CN102619570A - Turbine bucket for use in gas turbine engines and methods for fabricating the same - Google Patents

Turbine bucket for use in gas turbine engines and methods for fabricating the same Download PDF

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Publication number
CN102619570A
CN102619570A CN2012100592467A CN201210059246A CN102619570A CN 102619570 A CN102619570 A CN 102619570A CN 2012100592467 A CN2012100592467 A CN 2012100592467A CN 201210059246 A CN201210059246 A CN 201210059246A CN 102619570 A CN102619570 A CN 102619570A
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CN
China
Prior art keywords
turbine
radial distance
tip
guard shield
backplate
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Pending
Application number
CN2012100592467A
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Chinese (zh)
Inventor
S·K·贾因
R·N·苏哈
M·D·科利耶
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General Electric Co
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General Electric Co
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Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Publication of CN102619570A publication Critical patent/CN102619570A/en
Pending legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/34—Arrangement of components translated
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/38—Arrangement of components angled, e.g. sweep angle

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

A turbine bucket34 for use with a turbine engine 10. The turbine bucket includes a dovetail that is coupled to a rotor assembly that is positioned within a turbine casing40 . A platform extends from the dovetail. An airfoil extends from the platform. The airfoil includes a root end88 and a tip end90. The tip end90 extends outwardly from the root end towards the turbine casing. A tip shroud extends from the tip end. The tip shroud includes a shroud plate110. A first shroud rail 162 extending a first radial distance from the shroud plate towards the turbine casing. A second shroud rail extends a second radial distance164 from the shroud plate towards the turbine casing that is different than the first radial distance.

Description

Be used in turbine vane and production method thereof in the gas turbine engine
Technical field
The field of the invention relates in general to gas turbine engine, and especially relates to the turbine vane that uses with gas turbine engine.
Background technique
At least some known gas turbine engines comprise burner, the compressor that connects in the burner downstream, turbine and rotatably be connected the rotor assembly between compressor and the turbine.Some known rotor assembly comprise rotor shaft, at least one rotor disk that is connected with rotor shaft and a plurality of circumferentially spaced turbine vane that outwards stretches from each rotor disk.Each turbine vane comprises the aerofoil that extends radially outwardly from platform towards turbine shroud.
In the operating process of at least some known turbines, the compressor compresses air, its subsequently with the fuel mix that before was passed into burner.Mixture is lighted the generation hot combustion gas then, and it is passed into turbine afterwards.The turbine bucket of rotation or wheel blade guiding high temperature fluid for example combustion gas pass through turbine.Turbine extracts energy from combustion gas and is used for Driven Compressor, also produces effective merit to drive load, and for example generator perhaps advances aircraft flight.
At least some known turbine vanes comprise the guard shield that stretches out from the exterior extremity of aerofoil, to reduce the air-flow between aerofoil and turbine shroud.At least a portion combustion gas through the turbine guiding desirably do not import between tip shield and aerofoil as tip clearance loss.Such tip clearance is lost in the total losses of turbine stage and accounts for about 20-25%.Known tip shield comprises cavity, and this cavity is limited between the leading edge and turbine shroud of turbine vane.Cavity is caught combustion gas and is caused eddy current in cavity, to form.Eddy current is redirected combustion gas and in main flow path, interrupts combustion gas towards main flow path.Yet this interrupts in main airflow path, producing the secondary flow loss, and this has reduced the operating efficiency of turbine.
Summary of the invention
On the one hand, a kind of turbine vane that uses with turbogenerator is provided.This turbine vane comprises the dovetail that the rotor assembly interior with being placed on turbine shroud is connected.The platform that stretches out from dovetail.The aerofoil that stretches out from platform.This aerofoil comprises butt and tip.This tip stretches out towards turbine shroud from butt.Tip shield stretches out from the tip, and this tip shield comprises backplate.The first guard shield track extends first radial distance from backplate towards turbine shroud.The second guard shield track extends second radial distance that is different from first radial distance from backplate towards turbine shroud.
On the other hand, turbine engine system is provided.This turbine engine system comprises housing, and compressor and the turbine that is connected with mode that compressor is communicated with fluid are to admit the air of at least a portion by compressor discharge.Turbine is placed in the housing.Rotor shaft rotatably is connected with turbine.Rotor shaft limits central axis.A plurality of circumferentially spaced turbine vanes are connected with rotor shaft, and each of a plurality of turbine vanes comprises platform.The aerofoil that stretches out from platform.This aerofoil comprises butt and tip.This tip stretches out towards housing from butt.Tip shield stretches out from the tip, and this tip shield comprises backplate.The first guard shield track extends first radial distance from backplate towards turbine shroud.The second guard shield track extends second radial distance that is different from first radial distance from backplate towards housing.
Aspect another, a kind of method that is used for being manufactured on the turbine vane that turbogenerator uses is provided.This method comprises and forms hard ceramic turbine vane core substantially.This core is inserted in the punch die.Turbine vane is cast as and has the aerofoil that stretches out from platform.Wherein aerofoil comprises the tip shield that stretches out from the tip of aerofoil.Tip shield comprises backplate, and the first guard shield track extends first radial distance from backplate, and the second guard shield track extends second radial distance that is different from first radial distance from backplate.
Description of drawings
Fig. 1 is the schematic representation of exemplary turbogenerator.
Fig. 2 is the sectional drawing of the part of the exemplary rotor assembly of a part, and it can use with the gas turbine engine shown in Fig. 1.
Fig. 3 is the fragmentary cross-sectional view along the amplification of a part of rotor assembly shown in 3 Fig. 2 that got of zone.
Fig. 4 is the part perspective view of the rotor assembly shown in Fig. 3.
Fig. 5 is the fragmentary cross-sectional view of alternative rotor assembly, and it can use with the gas turbine engine shown in Fig. 1.
Component list
10 turbine engine systems
12 suction casings
14 compressor section
16 burner sections
18 turbine section
20 discharge section
22 rotor shafts
More than 24 burner
26 loads
28 rotor assembly
30 is multistage
32 stator stators
34 turbine vanes
36 rotor disks
38 central axis
40 turbine shrouds
42 center holes
44 disk bodies
46 inner edges radially
48 radially outer edges
50 upstream face
52 relative downstream surface
54 support arms
56 spaces
58 rows
60 rows
62 combustion gas path
64 aerofoils
66 tip shields
68 platforms
70 shanks
72 dovetails
74 front shrouds
76 back shrouds
Angel's wing before 78
Buffer cavity before 80
82 back angel's wings
84 back buffer cavitys
86 preceding bottom angel's wings
88 butts
90 tips
92 radial lengths
94 on the pressure side
96 suction side
98 leading edges
100 trailing edges
102 axial widths
104 first axial widths
106 second axial widths
108 guard shield tracks
110 backplates
112 front surfaces
Surface, 114 back
116 first outer rims
118 second outer rims
120Z type groove
122 circumferential width
124 axial lengths
126 sidewalls
128 upstream face
130 downstream surface
132 circumferential width
134 radially-outer surfaces
136 inner radial surface
138 radial heights
140 plate outer surfaces
142 radial distances
144 shell inner surface
144 internal surfaces
146 wearing layers
148 cuttings
150 first guard shield tracks
152 second guard shield tracks
154 first chambeies
154 ante-chambers
156 second chambeies
158 the 3rd chambeies
160 most advanced and sophisticated fluid flow path
162 first radial distances
164 second radial distances
166 first wearing faces
168 second wearing faces
170 tangent line raceway surfaces
172 guard shield surface plane
174 midline plane
176 combustion gas
178 radial distances
Embodiment
To describe exemplary method and system at this, reduce the tip shield that forms near the leading edge vortex of turbine vane, overcome the shortcoming of the known turbine vane of at least a portion through a kind of promotion is provided.More precisely, embodiment described here provides a kind of tip shield, and it comprises a plurality of guard shield tracks with radial height, and it has reduced to be limited to the size of the cavity between tip shield and the turbine shroud, makes the eddy current that has reduced in cavity form.
In this article, employed term " upper reaches " refers to the front end or the input end of gas turbine engine, and term " downstream " refers to the afterbody or the nozzle end of gas turbine engine.
Fig. 1 is the schematic representation of exemplary turbine engine system 10.In exemplary embodiment, turbine engine system 10 comprises entrance 12; Compressor section 14, it is connected in the downstream of entrance 12 and connects; Burner section 16 connects, and its downstream in compressor section 14 connect; Turbine section 18, its downstream that are connected burner section 16 connect; And exhaust section 20, it is connected with turbine section 18.Turbine section 18 is connected through rotor shaft 22 with compressor section 14.In exemplary embodiment, burner section 16 comprises a plurality of burners 24.Burner section 16 is connected with compressor section 14, therefore makes 14 one-tenth fluids of each burner 24 and compressor section be communicated with.Turbine section 18 is connected with compressor section 14, and is connected with load 26, and load for example, and is but whether restrictive, and is limited to the electrical power generator generator and/or the mechanical transmission application apparatus is connected.In exemplary embodiment, each compressor section 14 comprises at least one rotor assembly 28 that at least one is connected with rotor shaft 22 with turbine section 18.
In operation, entrance 12 guiding air are towards compressor section 14, and its air was compressed to higher pressure and temperature before being discharged into burner section 16.Burner section 16 is pressurized air and fuel mix, and the fire fuel air mixture to be producing combustion gas, and with combustion gas guided turbine section 18.More precisely, in burner 24, fuel (for example rock gas and/or fuel oil) is injected in the air-flow, and fuel air mixture is lighted the high-temperature combustion gas that is directed to turbine section 18 with generation.Turbine section 18 will be the mechanical rotation ability from the thermal power transfer of air-flow, can give turbine section 18 and rotor assembly 28 because combustion gas will rotate.
Fig. 2 is the fragmentary cross-sectional view of a part of rotor assembly 28, and it can use with turbine engine system 10.Fig. 3 is the fragmentary cross-sectional view of the amplification of 3 rotor assembly 28 got along the zone.Fig. 4 is the part perspective view of rotor assembly 28.In exemplary embodiment, turbine section 18 comprises a plurality of levels 30, and its stator stator 32 and one that comprises that separately a row is fixing is arranged the turbine vane 34 of rotation.Each extends radially outwardly turbine vane 34 since rotor disk 36.Each rotor disk 36 is connected with rotor shaft 22, and around the central axis that limits rotor shaft 22 38 rotations.Turbine shroud 40 around rotor assembly 28 and stator stator 32 at extending circumferentially.Stator stator 32 is connected with housing 40 separately, and radially extends to rotor shaft 22 to interior from housing 40.
In exemplary embodiment, each rotor disk 36 is annular, and comprises and axially pass wherein the center hole 42 that extends substantially.More precisely, disk body 44 extends radially outwardly from center hole 42, and is orientated substantially perpendicular to central axis 38.The size of center hole 42 is set to admit the rotor shaft 22 that passes wherein.Disk body 44 is radially radially extending between inner edge 46 and the radially outer edge 48, and axially 52 extensions from upstream face 50 to relative downstream surface.Each of upstream face 50 and downstream surface 52 is all extended between inner edge 46 and outer rim 48.Support arm 54 is connected between the contiguous rotor disk 36 to form rotor assembly 28.
Each turbine vane 34 is connected with disk body outer rim 48 and centers on rotor disk 36 circumferentially at interval.Contiguous rotor disk 36 be oriented make circumferentially spaced turbine vane 34 respectively arrange 58 between limit space 56.The size in space 56 is for admitting a row 60 stator stators 32, and stator stator 32 around rotor shaft 22 circumferentially at interval.Stator stator 32 is orientated combustion gas downstream towards turbine vane 34 guiding.Combustion gas path 62 is limited between turbine shroud 40 and each rotor disk 36.Turbine vane 34 and stator stator 32 respectively arrange 58 and 60 extend through combustion gas path 62 at least in part a part.
In exemplary embodiment, each turbine vane 34 extends radially outwardly from disk body 44, and each includes aerofoil 64, tip shield 66, platform 68, shank 70 and dovetail 72.Aerofoil 64 radially extends between platform 68 and tip shield 66 usually.Platform 68 extends between aerofoil 64 and shank 70, makes each aerofoil 64 extend radially outwardly to turbine shroud 40 from platform 68.Shank 70 extends radially inwardly dovetail 72 from platform 68.Dovetail 72 extends radially inwardly from shank 70, and makes turbine vane 34 to be connected with rotor disk 36 reliably.Shank 70 comprises front shroud 74 and relative back shroud 76.
In exemplary embodiment, preceding angel's wing 78 stretches out to promote sealing to be limited to the preceding buffer cavity 80 between rotor disk upstream face 50 and the stator stator 32 from front shroud 74.Back angel's wing 82 stretches out to promote sealing to be limited to the back buffer cavity 84 between rotor disk downstream surface 52 and the stator stator 32 from back shroud 76.In exemplary embodiment, preceding bottom angel's wing 86 stretches out to promote the sealing between turbine vane 34 and rotor disk 36 from front shroud 74.More precisely, preceding bottom angel's wing 86 is positioned between dovetail 72 and the preceding angel's wing 78.
In exemplary embodiment, aerofoil 64 radially extends between butt 88 and most advanced and sophisticated 90, and comprises the radial length 92 that is limited between butt 88 and most advanced and sophisticated 90.Butt 88 is near platform 68.Aerofoil 64 is 40 extensions from platform 68 radially outwards towards turbine shroud, make most advanced and sophisticated 90 contiguous turbine shrouds 40 place.In exemplary embodiment, aerofoil 64 have on the pressure side 94 with suction side 96.Each side 94 and 96 is axially extended between leading edge 98 and trailing edge 100 usually.94 concave surfaces normally on the pressure side, suction side 96 are convex surface normally.In exemplary embodiment, aerofoil 64 has the axial width 102 between leading edge of being limited to 98 and the trailing edge 100.In one embodiment, most advanced and sophisticated 90 have first axial width 104, and butt 88 has second axial width 106 wideer than first axial width 104.
In exemplary embodiment, tip shield 66 stretches out from the tip 90 of aerofoil 64 and between tip 90 and turbine shroud 40, extends.Tip shield 66 comprises a plurality of guard shield tracks 108 that stretch out from backplate 110.In one embodiment, guard shield track 108 is connected with backplate 110.In an alternative, guard shield track 108 integrally forms with backplate 110.In exemplary embodiment, backplate 110 is substantially rectangle, and between front surface 112 and relative back surperficial 114, extends, and between first outer rim 116 and relative circumferentially spaced second outer rim 118, extends.Z type groove 120 is limited in each of first outer rim 116 and second outer rim 118 to promote backplate 110 to be connected with the backplate 110 that is close to.In exemplary embodiment, backplate 110 has the circumferential width 122 that is limited between edge 116 and 118.Backplate 110 also has the axial length 124 that is limited between the surface 112 and 114.In exemplary embodiment, axial length 124 equals most advanced and sophisticated 90 axial width 104 approx.Alternatively, backplate 110 can have greater than with axial length 124 less than circumferential width 104.
In exemplary embodiment, each guard shield track 108 comprises sidewall 126, and it comprises upstream face 128 and downstream surface 130.Guard shield track 108 has the circumferential width 132 that is limited between plate edge 116 and 118.Each sidewall 126 radially extends between radially-outer surface 134 and inner radial surface 136 usually, and has the radial height 138 that is limited between the surface 136 and 134.Extend between surface 136 and 134 on each surface 128 and 130.Inner radial surface 136 is stretched out from the outer surface 140 of backplate 110.Sidewall 126 140 extends radial distances 142 to inner radial surface 134 towards turbine shroud 40 from outer surface.
In exemplary embodiment, turbine shroud 40 comprises the internal surface 144 of circumscribed rotor assembly 28.Internal surface 144 comprises wear-resistant material layer 146.Guard shield track 108 be provided with adjacent inner surface 144, make radially-outer surface 134 contact at least a portion wearing layers 146, make a part of wearing layer 146 when rotor assembly 28 rotations owing to turbine vane 34 thermal expansions are removed.In one embodiment, guard shield track 108 comprises at least one cutting 148, and it stretches out from surface 128 and 130.Each cutting 148 promotes to remove wearing layer 146 when rotor assembly 28 rotations.
In exemplary embodiment, tip shield 66 comprises the first guard shield track 150 and the second guard shield track 152.The first guard shield track 150 is provided with to such an extent that compare the second guard shield track 152 more near front surface 112.The first guard shield track 150 is provided with near front surface 112, makes the chamber 154 of winning, and promptly ante-chamber is limited between shell inner surface 144 and each upstream face 128.The second guard shield track 152 makes second chamber 156 from the first guard shield track 150 along central axis 38 axially at interval, and promptly inner chamber is limited to internal surface 144, and plate outer surface 140 is between track downstream surface 130 and the track upstream face 128.The second guard shield track 152 back surface 114 relatively is provided with, and makes the 3rd chamber 158, i.e. back cavity is limited to the downstream surface 130 of the second guard shield track 152, between plate outer surface 140 and the internal surface 144.In exemplary embodiment, the contiguous separately shell inner surface 144 of track 150 and 152 makes most advanced and sophisticated fluid flow path 160 be limited between tip shield 66 and the internal surface 144.At least a portion combustion gas that most advanced and sophisticated fluid flow path 160 guiding import through rotor assembly 28 between tip shield 66 and turbine shroud 40.
In exemplary embodiment, the first guard shield track 150 extends first radial distance 162 between plate outer surface 140 and track outer surface 134.The second guard shield track 152 extends second radial distance 164 between outer surface 140 and track outer surface 134.In exemplary embodiment, first radial distance 162 is different from second radial distance 164.For example: in one embodiment, second radial distance 164 of first radial distance, 162 to the second guard shield tracks 152 is short.In another alternative, first radial distance 162 be second radial distance 164 about 40% to about 60% between.
In exemplary embodiment, turbine shroud 40 comprises first wearing face 166 of the contiguous first guard shield track 150 and second wearing face 168 of the contiguous second guard shield track 152.Each of first wearing face 166 and second wearing face 168 is parallel with central axis 38 substantially.In exemplary embodiment, each track 150 is parallel with central axis 38 substantially with 152 radially-outer surface 134.Alternatively, outer surface 134 can be orientated corresponding central axis 38 inclinations.
In exemplary embodiment, tangent line raceway surface 170 is limited between the radially-outer surface 134 of track 150 and 152.Guard shield surface plane 172 is 112 114 qualifications to the surface, back along plate outer surface 140 from front surface.Midline plane 174 is defined as parallel with the central axis 38 of axle.In exemplary embodiment, relative second backplate of the first guard shield track 150 110 sizes be provided with and be oriented make tangent line orbit plane 170 with respect to guard shield surperficial 172 to be limited to first jiao of α between tangent line orbit plane 170 and the guard shield surperficial 172 1 Diagonally extending.Backplate 110 is directed obliquely with respect to midline plane 174, makes second inclined angle alpha 2Be limited between guard shield surface plane 172 and the midline plane 174.In one embodiment, first inclined angle alpha 1Greater than second inclined angle alpha 2In another interchangeable embodiment, first inclined angle alpha 1Less than second inclined angle alpha 2
In operation, compressor section 14 (as shown in Figure 1) pressurized air and the air that will compress are discharged in the burner section 16 (as shown in Figure 1) side by side to turbine section 18.Most of air of discharging from compressor section 14 guides towards burner section 16.More precisely, the pressurized air of pressurization is directed to burner 24 (as shown in Figure 1), lights after its air and the fuel mix to produce high-temperature combustion gas 176.Combustion gas 176 are towards combustion gas path 62 guidings, and wherein gas 176 impulse turbine wheel blades 34 act on the rotor assembly 28 with the promotion rotating force with stator stator 32.At least a portion combustion gas 176 impulse turbine wheel blades 34, gas is directed in the most advanced and sophisticated flow path 160 and between tip shield 66 and turbine shroud 40 and is directed.When combustion gas 176 flow through most advanced and sophisticated fluid flow path 160; The size setting of the first guard shield track 150 also is orientated the ante-chamber 154 that limits a certain size promoting to reduce the formation along the eddy current of most advanced and sophisticated fluid flow path 160, and the fluid that reduces between combustion gas path 62 and most advanced and sophisticated fluid flow path 160 is interfered.
Fig. 5 is the sectional drawing of the alternative of the rotor assembly 28 that shows among Fig. 3.Identical reference character is represented among identical parts utilization and Fig. 3 in Fig. 5.In alternative, the first guard shield track 150 comprises first orbital distance 162, and its second orbital distance 164 than the second guard shield track 152 is long.In one embodiment, the size of the first guard shield track, the 150 relative second guard shield tracks 152 is arranged so that tangent line orbit plane 170 is parallel with central axis 38 substantially.In such embodiment, each of first wearing face 166 and second wearing face 168 is parallel with central axis substantially, and they are placed on apart from midline plane 174 approximately equalised radial distances 178 places separately.In this embodiment; The second guard shield track 152 limits back cavity 158; Its size, shape and orientation are set to promote to reduce the formation along the eddy current of most advanced and sophisticated fluid flow path 160, and the fluid that reduces between combustion gas path 62 and most advanced and sophisticated fluid flow path 160 is interfered.
In one embodiment, turbine vane 34 is made through the casting core (not shown).This core is made through liquid ceramics and black lead wash are injected in the core punch die (not shown), and slurry is heated to form the solid ceramic blade core.This blade core suspends in blade punch die (not shown), and hot wax is injected in the blade punch die to surround the ceramic blade core.Hot wax solidifies and forms the wax blade that has the ceramic core that is suspended in the blade.
Wax blade with ceramic core repeatedly is immersed in the ceramic slurry to form ceramic package in the wax blade exterior.Core, wax and shell group are heated to high temperature to remove wax and to be formed on the mold that wherein has ceramic core then.Molten metal is injected in the hollow mold then.Molten metal replaces the wax blade, and forms the metal turbine vane, and ceramic core is held in place.Cooling turbine wheel blade and remove ceramic core then.
Above-mentioned turbine vane forms at least some shortcomings that eddy current has overcome known turbine vane through reducing in the combustion gas path between turbine vane and turbine shroud.More precisely, through the tip shield that comprises a plurality of guard shield tracks with different radial heights is provided, reduced to be limited to the size of the cavity between turbine vane and the turbine shroud.Through reducing the size in chamber, the forming of eddy current of the combustion gas that are redirected of the main flow path that limits between aerofoil reduced.In addition, the secondary flow loss that in main gas path, produces has reduced, thereby has reduced the gas energy loss, and has increased the working life of turbogenerator.
The turbine vane that is used for gas turbine engine and the exemplary embodiment of assembly method thereof have more than been described in detail.The concrete embodiment that this method and apparatus is not restricted to describe among this paper, but the parts of system and/or the step of method can be implemented with other parts and/or the method for describing in this article independently and respectively.For example: method and apparatus can be united use with other combustion systems and method, and what be not restricted to describe in this article only implements as gas turbine engine.But exemplary embodiment can be implemented with a lot of other combustion systems application and use explicitly.
Although a plurality of embodiments' of the present invention concrete characteristic shows in some drawings, and not demonstration in other accompanying drawings, this only is the convenience in order to show.In addition, above-mentioned " embodiment " and do not mean that and be interpreted as the existence of getting rid of the other embodiment who also comprises described characteristic.According to principle of the present invention, any characteristic among the figure can be quoted and/or ask for protection with any characteristic binding of other accompanying drawing.
The mode that this written explanation utilization is given an example is come open the present invention, comprises best mode of execution, makes that also any those skilled in the art can embodiment of the present invention, comprises making and utilizing other device or system and implement any method that combines.The scope of patentability of the present invention is defined by the claims, and can comprise other embodiment who expects easily to those skilled in the art.These other embodiment; If the element of construction that they have does not have different with the literal language of claim; If perhaps they comprise that literal language with claim does not have the equivalent structure unit of difference on the entity, mean that then these other embodiment is in the scope of claim.

Claims (10)

1. turbine vane (34) that uses with turbogenerator (10), said turbine vane comprises:
The dovetail (72) that the rotor assembly (28) interior with being placed on turbine shroud (40) is connected;
The platform (68) that stretches out from said dovetail;
From the aerofoil (64) that said platform stretches out, said aerofoil comprises butt (88) and most advanced and sophisticated (90), and said tip stretches out towards turbine shroud from said butt; With
The tip shield (66) that stretches out from said tip, said tip shield comprises:
Backplate (110);
Extend the first guard shield track (150) of first radial distance (162) from said backplate towards turbine shroud; With
Extend the second guard shield track (152) of second radial distance (164) from said backplate towards turbine shroud, this second radial distance is different from said first radial distance.
2. turbine vane according to claim 1 (34); It is characterized in that; Said backplate totally extends between front surface (112) and surface, back (114); The said first guard shield track (150) place than the said second guard shield track (152) more near said front surface, make axial flow path be limited between turbine shroud (40) and the said tip shield (66).
3. turbine vane according to claim 2 (34) is characterized in that, said first radial distance (162) is longer than said second radial distance (164).
4. turbine vane according to claim 2 (34) is characterized in that, said first radial distance (162) is shorter than said second radial distance (164).
5. turbine vane according to claim 1 (34) is characterized in that, said first radial distance (162) said second radial distance (164) about 40% to 60% between.
6. turbine vane according to claim 3 (34); It is characterized in that; The said first guard shield track (150) is placed apart from said front surface (112) a distance; Make between said front surface and turbine shroud (40), to define cavity that said cavity helps minimizing in said cavity, to form eddy current.
7. turbine vane according to claim 4 (44); It is characterized in that; The said second guard shield track (150; 152) place apart from (114) a distance, said back surface, make between surface, said back and turbine shroud (40), to define cavity (154), said cavity helps minimizing in said cavity, to form eddy current.
8. turbine vane according to claim 1 (34); It is characterized in that; The said first and second guard shield tracks (150,152) limit tangent line raceway surface (170), and said tangent line raceway surface (170) is with respect to the angular orientation of outer surface (140) to tilt of said backplate (110).
9. turbine engine system comprises:
Housing (40);
Compressor (14);
The turbine that is communicated with said compressor (14) fluid is with at least a portion air that admittance is discharged through said compressor, in the said housing that said turbine is placed;
The rotor shaft (22) that rotatably is connected with said turbine, said rotor shaft limits central axis (38); With
The a plurality of circumferentially spaced turbine vane (34) that is connected with said rotor shaft, each of said a plurality of turbine vanes includes:
Platform (68);
From the aerofoil (64) that said platform stretches out, said aerofoil comprises butt (88) and most advanced and sophisticated (90), and said tip stretches out from said butt towards said housing; With
The tip shield (66) that stretches out from said tip, said tip shield comprises:
Backplate (110);
Extend the first guard shield track (150) of first radial distance (162) from said backplate towards said housing; With
Extend the second guard shield track (152) of second radial distance (164) from said backplate towards said housing, this second radial distance is different from said first radial distance.
10. turbine engine system according to claim 9 (10); It is characterized in that; Said backplate totally extends between front surface (112) and surface, back (114); The said first guard shield track (150) place than the said second guard shield track more near said front surface, make axial flow path be limited between housing (40) and the said tip shield (66).
CN2012100592467A 2011-01-28 2012-01-28 Turbine bucket for use in gas turbine engines and methods for fabricating the same Pending CN102619570A (en)

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US13/015,747 US20120195742A1 (en) 2011-01-28 2011-01-28 Turbine bucket for use in gas turbine engines and methods for fabricating the same
US13/015,747 2011-01-28

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CN102619570A true CN102619570A (en) 2012-08-01

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