure surface and the contact surface 72 of second pressure surface.
Fig. 3 is the sectional elevation of a part of fan component 12 of being intercepted of the line 3-3 along Fig. 2.Fig. 4 is the sectional elevation of a part of fan component 12 of being intercepted of the line 4-4 along Fig. 3.Specifically, in Fig. 3 and 4, fan blade 30 is connected in the fan hub 24.More particularly, fan blade 30 is installed and is fixed in the dish-shaped groove 74, here is also referred to as to be placed in the dish-shaped groove 74, and wherein dish-shaped groove 74 limits in fan hub 24.In one embodiment, fan hub 24 comprises some dish-shaped grooves 74, and these grooves are formed in the fan hub 24 and are spaced along circumferencial direction around fan hub 24.
Each dish-shaped groove 74 extends a length 64 at least, so each dovetail 52 is installed in wherein fully.In the time of in each fan blade dovetail 52 is placed in corresponding dish-shaped groove 74, each fan blade 30 is extended radially outwardly from fan hub 24.Dish groove 74 comprises internal surface 76 radially, the part 78 of dish groove 74 is shaped as with the part of dovetail 52 complementary, therefore in the time of in dovetail 52 is placed in dish-shaped groove 74, the pressure surface 82 of the contact surface 72 contacts second dish-shaped groove of pressure surface 80, the second pressure surfaces of the contact surface 70 contiguous first dish-shaped grooves of first pressure surface.
In this exemplary embodiment, dovetail 52 comprises blade spacer region 84, and this spacer region 84 stretches out from the inner radial surface 86 of dovetail 52.Perhaps, dovetail 52 does not comprise spacer region 84.More particularly, spacer region 84 radially inwardly extends towards the inner radial surface 76 of fan hub 24 and dish-shaped groove.In the time of in fan blade 30 is placed in dish-shaped groove 74, blade spacer region 84 extends a distance 88 from the inner radial surface 86 of dovetail, has therefore limited minimum blade/dish type radial clearance 90 between the inner radial surface 76 of the inner radial surface 92 of spacer region 84 and dish-shaped groove.In this example embodiment, blade spacer region 84 is crossed over the length 64 of fan blade basically and is extended.Perhaps, in another embodiment, 84 parts of crossing over fan blade length 64 of blade spacer region.Extend in this exemplary embodiment, blade spacer region 84 is the individual part that connect dovetail 52.In this alternative embodiment, blade spacer region 84 forms one with the dovetail 52 of fan blade.
The dovetail 52 of fan blade also comprises protuberance 94, and this protuberance 94 stretches out from blade spacer region 84.More particularly, protuberance 94 from dovetail 52 towards axis 40, the inner radial surface 76 of fan hub 24 and dish-shaped groove radially inwardly extends.In the time of in fan blade 30 is placed in dish-shaped groove 74, protuberance 94 is arranged to 92 1 distances 96 of inner radial surface apart from the blade spacer region, makes the radial clearance 98 that has limited protuberance/dish-shaped groove between the radial surface 100 of the inner radial surface 76 of dish-shaped groove and protuberance 94.In one embodiment, gap 90 approximates 0.190 inch greatly, and gap 98 approximates 0.040 inch greatly.
In this exemplary embodiment, protuberance 94 is independent components, and it is connected to or rubs and is connected on the blade spacer region 84.In an alternative embodiment, protuberance 94 forms one with blade spacer region 84.In one embodiment, fan blade 30 does not comprise blade spacer region 84, and protuberance 94 outwards extends towards the inner radial surface 76 of axis 40, fan hub 24 and dish-shaped groove from the inner radial surface 86 of dovetail.In an alternative embodiment, fan blade 30 does not comprise blade spacer region 84, and protuberance 94 forms one with dovetail 52 or is connected on the dovetail 52.Protuberance 94 extends a distance 102 from fan blade rear end 68 towards fan blade front end 66.Although being shown to from the rear end 68 here, protuberance 94 extends a distance 102 towards front end 66, but will be appreciated that, protuberance 94 can be arranged on any position on the inner radial surface 92 of blade spacer region so that preventing fan blade 30 lost efficacy, and this point below will be described.For example, in an alternative embodiment, protuberance 94 is arranged to the front end 66 of contiguous fan blade.
Fan component 12 comprises axis 104, and axis 104 is tangent with the inner radial surface 76 of dish-shaped groove.Although axis 104 is shown to the head center of passing fan blade length 64 here, should be understood that axis 104 can pass any part of blade 30 along length 64, and tangent with the inner radial surface 76 of dish-shaped groove.
During the rotation of fan component 12, when the blade on the fan hub 24 that is installed to blade 30 upstream ends broke down or throw away (situation that promptly is called " blade is deviate from " here) from its corresponding dish-shaped groove, the part of this fan blade can be clashed into fan 30.This contact can cause fan blade 30 swings or be rotated around axis 104.Specifically, at first, therefore fan blade 30 radially inwardly forces the inner radial surface 76 of front end 66 towards dish-shaped groove, and forces the inner radial surface 76 of fan blade rear end 68 away from dish-shaped groove radially outwardly towards fan blade front end 66, be rotated around axis 104.More particularly, this bump can cause that fan blade front end 66 parts leave dish-shaped groove 74.When by disconnecting that blade clashes into that caused stress fluctuation is reflected by blade 30 and when transmitting, be reversed around rotatablely moving of axis 104, therefore cause that fan blade 30 rotates towards the rear end 68 of fan blade, so that force the inner radial surface 76 of fan blade front end 66 radially outwardly, and radially inwardly force the inner radial surface 76 of fan blade rear end 68 towards dish-shaped groove away from dish-shaped groove.More particularly, fan blade rear end 68 can be left dish-shaped groove 74 partly.
When fan blade rear end 68 was left dish-shaped groove 74 at least in part, the pressure between the contact surface 72 of pressure between the contact surface 70 of first pressure surface of fan blade and the pressure surface 80 of the first dish-shaped groove and second pressure surface of fan blade and the pressure surface 82 of the second dish-shaped groove concentrated on the front end 66 of fan blade.More particularly, relatively large pressure stress can concentrate on the rear end 68 of fan blade, and relatively large tensile stress can concentrate on the front end 66 of fan blade.The tensile stress in the fan blade 30 and the size of pressure stress can surpass the predetermined damaging thresholding of at least a portion fan blade 30, cause that therefore fan blade 30 partly or completely damages.But protuberance 94 has limited the motion of fan blade 30, and more particularly, restriction fan blade 30 is rotated around axis 104, therefore helps reducing issuable tensile stress in fan blade front end 66.More particularly, when dish-shaped groove 74 was left in fan blade rear end 68, therefore inwardly moving radially of protuberance 94 partial restriction fan blade rear ends 68 had only the tensile stress of limited size can concentrate on the front end 66 of fan blade.Correspondingly, protuberance 94 helps the stress intensity in the fan blade 30 is remained under the damaging thresholding of fan blade 30.
The expense of said apparatus is lower, and reliability is higher, thereby helps preventing damage parts.This device helps the stress that is produced at least a portion of part is remained under the predetermined damaging thresholding of this part.More particularly, this device has limited the motion of part at least in part, thereby makes under the damaging thresholding that tensile stress in this part and pressure stress remain on this part.Consequently, this device helps preventing damage parts in the mode of cheapness and reliable operation.
Described the exemplary embodiment of blade and assembly in the above in detail.These systems are not limited to specific embodiment as described herein, and on the contrary, some parts of each assembly can use dividually, independently with other part as described herein.The part of each blade and assembly also can combine use with the part of other device and assembly.
Although described the present invention with regard to various specific embodiments, those of ordinary skills should be understood that the present invention can realize by the distortion in the spirit and scope that are in claims.
Claims (20)
1. a manufacturing is used for the method for gas turbine engine (10) rotor assembly (12), and described method comprises:
Form blade (30), wherein this blade comprises an aerofoil (50), and this aerofoil (50) extends from all-in-one-piece dovetail (52), and this dovetail (52) is used for a blade installation in rotor assembly; And
One protuberance (94) is extended from least a portion of blade, make the stress that is produced at least a portion (86) at blade help remaining under the predetermined damaging thresholding of this blade, thereby help preventing that this blade from damaging.
2. the method for claim 1, it is characterized in that, the step that protuberance (94) is extended from least a portion (86) of blade comprises: protuberance is connected at least a portion of this blade, this protuberance is stretched out from least a portion of this blade.
3. the method for claim 1, it is characterized in that, the step that protuberance (94) is extended from least a portion (86) of blade (30) comprises: at least a portion at this blade is integrally formed into protuberance, and this protuberance is stretched out from least a portion of this blade.
4. the method for claim 1 is characterized in that, the step that protuberance (94) is extended from least a portion (86) of blade (30) comprises: utilize this protuberance to be beneficial to limit at least in part the motion of at least a portion of this blade.
5. the method for claim 1, it is characterized in that the step that protuberance (94) is extended from least a portion (86) of blade (30) comprises: utilize this protuberance to be beneficial to the tensile stress at least a portion of this blade is remained under the predetermined damaging thresholding of this blade.
6. the method for claim 1 is characterized in that, the step that protuberance (94) is extended from least a portion (86) of blade (30) comprises: utilize this protuberance to be beneficial to limit at least in part at least a portion generation rotation of this blade.
7. the method for claim 1, it is characterized in that, the step that protuberance (94) is extended from least a portion (86) of blade (30) comprises: during at least a situation damage of second gas turbine engine blade and second gas turbine engine blade generation leaf abscission both of these case, utilize this protuberance to be beneficial to the stress at least a portion of this blade is remained under the predetermined damaging thresholding of this blade.
8. method as claimed in claim 7, utilize protuberance (94) to keep the step of the stress at least a portion of this blade (30) to comprise: damage and during second gas turbine engine blade produces at least a situation in the leaf abscission both of these case at second gas turbine engine blade, at least a portion of utilizing this protuberance to be beneficial to limit at least in part this blade produces moves.
9. the blade of a gas turbine engine (30), this blade comprises:
One aerofoil (50);
One dovetail (52), it and described aerofoil form one; And
Protuberance (94), its at least one from described aerofoil and described dovetail stretches out, and described protuberance is shaped as the motion that helps limiting at least in part described blade, thereby helps preventing that described blade from damaging.
10. blade as claimed in claim 9 (30), it is characterized in that described protuberance (94) also is shaped as the stress that helps being produced at least one in described aerofoil (50) and described dovetail (52) of handle and remains under the predetermined damaging thresholding of described blade.
11. blade as claimed in claim 10 (30), it is characterized in that described protuberance (94) also is shaped as the tensile stress that helps at least one in described aerofoil (50) and described dovetail (52) of handle and remains under the predetermined damaging thresholding of described blade.
12. blade as claimed in claim 9 (30) is characterized in that, described protuberance (94) extends from described dovetail (52) radially outwardly.
13. the fan component of a gas turbine engine (12), described fan component comprises:
One fan hub (24); With
At least one fan blade (30), this fan blade extends radially outwardly from described fan hub, and described fan blade comprises: a dovetail (52); One aerofoil (50), it stretches out from described dovetail; And protuberance (94), it stretches out from described dovetail, thus the stress that is produced at least one in described dovetail and described aerofoil of handle remains under the predetermined damaging thresholding of described fan blade.
14. fan component as claimed in claim 13 (12), it is characterized in that, described protuberance (94) is shaped as the motion that helps limiting at least in part described fan blade (30), and the stress that helps being produced at least one in described fan blade aerofoil (50) and described fan blade dovetail (52) of handle is remained under the predetermined damaging thresholding of described fan blade.
15. fan component as claimed in claim 13 (12) is characterized in that, described protuberance (94) is connected on the described dovetail (52).
16. fan component as claimed in claim 13 (12) is characterized in that, described protuberance (94) forms one with described dovetail (52).
17. fan component as claimed in claim 13 (12) is characterized in that, described dovetail (52) comprises a spacer region (84), and this spacer region (84) stretches out from described dovetail, and described protuberance stretches out from described spacer region.
18. fan component as claimed in claim 13 (12), it is characterized in that, described fan blade (30) also comprises an aerofoil top (32), described aerofoil (50) extends between described dovetail (52) and described aerofoil top, and described protuberance (94) stretches out from described dovetail section along the direction away from described aerofoil top.
19. fan component as claimed in claim 13 (12), it is characterized in that, described fan hub (24) comprises at least one dish-shaped groove (74), described dovetail (52) is installed in the described dish-shaped groove at least in part, described fan blade (30) is fixed with respect to described fan matrix, described protuberance (94) radially extends to the described dish-shaped groove from described dovetail, moves in described dish-shaped groove thereby help limiting at least in part described fan blade.
20. fan component as claimed in claim 19 (12) is characterized in that, described protuberance (94) also is shaped as and helps limiting at least in part the rotation of described fan blade (30) with respect to described dish-shaped groove (74).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CNA2003101024132A CN1497131A (en) | 2002-10-18 | 2003-10-17 | Method and device for preventing damaging blade of gas turbine engine |
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US10/273969 | 2002-10-18 | ||
| CNA2003101024132A CN1497131A (en) | 2002-10-18 | 2003-10-17 | Method and device for preventing damaging blade of gas turbine engine |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN1497131A true CN1497131A (en) | 2004-05-19 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CNA2003101024132A Pending CN1497131A (en) | 2002-10-18 | 2003-10-17 | Method and device for preventing damaging blade of gas turbine engine |
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| CN (1) | CN1497131A (en) |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1854465B (en) * | 2005-04-28 | 2010-11-03 | 通用电气公司 | Finger dovetail attachment between a turbine rotor wheel and bucket for stress reduction |
| CN102121400A (en) * | 2009-11-10 | 2011-07-13 | 阿尔斯托姆科技有限公司 | Rotor for an axial-throughflow turbomachine and moving blade for such a rotor |
| US8272841B2 (en) | 2006-11-02 | 2012-09-25 | Ge Aviation Uk | Propeller blade retention |
| CN103244198A (en) * | 2012-02-10 | 2013-08-14 | 通用电气公司 | Turbine assembly |
| CN104213941A (en) * | 2013-05-29 | 2014-12-17 | 阿尔斯通技术有限公司 | Blade of a turbine |
| CN104712374A (en) * | 2013-12-17 | 2015-06-17 | 通用电气公司 | Rotor wheel assembly and assembling method thereof and corresponding turbine engine |
-
2003
- 2003-10-17 CN CNA2003101024132A patent/CN1497131A/en active Pending
Cited By (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1854465B (en) * | 2005-04-28 | 2010-11-03 | 通用电气公司 | Finger dovetail attachment between a turbine rotor wheel and bucket for stress reduction |
| US8272841B2 (en) | 2006-11-02 | 2012-09-25 | Ge Aviation Uk | Propeller blade retention |
| CN101553399B (en) * | 2006-11-02 | 2013-10-02 | Ge航空英国公司 | Propeller blade retention |
| CN102121400A (en) * | 2009-11-10 | 2011-07-13 | 阿尔斯托姆科技有限公司 | Rotor for an axial-throughflow turbomachine and moving blade for such a rotor |
| CN102121400B (en) * | 2009-11-10 | 2015-12-16 | 阿尔斯托姆科技有限公司 | For the rotor of axial flow formula turbo machine and the moving vane for this rotor |
| CN103244198A (en) * | 2012-02-10 | 2013-08-14 | 通用电气公司 | Turbine assembly |
| CN104213941A (en) * | 2013-05-29 | 2014-12-17 | 阿尔斯通技术有限公司 | Blade of a turbine |
| CN104213941B (en) * | 2013-05-29 | 2016-03-23 | 阿尔斯通技术有限公司 | The blade of turbine |
| CN104712374A (en) * | 2013-12-17 | 2015-06-17 | 通用电气公司 | Rotor wheel assembly and assembling method thereof and corresponding turbine engine |
| CN104712374B (en) * | 2013-12-17 | 2018-07-17 | 通用电气公司 | Rotor wheel assembly and its assemble method and corresponding turbogenerator |
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