CN107131005A - Turbine engine shroud component - Google Patents
Turbine engine shroud component Download PDFInfo
- Publication number
- CN107131005A CN107131005A CN201710113770.0A CN201710113770A CN107131005A CN 107131005 A CN107131005 A CN 107131005A CN 201710113770 A CN201710113770 A CN 201710113770A CN 107131005 A CN107131005 A CN 107131005A
- Authority
- CN
- China
- Prior art keywords
- turbogenerator
- flange
- airfoil
- blocking element
- blade
- 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
Links
- 230000000903 blocking effect Effects 0.000 claims abstract description 33
- 230000036316 preload Effects 0.000 claims description 5
- 239000003570 air Substances 0.000 description 9
- 238000005452 bending Methods 0.000 description 8
- 238000000034 method Methods 0.000 description 6
- 230000003068 static effect Effects 0.000 description 5
- 230000015572 biosynthetic process Effects 0.000 description 4
- 239000000567 combustion gas Substances 0.000 description 4
- 239000012530 fluid Substances 0.000 description 4
- 238000011144 upstream manufacturing Methods 0.000 description 4
- 238000001816 cooling Methods 0.000 description 3
- 239000011258 core-shell material Substances 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
- 238000005516 engineering process Methods 0.000 description 3
- 239000007789 gas Substances 0.000 description 3
- 238000004519 manufacturing process Methods 0.000 description 3
- 230000007935 neutral effect Effects 0.000 description 3
- 239000012080 ambient air Substances 0.000 description 2
- 230000000712 assembly Effects 0.000 description 2
- 238000000429 assembly Methods 0.000 description 2
- 230000006835 compression Effects 0.000 description 2
- 238000007906 compression Methods 0.000 description 2
- 238000010276 construction Methods 0.000 description 2
- 238000013016 damping Methods 0.000 description 2
- 230000008450 motivation Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 238000002485 combustion reaction Methods 0.000 description 1
- 230000001276 controlling effect Effects 0.000 description 1
- 239000012809 cooling fluid Substances 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 239000000203 mixture Substances 0.000 description 1
- ORQBXQOJMQIAOY-UHFFFAOYSA-N nobelium Chemical compound [No] ORQBXQOJMQIAOY-UHFFFAOYSA-N 0.000 description 1
- 238000007363 ring formation reaction Methods 0.000 description 1
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
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
-
- 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
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
-
- 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
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
- F01D25/246—Fastening of diaphragms or stator-rings
-
- 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
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
-
- 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
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
-
- 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
- F05D2250/32—Arrangement of components according to their shape
-
- 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
- F05D2260/00—Function
- F05D2260/30—Retaining components in desired mutual position
- F05D2260/36—Retaining components in desired mutual position by a form fit connection, e.g. by interlocking
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Architecture (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application is related to turbine engine shroud component.Wherein, one kind is used for the interlocking cover assembly (100 of turbogenerator (10), 200) circumferentially spaced airfoil (70) is radially extended with multiple, it terminates at sheath elements (102,202), and the opposite radial side (106,108) with the blocking element of band first (134) and the second blocking element (138).
Description
Technical field
Turbogenerator, and particularly combustion gas or combustion turbine engine, are to flow to multiple rotation whirlpools from through engine
The pressurized combustion gases of impeller blade obtain the rotary engine of energy.
Background technology
Rotary turbine blade can be by supporting hood shield, and shield interlocks to form the circumferential shell of turbine.Z-shaped interlocking is needed in system
The typical construction for the blade assembly with shield made and distorted in advance during assembling is selected.Eliminated while interlocking formation is kept
Distortion in advance will be beneficial for cover assembly manufacture.
The content of the invention
On the one hand, embodiment is related to a kind of turbogenerator, it include with circumferentially around rotor intervals open it is multiple
The rotor of the airfoil (airfoil) radially extended, wherein airfoil terminates at tip (tip), and cover assembly is external
(circumscribing) airfoil, and the sheath elements including being installed on each tip, and with the blocking element of band first
(interlock element) and the second blocking element opposite radial side, the first blocking element of one of sheath elements
Match (mate) with the second blocking element of circumferentially-adjacent element, with formed around airfoil circumference adjacent sheath elements it
Between multiple locking parts (interlocks).
On the other hand, embodiment is related to a kind of interlocking cover assembly for turbogenerator, including multiple radially prolongs
Circumferentially spaced airfoil is stretched, it terminates at sheath elements and the phase with the blocking element of band first and the second blocking element
To radial side, it interlocks to form the locking part between circumferentially-adjacent airfoil.
Another aspect, embodiment is related to a kind of side for the shield for being formed and surrounding multiple rotating vanes in turbogenerator
Method, including the locking part formed between the circumferentially-adjacent tip of blade, and preload (preloading) locking part.
A kind of turbogenerator of technical scheme 1., including：
Rotor, it has the multiple airfoils radially extended opened circumferentially around the rotor intervals, and the airfoil
Terminate at tip；And
Cover assembly, its external described airfoil and the sheath elements including being installed on each tip, and the sheath elements
Relative radial side with the blocking element of band first and the second blocking element；
Wherein, the first blocking element of a sheath elements is matched with the second blocking element of circumferentially-adjacent element, to be formed
Multiple locking parts between the adjacent sheath elements of the circumference of the airfoil.
Turbogenerator of the technical scheme 2. according to technical scheme 1, it is characterised in that the sheath elements and institute
Airfoil is stated to be integrally formed.
Turbogenerator of the technical scheme 3. according to technical scheme 2, it is characterised in that the airfoil is blade.
Turbogenerator of the technical scheme 4. according to technical scheme 3, it is characterised in that the blade terminate at
The relative dovetail of the tip, and the dovetail is installed on the rotor.
Turbogenerator of the technical scheme 5. according to technical scheme 1, it is characterised in that airfoil is sized to
So that being arch (sprung) when the airfoil and adjacent aerofoil part are interlocked, to apply preloading to the locking part
On.
Turbogenerator of the technical scheme 6. according to technical scheme 1, it is characterised in that the first locking part bag
The first flange is included, second locking part includes the second flange, and the first flange overlying (overly) and adjoining
(abuts) second flange.
Turbogenerator of the technical scheme 7. according to technical scheme 6, it is characterised in that first locking part is also
Including with the circumferentially spaced bearing of the first flange, and second flange be located at the bearing in.
Turbogenerator of the technical scheme 8. according to technical scheme 7, it is characterised in that the bearing is by the first footpath
Formed to edge and first flange.
Turbogenerator of the technical scheme 9. according to technical scheme 8, it is characterised in that the first flange edge week
To protruding past first radial edges.
Turbogenerator of the technical scheme 10. according to technical scheme 9, it is characterised in that first flange is in institute
State the outer circumferential of the first radial edges.
Turbogenerator of the technical scheme 11. according to technical scheme 10, it is characterised in that second flange is
Second radial edges of adjacent sheath elements.
Turbogenerator of the technical scheme 12. according to technical scheme 11, it is characterised in that airfoil size is set
Into make it that when the airfoil and adjacent aerofoil part are interlocked be arch, it is applied to preload on the locking part, so that
Cause second radial edges towards the first flange biases.
A kind of interlocking cover assembly for turbogenerator of technical scheme 13., including multiple radially extend circumferentially
Airfoil spaced apart, it terminates at sheath elements and has the blocking element of band first and the second blocking element relatively radially
Side, it interlocks to form the locking part between circumferentially-adjacent airfoil.
Interlocking cover assembly of the technical scheme 14. according to technical scheme 13, it is characterised in that airfoil size is set
It is set to so that being arch when the airfoil and adjacent aerofoil part are interlocked, to apply preloading to the locking part.
Interlocking cover assembly of the technical scheme 15. according to technical scheme 14, it is characterised in that first interlocking
Part includes the first flange, and second locking part includes the second flange, and the first flange overlying and adjacent described second convex
Edge.
Interlocking cover assembly of the technical scheme 16. according to technical scheme 15, it is characterised in that first interlocking
Part also include with the circumferentially spaced bearing of the first flange, and second flange be located at the bearing in.
Interlocking cover assembly of the technical scheme 17. according to technical scheme 16, it is characterised in that the bearing is by
One radial edges and first flange are formed.
Interlocking cover assembly of the technical scheme 18. according to technical scheme 17, it is characterised in that first flange
Circumferentially protrude past first radial edges.
Interlocking cover assembly of the technical scheme 19. according to technical scheme 18, it is characterised in that first flange
In the outer circumferential of first radial edges.
Interlocking cover assembly of the technical scheme 20. according to technical scheme 19, it is characterised in that second flange
For the second radial edges of adjacent sheath elements.
A kind of method for forming the shield for surrounding multiple rotating vanes in turbogenerator of technical scheme 21., including shape
Locking part between the circumferentially-adjacent tip of blade, and preload the locking part.
Method of the technical scheme 22. according to technical scheme 21, it is characterised in that formation locking part is included in described
Locking part is formed in the circumferentially opposite sides of blade.
Method of the technical scheme 23. according to technical scheme 21, it is characterised in that preloading the locking part includes making
The blade arches upward (springing).
Brief description of the drawings
In the accompanying drawings：
Fig. 1 is the schematic sectional view of the turbogenerator for aircraft.
Fig. 2 is multiple airfoils of assembling.
Fig. 3 is the perspective view of sheath elements.
Fig. 4 is another perspective view of sheath elements.
Fig. 5 is the diagram of cover assembly.
Fig. 6 is the section view of Fig. 5 cover assembly.
Fig. 7 is the section view of the second embodiment of Fig. 5 cover assembly.
Fig. 8 is single airfoil component.
Parts List：
10 engines
12 center lines
14 is anterior
16 rear portions
18 fan sections
20 fans
22 compressor sections
24 LP compressors
26 HP compressors
28 burning blocks
30 burners
32 turbines
34 HP turbines
36 LP turbines
38 exhaust sections
40 fan hubs
42 fan blade
44 cores
46 core shells
48 HP rotating shafts
50 LP rotating shafts
52 HP compressor stages
53 rotors
54 HP compressor stages
56 LP compressor blades
58 HP compressor blades
59 disks
60 LP compressor vanes
61 disks
62 HP compressor vanes
63 stators
64 HP stage of turbines
66 LP stage of turbines
68 HP turbo blades
70 LP turbo blades
71 disks
72 HP turbine guide vanes
73 disks
74 LP turbine guide vanes
76 pressurized ambient airs
77 release air
78 air streams
80 exit guide blade components
82 airfoil stators
84 fan exhaust sides
88 roots
90 tips
96 inner ring
98 outer shrouds
100 cover assemblies
102 sheath elements
104 flanges
106 first radial edges
107 radial sides
108 second radial edges
109 radial sides
110 bearings
112 parts
114 airfoil components
116 dovetails
118 arches (sprung)
120 initial parallel positions
122 neutral axis
124 final bending positions
126 initial bending positions
128 final parallel positions
130 locking parts
132 first flanges
134 first blocking elements
136 second flanges
138 second blocking elements
200 cover assemblies
202 sheath elements
204 flanges
206 first radial edges
208 second radial edges
210 bearings
212 parts
214 airfoil components
216 dovetails
218 arches
220 initial parallel positions
222 neutral axis
224 final bending positions
226 initial bending positions
228 final parallel positions
230 locking parts
232 first flanges
234 first blocking elements
236 second flanges
238 second blocking elements.
Embodiment
The embodiment introduced of the present invention is directed to the cover assembly for airfoil.For purposes of illustration, by reference
The present invention is described for the turbine of aircraft turbine engine.It is to be understood, however, that the present invention is not so limited,
And can have in-engine common application, including compressor, and non-aircraft in application, e.g., other Mobile solutions and
Non-moving industry, business and residential application.
Term " preceding " as used herein or " upstream " refer to along being moved towards the direction of motor inlet, or component compared to
Another component is relatively closer to motor inlet.The term " rear " used together with " preceding " or " upstream " or " downstream " refer to towards hair
Motivation is on the rear portion of engine centerline or the direction of outlet.
In addition, term " radial direction " or " radially " refer to outside the central longitudinal axis and engine of engine as used herein
The dimension extended between week.
All directions represent (for example, radially, axially, proximal and distal, it is upper and lower, upward, downward, left and right, lateral, preceding,
Afterwards, top, bottom, top, lower section, vertical, level, clockwise, counterclockwise, upstream, downstream, rear etc.) it is only used for recognizing mesh
, to contribute to reader to understand the present invention, and do not produce especially with regard to position of the invention, the limitation for orienting or using.Even
Connecing expression (for example, attachment, connection, connection and link) will be to be construed broadly as, and may include a series of centre between elements
Relative movement between part, and element, unless otherwise noted.Therefore, to be not necessarily referring to two elements direct for connection expression
Ground connect, and with each other in fixed relationship.Schematic diagram merely for diagram purpose, and reflect in appended accompanying drawing size, position,
Order and relative size alterable.
Fig. 1 is the schematic sectional view of the turbogenerator 10 for aircraft.Engine 10 has from preceding 14 backward 16
The axis or center line 12 of the generally longitudinal extension of extension.Engine 10 is included with downstream series flow relationship：Including fan 20
Fan section 18 including booster or low pressure (LP) compressor 24 and high pressure (HP) compressor 26 compressor section 22, bag
Include the burning block 28 of burner 30 including the turbine 32 of HP turbines 34 and LP turbines 36, and exhaust section 38.
Fan section 18 includes the fan hub 40 of wrapping fan 20.Fan 20 includes what is radially extended around center line 12
Multiple fan blade 42.The core 44 of HP compressors 26, burner 30 and the formation engine 10 of HP turbines 34, it generates combustion gas
Body.Core 44 is wrapped by core shell 46, and core shell 46 can couple with fan hub 40.
HP turbines 34 are drivingly connected to by the HP axles or rotating shaft 48 being coaxially disposed around the center line 12 of engine 10
On HP compressors 26.The LP axles in larger-diameter annular HP rotating shafts 48 are placed coaxially on around the center line 12 of engine 10
Or LP turbines 36 are drivingly connected on LP compressors 24 and fan 20 by rotating shaft 50.
Compressor 24 and HP compressors 26 include multiple compressor stages 52,54, one of which compressor blade 56,58 respectively
On corresponding one group of static compressor vanes 60,62 (also referred to as nozzle) rotation, to compress or pressurize through the fluid of level
Stream.In single compressor stage 52,54, multiple compressor blades 56,58 can cyclization provide, and can be on center line 12 from blade
Platform is extended radially outward to blade tips, while corresponding static compressor vanes 60,62 are positioned at rotating vane 56,58
Upstream and adjacent thereto.It will be noted that, the number of blade, stator and compressor stage shown in Fig. 1 is only for exemplary mesh
Selection, and other numbers are possible.
The blade 56,58 of compressor stage may be mounted on disk 59, and disk 53 is installed to the correspondence one in HP and LP rotating shafts 48,50
On individual, wherein each grade has the disk 59,61 of its own.The stator 60,62 of the level of compressor can be by being circumferentially installed to core
On heart shell 46.
Turbine 34 and LP turbines 36 include multiple stage of turbines 64,66 respectively, and one of which turbo blade 68,70 is on correspondence
One group of static turbine guide vane 72,74 (also referred to as nozzle) rotation, to obtain energy from the fluid stream through level.In single stage of turbine
In 64,66, multiple turbine guide vanes 72,74 may be provided in ring, and can be extended radially outward on center line 12, and correspond to simultaneously
Rotating vane 68,70 be positioned at the static downstream of turbine guide vane 72,74 and adjacent thereto, and can also be last from bucket platform to blade
The tip, is extended radially outward on center line 12.It will be noted that, the number of blade, stator and stage of turbine shown in Fig. 1 is only
Select for exemplary purposes, and other numbers are possible.
The blade 68,70 of stage of turbine may be mounted on disk 71, and disk 71 is installed to the correspondence one in HP and LP rotating shafts 48,50
On, wherein each grade has the disk 71,73 of its own.The stator 72,74 of the level of compressor can be by being circumferentially installed to core
On shell 46.
The part for the engine 10 for being installed in rotating shaft 48,50 and being rotated together with one or both therein also individually or
Commonly referred to as rotor 53.Including the stationary part of the engine 10 of part that is installed in core shell 46 also separately or together
Ground is referred to as stator 63.
In operation, the air stream shunting of outflow fan section 18 a so that part for air stream is sent to LP compressors
24, then pressurized ambient air 76 is supplied to the HP compressors 26 of further pressurized ambient air by it.From HP compressors 26
Forced air 76 mix and light with fuel in burner 30, so as to generate burning gases.Some work(by HP turbines 34 from
These gases are obtained, the driving HP of HP turbines 34 compressors 26.Combustion gases exhaust is into LP turbines 36, and it obtains additional work(to drive
LP compressors 24 are moved, and discharge gas is finally discharged via exhaust section 38 from engine 10.The driving of LP turbines 36 can drive
LP rotating shafts 50 come rotary fan 20 and LP compressors 24.
The remainder of air stream 78 bypasses LP compressors 24 and engine core 44, and is set out via static stator drainage
Motivation component 10, and include the exit guide blade component 80 of multiple airfoil stators 82 more specifically at fan exhaust side 84.More
Specifically, the circumference of the airfoil stator 82 radially extended, which comes fan section 18, to be nearby used to apply the one of air stream 78
A little direction controllings.
Some surrounding airs supplied by fan 20 can bypass engine core 44, and (outstanding for the part of engine 10
It is hot part) cooling, and/or for cooling down or energizing to the other side of aircraft.Under the background of turbogenerator,
The heat part of engine is usually burner 30 and the component in the downstream of burner 40, especially turbine 32, wherein HP turbines
34 be most hot part, because it is directly in the downstream of burning block 28.Other cooling fluid sources can be but be not limited to compress from LP
The fluid that machine 24 or HP compressors 26 are discharged.The fluid can be to release air 77, and it may include from LP compressors 24 or HP compressions
The air that machine 26 absorbs, it bypasses burner 30 as the cooling source for turbine 32.This is common engine construction, no
It is intended to restricted.
Fig. 2 show it is multiple radially extend circumferentially spaced airfoil or blade 70, wherein each blade 70 is equal
Extend from root 88 and terminate at tip (Fig. 3), arrangement is circumferentially arranged and supported by arc inner ring 96 and arc outer shroud 98.Arc
Outer shroud 98 includes being made up of the separate sheath element 102 separated of the opposite radial side 107 and 109 with external blade 70 together
Cover assembly 100.
As shown in Figures 3 and 4, each sheath elements 102 is integrally formed at tip 90 with blade 70, and including flange
104.Flange 104 includes the first radial edges 106 and the second radial edges 108, and bearing 110, and it is circumferentially protruded past
The first radial edges 106 shown in Fig. 3.Bearing 110 is by convex in the part 112 in the first radial edges 106 and flange 104
Edge 104 is formed, the outer circumferential of shown the first longitudinal edge along 106 in Fig. 4 of part 112.
Fig. 5 shows airfoil component 114, and wherein sheath elements 102 are integrally formed with blade 70, and its Leaf is terminated at
Dovetail 116.Dovetail 116 is formed as being installed on rotor 53.In assembling, blade 70 is arranged to arch 118, with to connection
Latch fitting, which applies, to be preloaded.In interlocking, blade 70 can be from neutral axis of the notable parallel position 120 relative to airfoil component 114
Arch shown in line 122 to the solid line 118 of bending position 124.Interlocking when, blade 70 can also be from bend substantially 126 just
Beginning position to parallel position 128 arch 118.
No matter the initial or final position of blade 70, final position 124,128 will all cause the second radial edges 108 outside
Bias and bearing inward bias.The bias is by the compression stress F from dovetail 116CCause, it changes into the second radial edges
Upward power F at 1082With the downward power F from bearing 1101。
Circumferentially-adjacent sheath elements 102 lock together, to form multiple locking parts between adjacent sheath elements 102
130, to form cover assembly 100 as shown in Figure 6.The exemplary embodiment of Fig. 6 cover assembly 100 is shown in Fig. 7
Section view, wherein the first flange 132 with the first blocking element 134 and second convex with the second blocking element 138
Edge 136 is matched, wherein the overlying of the first flange 132 and adjacent second flange 136.
When cover assembly 100 is assembled, the second radial edges 108 of the second flange 136 will be due to power F1And F2It is convex towards first
Edge 132 is biased.The bias allows frictional force formation between the first blocking element and the second blocking element, and this is first by each shield
Part 102 is attached to the sheath elements 102 of next radially adjoining.
The second embodiment of cover assembly has been envisioned in Fig. 8.Second embodiment is similar to first embodiment；Therefore, phase
The like numeral of increase by 100 respectively will be indicated like part, wherein will be appreciated that the description of the similar portion of first embodiment is fitted
For additional embodiment, unless otherwise noted.
In a second embodiment, the first blocking element 234 formed on the first flange 232 includes the bearing 210 of angulation,
It is formed as the second blocking element 238 for storing the angulation formed on the second flange 236.Therefore, each sheath elements 202
Flange 204 includes the bearing 210 of angulation, and is formed as the part 240 of angulation being coupled in the bearing 210 of angulation.To the greatest extent
Pipe is shown as being formed two inclined-planes 242 on summit 246, and 244, but the bearing 210 of angulation and the part 240 of angulation can be the first interlocking
Element 234 is formed as storing any shape of the second blocking element 238.
It is a kind of formed surround turbogenerator in multiple rotating vanes include sheath elements integrally formed with blade
The method of cover assembly include the locking part formed between the circumferentially-adjacent tip of blade, and preloading locking part.Locking part
Preload, wherein making blocking element be biased towards another blocking element.
Embodiment as described herein has the benefit on production, performance and damping capacity.Shield blade assembly it is existing
Technology includes Z-shaped locking part.Implementation airfoil and shield radially bending ensure that the contact between blocking element, with operation shape
Realize that outer shroud is preloaded under state, without the typical torsional bending used in the Z-shaped guard design that needs to distort in advance.Such bending
Also the blade balance for centrifugal force is only needed, and it is common to improve blade caused by the increase contact area between blocking element surface
Damping under shaking.This increase of contact area additionally provides the reduction of outer flow passage leakage.The simplification shape of design and disappear
It is easy to manufacture the need for except advance distortion.
It is to be appreciated that the application of disclosed design is not limited to the propeller for turboprop with fan and booster section
Machine (turbine engines), but it is equally applicable to turbojet (turbojets) and turbo-dynamo (turbo
engines)。
This written description discloses (including optimal mode) of the invention using example, and also makes any technology of this area
Personnel can implement the present invention (including making and use any device or system, and any method being incorporated to of execution).This hair
Bright the scope of the claims is defined by the claims, and may include other examples that those skilled in the art expects.If it is such its
Its embodiment has the structural detail for the written language for being not different from claim, or if they include and claim
Equivalent structural elements of the written language without essential difference, then expect such other examples within the scope of the claims.
Claims (10)
1. a kind of turbogenerator (10), including：
Rotor (53), it has the multiple airfoils (70) radially extended being spaced apart circumferentially around the rotor (53),
And the airfoil (70) terminates at tip (90)；And
Cover assembly (100,200), its external described airfoil (70) and the shield member including being installed on each tip (90)
Part (100,200), and the sheath elements (100,200) have the blocking element of band first (134,234) and the second blocking element
(138,238) relative radial side (106,108,206,208)；
Wherein, the first blocking element (134,234) of a sheath elements (102,202) and circumferentially-adjacent element (102,
202) the second blocking element (138,238) matching, to form the adjacent sheath elements of the circumference around the airfoil (70)
(102,202) multiple locking parts (130,230) between.
2. turbogenerator (10) according to claim 1, it is characterised in that the sheath elements (102,202) and institute
Airfoil (70) is stated to be integrally formed.
3. turbogenerator (10) according to claim 2, it is characterised in that the airfoil (70) is blade (70).
4. turbogenerator (10) according to claim 3, it is characterised in that the blade (70) terminates at and the end
The relative dovetail (116,216) of the tip (90), and the dovetail is installed on the rotor (53).
5. turbogenerator (10) according to claim 1, it is characterised in that the first locking part (134,234) bag
Include the first flange (132,232), second locking part (138,238) includes the second flange (136,236), and it is described first convex
Edge (132,232) overlying and adjacent second flange (136,236).
6. turbogenerator (10) according to claim 5, it is characterised in that first locking part (134,234) is also
Including with the circumferentially spaced bearing of first flange (132,232) (110,210), and second flange (136,
236) it is located in the bearing (110,210).
7. turbogenerator (10) according to claim 6, it is characterised in that the bearing is by the first radial edges
(106,206) formed with first flange (132,232).
8. turbogenerator (10) according to claim 7, it is characterised in that first flange (132,232) is along week
To protruding past first radial edges (106,206) and in the outer circumferential of first radial edges (106,206).
9. turbogenerator (10) according to claim 8, it is characterised in that second flange (136,236) is phase
The second radial edges (108,208) of adjacent sheath elements (102,202).
10. turbogenerator (10) according to claim 9, it is characterised in that airfoil is dimensioned so that in institute
It is arch to state when airfoil (70) is interlocked with adjacent aerofoil part (70), is applied to preload on the locking part, so as to draw
Second radial edges are played towards the first flange biases.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| PLP.416301 | 2016-02-29 | ||
| PL416301A PL416301A1 (en) | 2016-02-29 | 2016-02-29 | Turbine engine shrouding bandage unit |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN107131005A true CN107131005A (en) | 2017-09-05 |
Family
ID=58094284
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201710113770.0A Pending CN107131005A (en) | 2016-02-29 | 2017-02-28 | Turbine engine shroud component |
Country Status (6)
| Country | Link |
|---|---|
| US (1) | US20170306768A1 (en) |
| EP (1) | EP3225794A1 (en) |
| JP (1) | JP2017198190A (en) |
| CN (1) | CN107131005A (en) |
| CA (1) | CA2958106A1 (en) |
| PL (1) | PL416301A1 (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114207256A (en) * | 2019-07-17 | 2022-03-18 | 赛峰飞机发动机公司 | Tool and method for removing and moving a TRV-type turbine casing from an aircraft turbine engine |
| CN115217628A (en) * | 2021-04-16 | 2022-10-21 | 通用电气公司 | Airfoil assembly |
| CN117869016A (en) * | 2024-03-12 | 2024-04-12 | 中国航发四川燃气涡轮研究院 | Cooling unit for reducing heat conduction of turbine outer ring and analysis method thereof |
Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10738634B2 (en) * | 2018-07-19 | 2020-08-11 | Raytheon Technologies Corporation | Contact coupled singlets |
| WO2020263394A1 (en) * | 2019-06-28 | 2020-12-30 | Siemens Aktiengesellschaft | Outlet guide vane assembly in gas turbine engine |
| US11208903B1 (en) | 2020-11-20 | 2021-12-28 | Solar Turbines Incorporated | Stiffness coupling and vibration damping for turbine blade shroud |
| JP7785560B2 (en) * | 2022-02-16 | 2025-12-15 | 三菱重工航空エンジン株式会社 | turbine |
| FR3137121B1 (en) * | 2022-06-22 | 2024-06-21 | Safran Aircraft Engines | Bladed assembly with inter-platform connection by interposed rolling element |
| KR102803494B1 (en) | 2023-01-27 | 2025-05-02 | 두산에너빌리티 주식회사 | Rotor disk assembly and gas turbine comprising it |
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| CN103422903A (en) * | 2012-05-21 | 2013-12-04 | 阿尔斯通技术有限公司 | Turbine diaphragm construction |
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| US20120051930A1 (en) * | 2010-08-31 | 2012-03-01 | General Electric Company | Shrouded turbine blade with contoured platform and axial dovetail |
-
2016
- 2016-02-29 PL PL416301A patent/PL416301A1/en unknown
-
2017
- 2017-02-16 JP JP2017026516A patent/JP2017198190A/en active Pending
- 2017-02-16 CA CA2958106A patent/CA2958106A1/en not_active Abandoned
- 2017-02-20 EP EP17156817.3A patent/EP3225794A1/en not_active Withdrawn
- 2017-02-24 US US15/441,597 patent/US20170306768A1/en not_active Abandoned
- 2017-02-28 CN CN201710113770.0A patent/CN107131005A/en active Pending
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|---|---|---|---|---|
| EP1939401A2 (en) * | 2006-12-19 | 2008-07-02 | General Electric Company | Methods and apparatus for load transfer in rotor assemblies |
| CN101988392A (en) * | 2009-07-31 | 2011-03-23 | 通用电气公司 | Rotor blades for turbine engines |
| CN103168150A (en) * | 2010-10-29 | 2013-06-19 | 通用电气公司 | Anti-rotation shroud for turbine engines |
| EP2551459A2 (en) * | 2011-07-28 | 2013-01-30 | General Electric Company | Cap for ceramic blade tip shroud |
| CN103195505A (en) * | 2012-01-04 | 2013-07-10 | 通用电气公司 | Device and method for aligning tip shrouds |
| CN103422903A (en) * | 2012-05-21 | 2013-12-04 | 阿尔斯通技术有限公司 | Turbine diaphragm construction |
Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114207256A (en) * | 2019-07-17 | 2022-03-18 | 赛峰飞机发动机公司 | Tool and method for removing and moving a TRV-type turbine casing from an aircraft turbine engine |
| CN114207256B (en) * | 2019-07-17 | 2024-04-16 | 赛峰飞机发动机公司 | Tool and method for disassembling and moving a TRV-type turbine casing of an aircraft turbine engine |
| CN115217628A (en) * | 2021-04-16 | 2022-10-21 | 通用电气公司 | Airfoil assembly |
| CN117869016A (en) * | 2024-03-12 | 2024-04-12 | 中国航发四川燃气涡轮研究院 | Cooling unit for reducing heat conduction of turbine outer ring and analysis method thereof |
| CN117869016B (en) * | 2024-03-12 | 2024-05-17 | 中国航发四川燃气涡轮研究院 | Cooling unit for reducing heat conduction of turbine outer ring and analysis method thereof |
Also Published As
| Publication number | Publication date |
|---|---|
| JP2017198190A (en) | 2017-11-02 |
| CA2958106A1 (en) | 2017-08-29 |
| EP3225794A1 (en) | 2017-10-04 |
| PL416301A1 (en) | 2017-09-11 |
| US20170306768A1 (en) | 2017-10-26 |
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